WorldWideScience

Sample records for thermal rocket engines

  1. Unique nuclear thermal rocket engine

    International Nuclear Information System (INIS)

    Culver, D.W.; Rochow, R.

    1993-06-01

    In January, 1992, a new, advanced nuclear thermal rocket engine (NTRE) concept intended for manned missions to the moon and to Mars was introduced (Culver, 1992). This NTRE promises to be both shorter and lighter in weight than conventionally designed engines, because its forward flowing reactor is located within an expansion-deflection rocket nozzle. The concept has matured during the year, and this paper discusses a nearer term version that resolves four open issues identified in the initial concept: (1) the reactor design and cooling scheme simplification while retaining a high pressure power balance option; (2) elimination need for a new, uncooled nozzle throat material suitable for long life application; (3) a practical provision for reactor power control; and (4) use of near-term, long-life turbopumps

  2. Grooved Fuel Rings for Nuclear Thermal Rocket Engines

    Science.gov (United States)

    Emrich, William

    2009-01-01

    An alternative design concept for nuclear thermal rocket engines for interplanetary spacecraft calls for the use of grooved-ring fuel elements. Beyond spacecraft rocket engines, this concept also has potential for the design of terrestrial and spacecraft nuclear electric-power plants. The grooved ring fuel design attempts to retain the best features of the particle bed fuel element while eliminating most of its design deficiencies. In the grooved ring design, the hydrogen propellant enters the fuel element in a manner similar to that of the Particle Bed Reactor (PBR) fuel element.

  3. Boiler and Pressure Balls Monopropellant Thermal Rocket Engine

    Science.gov (United States)

    Greene, William D. (Inventor)

    2009-01-01

    The proposed technology is a rocket engine cycle utilizing as the propulsive fluid a low molecular weight, cryogenic fluid, typically liquid hydrogen, pressure driven, heated, and expelled through a nozzle to generate high velocity and high specific impulse discharge gas. The proposed technology feeds the propellant through the engine cycle without the use of a separate pressurization fluid and without the use of turbomachinery. Advantages of the proposed technology are found in those elements of state-of-the-art systems that it avoids. It does not require a separate pressurization fluid or a thick-walled primary propellant tank as is typically required for a classical pressure-fed system. Further, it does not require the acceptance of intrinsic reliability risks associated with the use of turbomachinery

  4. Design and analysis of a single stage to orbit nuclear thermal rocket reactor engine

    Energy Technology Data Exchange (ETDEWEB)

    Labib, Satira, E-mail: Satira.Labib@duke-energy.com; King, Jeffrey, E-mail: kingjc@mines.edu

    2015-06-15

    Graphical abstract: - Highlights: • Three NTR reactors are optimized for the single stage launch of 1–15 MT payloads. • The proposed rocket engines have specific impulses in excess of 700 s. • Reactivity and submersion criticality requirements are satisfied for each reactor. - Abstract: Recent advances in the development of high power density fuel materials have renewed interest in nuclear thermal rockets (NTRs) as a viable propulsion technology for future space exploration. This paper describes the design of three NTR reactor engines designed for the single stage to orbit launch of payloads from 1 to 15 metric tons. Thermal hydraulic and rocket engine analyses indicate that the proposed rocket engines are able to reach specific impulses in excess of 800 s. Neutronics analyses performed using MCNP5 demonstrate that the hot excess reactivity, shutdown margin, and submersion criticality requirements are satisfied for each NTR reactor. The reactors each consist of a 40 cm diameter core packed with hexagonal tungsten cermet fuel elements. The core is surrounded by radial and axial beryllium reflectors and eight boron carbide control drums. The 40 cm long reactor meets the submersion criticality requirements (a shutdown margin of at least $1 subcritical in all submersion scenarios) with no further modifications. The 80 and 120 cm long reactors include small amounts of gadolinium nitride as a spectral shift absorber to keep them subcritical upon submersion in seawater or wet sand following a launch abort.

  5. High Thermal Conductivity NARloy-Z-Diamond Composite Combustion Chamber Liner For Advanced Rocket Engines

    Science.gov (United States)

    Bhat, Biliyar N.; Ellis, David; Singh, Jogender

    2014-01-01

    Advanced high thermal conductivity materials research conducted at NASA Marshall Space Flight Center (MSFC) with state of the art combustion chamber liner material NARloy-Z showed that its thermal conductivity can be increased significantly by adding diamond particles and sintering it at high temperatures. For instance, NARloy-Z containing 40 vol. percent diamond particles, sintered at 975C to full density by using the Field assisted Sintering Technology (FAST) showed 69 percent higher thermal conductivity than baseline NARloy-Z. Furthermore, NARloy-Z-40vol. percent D is 30 percent lighter than NARloy-Z and hence the density normalized thermal conductivity is 140 percent better. These attributes will improve the performance and life of the advanced rocket engines significantly. By one estimate, increased thermal conductivity will directly translate into increased turbopump power up to 2X and increased chamber pressure for improved thrust and ISP, resulting in an expected 20 percent improvement in engine performance. Follow on research is now being conducted to demonstrate the benefits of this high thermal conductivity NARloy-Z-D composite for combustion chamber liner applications in advanced rocket engines. The work consists of a) Optimizing the chemistry and heat treatment for NARloy-Z-D composite, b) Developing design properties (thermal and mechanical) for the optimized NARloy-Z-D, c) Fabrication of net shape subscale combustion chamber liner, and d) Hot fire testing of the liner for performance. FAST is used for consolidating and sintering NARlo-Z-D. The subscale cylindrical liner with built in channels for coolant flow is also fabricated near net shape using the FAST process. The liner will be assembled into a test rig and hot fire tested in the MSFC test facility to determine performance. This paper describes the development of this novel high thermal conductivity NARloy-Z-D composite material, and the advanced net shape technology to fabricate the combustion

  6. A unique nuclear thermal rocket engine using a particle bed reactor

    Science.gov (United States)

    Culver, Donald W.; Dahl, Wayne B.; McIlwain, Melvin C.

    1992-01-01

    Aerojet Propulsion Division (APD) studied 75-klb thrust Nuclear Thermal Rocket Engines (NTRE) with particle bed reactors (PBR) for application to NASA's manned Mars mission and prepared a conceptual design description of a unique engine that best satisfied mission-defined propulsion requirements and customer criteria. This paper describes the selection of a sprint-type Mars transfer mission and its impact on propulsion system design and operation. It shows how our NTRE concept was developed from this information. The resulting, unusual engine design is short, lightweight, and capable of high specific impulse operation, all factors that decrease Earth to orbit launch costs. Many unusual features of the NTRE are discussed, including nozzle area ratio variation and nozzle closure for closed loop after cooling. Mission performance calculations reveal that other well known engine options do not support this mission.

  7. Gas core nuclear thermal rocket engine research and development in the former USSR

    International Nuclear Information System (INIS)

    Koehlinger, M.W.; Bennett, R.G.; Motloch, C.G.; Gurfink, M.M.

    1992-09-01

    Beginning in 1957 and continuing into the mid 1970s, the USSR conducted an extensive investigation into the use of both solid and gas core nuclear thermal rocket engines for space missions. During this time the scientific and engineering. problems associated with the development of a solid core engine were resolved. At the same time research was undertaken on a gas core engine, and some of the basic engineering problems associated with the concept were investigated. At the conclusion of the program, the basic principles of the solid core concept were established. However, a prototype solid core engine was not built because no established mission required such an engine. For the gas core concept, some of the basic physical processes involved were studied both theoretically and experimentally. However, no simple method of conducting proof-of-principle tests in a neutron flux was devised. This report focuses primarily on the development of the. gas core concept in the former USSR. A variety of gas core engine system parameters and designs are presented, along with a summary discussion of the basic physical principles and limitations involved in their design. The parallel development of the solid core concept is briefly described to provide an overall perspective of the magnitude of the nuclear thermal propulsion program and a technical comparison with the gas core concept

  8. To MARS and Beyond with Nuclear Power - Design Concept of Korea Advanced Nuclear Thermal Engine Rocket

    International Nuclear Information System (INIS)

    Nam, Seung Hyun; Chang, Soon Heung

    2013-01-01

    The President Park of ROK has also expressed support for space program promotion, praising the success of NARO as evidence of a positive outlook. These events hint a strong signal that ROK's space program will be accelerated by the national eager desire. In this national eager desire for space program, the policymakers and the aerospace engineers need to pay attention to the advanced nuclear technology of ROK that is set to a major world nuclear energy country, even exporting the technology. The space nuclear application is a very much attractive option because its energy density is the most enormous among available energy sources in space. This paper presents the design concept of Korea Advanced Nuclear Thermal Engine Rocket (KANuTER) that is one of the advanced nuclear thermal rocket engine developing in Korea Advanced Institute of Science and Technology (KAIST) for space application. Solar system exploration relying on CRs suffers from long trip time and high cost. In this regard, nuclear propulsion is a very attractive option for that because of higher performance and already demonstrated technology. Although ROK was a late entrant into elite global space club, its prospect as a space racer is very bright because of the national eager desire and its advanced technology. Especially it is greatly meaningful that ROK has potential capability to launch its nuclear technology into space as a global nuclear energy leader and a soaring space adventurer. In this regard, KANuTER will be a kind of bridgehead for Korean space nuclear application

  9. Thermohydraulic Design Analysis Modeling for Korea Advanced NUclear Thermal Engine Rocket for Space Application

    Energy Technology Data Exchange (ETDEWEB)

    Nam, Seung Hyun; Choi, Jae Young; Venneria, Paolo F.; Jeong, Yong Hoon; Chang, Soon Heung [KAIST, Daejeon (Korea, Republic of)

    2015-05-15

    Space exploration is a realistic and profitable goal for long-term humanity survival, although the harsh space environment imposes lots of severe challenges to space pioneers. To date, almost all space programs have relied upon Chemical Rockets (CRs) rating superior thrust level to transit from the Earth's surface to its orbit. However, CRs inherently have insurmountable barrier to carry out deep space missions beyond Earth's orbit due to its low propellant efficiency, and ensuing enormous propellant requirement and launch costs. Meanwhile, nuclear rockets typically offer at least two times the propellant efficiency of a CR and thus notably reduce the propellant demand. Particularly, a Nuclear Thermal Rocket (NTR) is a leading candidate for near-term manned missions to Mars and beyond because it satisfies a relatively high thrust as well as a high efficiency. The superior efficiency of NTRs is due to both high energy density of nuclear fuel and the low molecular weight propellant of Hydrogen (H{sub 2}) over the chemical reaction by-products. A NTR uses thermal energy released from a nuclear fission reactor to heat the H{sub 2} propellant and then exhausted the highly heated propellant through a propelling nozzle to produce thrust. A propellant efficiency parameter of rocket engines is specific impulse (I{sub s}p) which represents the ratio of the thrust over the propellant consumption rate. If the average exhaust H{sub 2} temperature of a NTR is around 3,000 K, the I{sub s}p can be achieved as high as 1,000 s as compared with only 450 - 500 s of the best CRs. For this reason, NTRs are favored for various space applications such as orbital tugs, lunar transports, and manned missions to Mars and beyond. The best known NTR development effort was conducted from 1955 to1974 under the ROVER and NERVA programs in the USA. These programs had successfully designed and tested many different reactors and engines. After these projects, the researches on NERVA derived

  10. Nuclear Rocket Engine Reactor

    CERN Document Server

    Lanin, Anatoly

    2013-01-01

    The development of a nuclear rocket engine reactor (NRER ) is presented in this book. The working capacity of an active zone NRER under mechanical and thermal load, intensive neutron fluxes, high energy generation (up to 30 MBT/l) in a working medium (hydrogen) at temperatures up to 3100 K is displayed. Design principles and bearing capacity of reactors area discussed on the basis of simulation experiments and test data of a prototype reactor. Property data of dense constructional, porous thermal insulating and fuel materials like carbide and uranium carbide compounds in the temperatures interval 300 - 3000 K are presented. Technological aspects of strength and thermal strength resistance of materials are considered. The design procedure of possible emergency processes in the NRER is developed and risks for their origination are evaluated. Prospects of the NRER development for pilotless space devices and piloted interplanetary ships are viewed.

  11. Coil-On-Plug Ignition for LOX/Methane Liquid Rocket Engines in Thermal Vacuum Environments

    Science.gov (United States)

    Melcher, John C.; Atwell, Matthew J.; Morehead, Robert L.; Hurlbert, Eric A.; Bugarin, Luz; Chaidez, Mariana

    2017-01-01

    A coil-on-plug ignition system has been developed and tested for Liquid Oxygen (LOX) / liquid methane rocket engines operating in thermal vacuum conditions. The igniters were developed and tested as part of the Integrated Cryogenic Propulsion Test Article (ICPTA), previously tested as part of the Project Morpheus test vehicle. The ICPTA uses an integrated, pressure-fed, cryogenic LOX/methane propulsion system including a reaction control system (RCS) and a main engine. The ICPTA was tested at NASA Glenn Research Center's Plum Brook Station in the Spacecraft Propulsion Research Facility (B-2) under vacuum and thermal vacuum conditions. In order to successfully demonstrate ignition reliability in the vacuum conditions and eliminate corona discharge issues, a coil-on-plug ignition system has been developed. The ICPTA uses spark-plug ignition for both the main engine igniter and the RCS. The coil-on-plug configuration eliminates the conventional high-voltage spark plug cable by combining the coil and the spark-plug into a single component. Prior to ICPTA testing at Plum Brook, component-level reaction control engine (RCE) and main engine igniter testing was conducted at NASA Johnson Space Center (JSC), which demonstrated successful hot-fire ignition using the coil-on-plug from sea-level ambient conditions down to 10(exp.-2) torr. Integrated vehicle hot-fire testing at JSC demonstrated electrical and command/data system performance. Lastly, Plum Brook testing demonstrated successful ignitions at simulated altitude conditions at 30 torr and cold thermal-vacuum conditions at 6 torr. The test campaign successfully proved that coil-on-plug technology will enable integrated LOX/methane propulsion systems in future spacecraft.

  12. Kinetic—a system code for analyzing nuclear thermal propulsion rocket engine transients

    Science.gov (United States)

    Schmidt, Eldon; Lazareth, Otto; Ludewig, Hans

    1993-01-01

    A system code suitable for analyzing Nuclear Thermal Propulsion (NTP) rocket engines is described in this paper. The code consists of a point reactor model and nodes to describe the fluid dynamics and heat transfer mechanism. Feedback from the fuel, coolant, moderator and reflector are allowed for, and the control of the reactor is by motion of controls element (drums or rods). The worth of the control element and feedback coefficients are predetermined. Separate models for the turbo-pump assembly (TPA) and nozzle are also included. The model to be described in this paper is specific for the Particle Bed Reactor (PBR). An illustrative problem is solved. This problem consists of a PBR operating in a blowdown mode.

  13. KINETIC: A system code for analyzing Nuclear thermal propulsion rocket engine transients

    Science.gov (United States)

    Schmidt, E.; Lazareth, O.; Ludewig, H.

    1993-07-01

    A system code suitable for analyzing Nuclear Thermal Propulsion (NTP) rocket engines is described in this paper. The code consists of a point reactor model and nodes to describe the fluid dynamics and heat transfer mechanism. Feedback from the fuel coolant, moderator and reflector are allowed for, and the control of the reactor is by motion of control elements (drums or rods). The worth of the control clement and feedback coefficients are predetermined. Separate models for the turbo-pump assembly (TPA) and nozzle are also included. The model to be described in this paper is specific for the Particle Bed Reactor (PBR). An illustrative problem is solved. This problem consists of a PBR operating in a blowdown mode.

  14. High Thermal Conductivity NARloy-Z-Diamond Composite Liner for Advanced Rocket Engines

    Science.gov (United States)

    Bhat, Biliyar; Greene, Sandra

    2015-01-01

    NARloy-Z (Cu-3Ag-0.5Zr) alloy is state-of-the-art combustion chamber liner material used in liquid propulsion engines such as the RS-68 and RS-25. The performance of future liquid propulsion systems can be improved significantly by increasing the heat transfer through the combustion chamber liner. Prior work1 done at NASA Marshall Space Flight Center (MSFC) has shown that the thermal conductivity of NARloy-Z alloy can be improved significantly by embedding high thermal conductivity diamond particles in the alloy matrix to form NARloy-Z-diamond composite (fig. 1). NARloy-Z-diamond composite containing 40vol% diamond showed 69% higher thermal conductivity than NARloy-Z. It is 24% lighter than NARloy-Z and hence the density normalized thermal conductivity is 120% better. These attributes will improve the performance and life of the advanced rocket engines significantly. The research work consists of (a) developing design properties (thermal and mechanical) of NARloy-Z-D composite, (b) fabrication of net shape subscale combustion chamber liner, and (c) hot-fire testing of the liner to test performance. Initially, NARloy-Z-D composite slabs were made using the Field Assisted Sintering Technology (FAST) for the purpose of determining design properties. In the next step, a cylindrical shape was fabricated to demonstrate feasibility (fig. 3). The liner consists of six cylinders which are sintered separately and then stacked and diffusion bonded to make the liner (fig. 4). The liner will be heat treated, finish-machined, and assembled into a combustion chamber and hot-fire tested in the MSFC test facility (TF 115) to determine perform.

  15. Conceptual Engine System Design for NERVA derived 66.7KN and 111.2KN Thrust Nuclear Thermal Rockets

    International Nuclear Information System (INIS)

    Fittje, James E.; Buehrle, Robert J.

    2006-01-01

    The Nuclear Thermal Rocket concept is being evaluated as an advanced propulsion concept for missions to the moon and Mars. A tremendous effort was undertaken during the 1960's and 1970's to develop and test NERVA derived Nuclear Thermal Rockets in the 111.2 KN to 1112 KN pound thrust class. NASA GRC is leveraging this past NTR investment in their vehicle concepts and mission analysis studies, and has been evaluating NERVA derived engines in the 66.7 KN to the 111.2 KN thrust range. The liquid hydrogen propellant feed system, including the turbopumps, is an essential component of the overall operation of this system. The NASA GRC team is evaluating numerous propellant feed system designs with both single and twin turbopumps. The Nuclear Engine System Simulation code is being exercised to analyze thermodynamic cycle points for these selected concepts. This paper will present propellant feed system concepts and the corresponding thermodynamic cycle points for 66.7 KN and 111.2 KN thrust NTR engine systems. A pump out condition for a twin turbopump concept will also be evaluated, and the NESS code will be assessed against the Small Nuclear Rocket Engine preliminary thermodynamic data

  16. Cryogenic rocket engine development at Delft aerospace rocket engineering

    NARCIS (Netherlands)

    Wink, J; Hermsen, R.; Huijsman, R; Akkermans, C.; Denies, L.; Barreiro, F.; Schutte, A.; Cervone, A.; Zandbergen, B.T.C.

    2016-01-01

    This paper describes the current developments regarding cryogenic rocket engine technology at Delft Aerospace Rocket Engineering (DARE). DARE is a student society based at Delft University of Technology with the goal of being the first student group in the world to launch a rocket into space. After

  17. Fabrication of High Thermal Conductivity NARloy-Z-Diamond Composite Combustion Chamber Liner for Advanced Rocket Engines

    Science.gov (United States)

    Bhat, Biliyar N.; Greene, Sandra E.; Singh, Jogender

    2016-01-01

    This paper describes the process development for fabricating a high thermal conductivity NARloy-Z-Diamond composite (NARloy-Z-D) combustion chamber liner for application in advanced rocket engines. The fabrication process is challenging and this paper presents some details of these challenges and approaches used to address them. Prior research conducted at NASA-MSFC and Penn State had shown that NARloy-Z-40%D composite material has significantly higher thermal conductivity than the state of the art NARloy-Z alloy. Furthermore, NARloy-Z-40 %D is much lighter than NARloy-Z. These attributes help to improve the performance of the advanced rocket engines. Increased thermal conductivity will directly translate into increased turbopump power, increased chamber pressure for improved thrust and specific impulse. Early work on NARloy-Z-D composites used the Field Assisted Sintering Technology (FAST, Ref. 1, 2) for fabricating discs. NARloy-Z-D composites containing 10, 20 and 40vol% of high thermal conductivity diamond powder were investigated. Thermal conductivity (TC) data. TC increased with increasing diamond content and showed 50% improvement over pure copper at 40vol% diamond. This composition was selected for fabricating the combustion chamber liner using the FAST technique.

  18. Liquid Rocket Engine Testing

    Science.gov (United States)

    2016-10-21

    and storable propellants • Liquid Oxygen (LOX) • RP-1 (Kerosene, very similar to JP-8) • Liquid Hydrogen • Liquid methane • Pressure = Performance in...booster rocket engines • 6000-10000 psia capabilities – Can use gaseous nitrogen, helium, or hydrogen to pressurize propellant tanks 9Distribution A...demonstrator of a Kerosene-LOX, 250,000 lbf, 3000 psi oxygen-rich staged combustion engine (ORSC) • AFRL’s Test Stand 2A recently completed a two-year

  19. Affordable Development and Demonstration of a Small Nuclear Thermal Rocket (NTR) Engine and Stage: How Small Is Big Enough?

    Science.gov (United States)

    Borowski, Stanley K.; Sefcik, Robert J.; Fittje, James E.; McCurdy, David R.; Qualls, Arthur L.; Schnitzler, Bruce G.; Werner, James E.; Weitzberg, Abraham; Joyner, Claude R.

    2016-01-01

    The Nuclear Thermal Rocket (NTR) derives its energy from fission of uranium-235 atoms contained within fuel elements that comprise the engine's reactor core. It generates high thrust and has a specific impulse potential of approximately 900 specific impulse - a 100 percent increase over today's best chemical rockets. The Nuclear Thermal Propulsion (NTP) project, funded by NASA's Advanced Exploration Systems (AES) program, includes five key task activities: (1) Recapture, demonstration, and validation of heritage graphite composite (GC) fuel (selected as the Lead Fuel option); (2) Engine Conceptual Design; (3) Operating Requirements Definition; (4) Identification of Affordable Options for Ground Testing; and (5) Formulation of an Affordable Development Strategy. During fiscal year (FY) 2014, a preliminary Design Development Test and Evaluation (DDT&E) plan and schedule for NTP development was outlined by the NASA Glenn Research Center (GRC), Department of Energy (DOE) and industry that involved significant system-level demonstration projects that included Ground Technology Demonstration (GTD) tests at the Nevada National Security Site (NNSS), followed by a Flight Technology Demonstration (FTD) mission. To reduce cost for the GTD tests and FTD mission, small NTR engines, in either the 7.5 or 16.5 kilopound-force thrust class, were considered. Both engine options used GC fuel and a common fuel element (FE) design. The small approximately 7.5 kilopound-force criticality-limited engine produces approximately157 thermal megawatts and its core is configured with parallel rows of hexagonal-shaped FEs and tie tubes (TTs) with a FE to TT ratio of approximately 1:1. The larger approximately 16.5 kilopound-force Small Nuclear Rocket Engine (SNRE), developed by Los Alamos National Laboratory (LANL) at the end of the Rover program, produces approximately 367 thermal megawatts and has a FE to TT ratio of approximately 2:1. Although both engines use a common 35-inch (approximately

  20. Thermal-structural analysis of regeneratively-cooled thrust chamber wall in reusable LOX/Methane rocket engines

    Directory of Open Access Journals (Sweden)

    Jiawen SONG

    2017-06-01

    Full Text Available To predict the thermal and structural responses of the thrust chamber wall under cyclic work, a 3-D fluid-structural coupling computational methodology is developed. The thermal and mechanical loads are determined by a validated 3-D finite volume fluid-thermal coupling computational method. With the specified loads, the nonlinear thermal-structural finite element analysis is applied to obtaining the 3-D thermal and structural responses. The Chaboche nonlinear kinematic hardening model calibrated by experimental data is adopted to predict the cyclic plastic behavior of the inner wall. The methodology is further applied to the thrust chamber of LOX/Methane rocket engines. The results show that both the maximum temperature at hot run phase and the maximum circumferential residual strain of the inner wall appear at the convergent part of the chamber. Structural analysis for multiple work cycles reveals that the failure of the inner wall may be controlled by the low-cycle fatigue when the Chaboche model parameter γ3 = 0, and the damage caused by the thermal-mechanical ratcheting of the inner wall cannot be ignored when γ3 > 0. The results of sensitivity analysis indicate that mechanical loads have a strong influence on the strains in the inner wall.

  1. The Thermal State Computational Research of the Low-Thrust Oxygen-Methane Gaseous-Propellant Rocket Engine in the Pulse Mode of Operation

    OpenAIRE

    O. A. Vorozheeva; D. A. Yagodnikov

    2014-01-01

    Currently promising development direction of space propulsion engineering is to use, as spacecraft controls, low-thrust rocket engines (RDTM) on clean fuels, such as oxygen-methane. Modern RDTM are characterized by a lack regenerative cooling and pulse mode of operation, during which there is accumulation of heat energy to lead to the high thermal stress of RDTM structural elements. To get an idea about the thermal state of its elements, which further will reduce the number of fire tests is t...

  2. Coil-On-Plug Ignition for Oxygen/Methane Liquid Rocket Engines in Thermal-Vacuum Environments

    Science.gov (United States)

    Melcher, John C.; Atwell, Matthew J.; Morehead, Robert L.; Hurlbert, Eric A.; Bugarin, Luz; Chaidez, Mariana

    2017-01-01

    A coil-on-plug ignition system has been developed and tested for Liquid Oxygen (LOX)/liquid methane (LCH4) rocket engines operating in thermal vacuum conditions. The igniters were developed and tested as part of the Integrated Cryogenic Propulsion Test Article (ICPTA), previously tested as part of the Project Morpheus test vehicle. The ICPTA uses an integrated, pressure-fed, cryogenic LOX/LCH4 propulsion system including a reaction control system (RCS) and a main engine. The ICPTA was tested at NASA Glenn Research Center's Plum Brook Station in the Spacecraft Propulsion Research Facility (B-2) under vacuum and thermal vacuum conditions. A coil-on-plug ignition system has been developed to successfully demonstrate ignition reliability at these conditions while preventing corona discharge issues. The ICPTA uses spark plug ignition for both the main engine igniter and the RCS. The coil-on-plug configuration eliminates the conventional high-voltage spark plug cable by combining the coil and the spark plug into a single component. Prior to ICPTA testing at Plum Brook, component-level reaction control engine (RCE) and main engine igniter testing was conducted at NASA Johnson Space Center (JSC), which demonstrated successful hot-fire ignition using the coil-on-plug from sea-level ambient conditions down to 10(exp -2) torr. Integrated vehicle hot-fire testing at JSC demonstrated electrical and command/data system performance. Lastly, hot-fire testing at Plum Brook demonstrated successful ignitions at simulated altitude conditions at 30 torr and cold thermal-vacuum conditions at 6 torr. The test campaign successfully proved that coil-on-plug technology will enable integrated LOX/LCH4 propulsion systems in future spacecraft.

  3. Thermal Hydraulics Design and Analysis Methodology for a Solid-Core Nuclear Thermal Rocket Engine Thrust Chamber

    Science.gov (United States)

    Wang, Ten-See; Canabal, Francisco; Chen, Yen-Sen; Cheng, Gary; Ito, Yasushi

    2013-01-01

    Nuclear thermal propulsion is a leading candidate for in-space propulsion for human Mars missions. This chapter describes a thermal hydraulics design and analysis methodology developed at the NASA Marshall Space Flight Center, in support of the nuclear thermal propulsion development effort. The objective of this campaign is to bridge the design methods in the Rover/NERVA era, with a modern computational fluid dynamics and heat transfer methodology, to predict thermal, fluid, and hydrogen environments of a hypothetical solid-core, nuclear thermal engine the Small Engine, designed in the 1960s. The computational methodology is based on an unstructured-grid, pressure-based, all speeds, chemically reacting, computational fluid dynamics and heat transfer platform, while formulations of flow and heat transfer through porous and solid media were implemented to describe those of hydrogen flow channels inside the solid24 core. Design analyses of a single flow element and the entire solid-core thrust chamber of the Small Engine were performed and the results are presented herein

  4. The Strutjet Rocket Based Combined Cycle Engine

    Science.gov (United States)

    Siebenhaar, A.; Bulman, M. J.; Bonnar, D. K.

    1998-01-01

    . RBCC engines exhibit a high potential for lowering the operating cost of launching payloads into orbit. Two sources of cost reductions can be identified. First, RBCC powered vehicles require only 20% takeoff thrust compared to conventional rockets, thereby lowering the thrust requirements and the replacement cost of the engines. Second, due to the higher structural and thermal margins achievable with RBCC engines coupled with a higher degree of subsystem redundance lower maintenance and operating cost are obtainable.

  5. Centrifugal pumps for rocket engines

    Science.gov (United States)

    Campbell, W. E.; Farquhar, J.

    1974-01-01

    The use of centrifugal pumps for rocket engines is described in terms of general requirements of operational and planned systems. Hydrodynamic and mechanical design considerations and techniques and test procedures are summarized. Some of the pump development experiences, in terms of both problems and solutions, are highlighted.

  6. Rocket Engine Altitude Simulation Technologies

    Science.gov (United States)

    Woods, Jody L.; Lansaw, John

    2010-01-01

    John C. Stennis Space Center is embarking on a very ambitious era in its rocket engine propulsion test history. The first new large rocket engine test stand to be built at Stennis Space Center in over 40 years is under construction. The new A3 Test Stand is designed to test very large (294,000 Ibf thrust) cryogenic propellant rocket engines at a simulated altitude of 100,000 feet. A3 Test Stand will have an engine testing chamber where the engine will be fired after the air in the chamber has been evacuated to a pressure at the simulated altitude of less than 0.16 PSIA. This will result in a very unique environment with extremely low pressures inside a very large chamber and ambient pressures outside this chamber. The test chamber is evacuated of air using a 2-stage diffuser / ejector system powered by 5000 lb/sec of steam produced by 27 chemical steam generators. This large amount of power and flow during an engine test will result in a significant acoustic and vibrational environment in and around A3 Test Stand.

  7. Development of Kabila rocket: A radioisotope heated thermionic plasma rocket engine

    OpenAIRE

    Mboyi, Kalomba; Ren, Junxue; Liu, Yu

    2015-01-01

    A new type of plasma rocket engine, the Kabila rocket, using a radioisotope heated thermionic heating chamber instead of a conventional combustion chamber or catalyst bed is introduced and it achieves specific impulses similar to the ones of conventional solid and bipropellant rockets. Curium-244 is chosen as a radioisotope heat source and a thermal reductive layer is also used to obtain precise thermionic emissions. The self-sufficiency principle is applied by simultaneously heating up the e...

  8. Rocket Engine Numerical Simulator (RENS)

    Science.gov (United States)

    Davidian, Kenneth O.

    1997-01-01

    Work is being done at three universities to help today's NASA engineers use the knowledge and experience of their Apolloera predecessors in designing liquid rocket engines. Ground-breaking work is being done in important subject areas to create a prototype of the most important functions for the Rocket Engine Numerical Simulator (RENS). The goal of RENS is to develop an interactive, realtime application that engineers can utilize for comprehensive preliminary propulsion system design functions. RENS will employ computer science and artificial intelligence research in knowledge acquisition, computer code parallelization and objectification, expert system architecture design, and object-oriented programming. In 1995, a 3year grant from the NASA Lewis Research Center was awarded to Dr. Douglas Moreman and Dr. John Dyer of Southern University at Baton Rouge, Louisiana, to begin acquiring knowledge in liquid rocket propulsion systems. Resources of the University of West Florida in Pensacola were enlisted to begin the process of enlisting knowledge from senior NASA engineers who are recognized experts in liquid rocket engine propulsion systems. Dr. John Coffey of the University of West Florida is utilizing his expertise in interviewing and concept mapping techniques to encode, classify, and integrate information obtained through personal interviews. The expertise extracted from the NASA engineers has been put into concept maps with supporting textual, audio, graphic, and video material. A fundamental concept map was delivered by the end of the first year of work and the development of maps containing increasing amounts of information is continuing. Find out more information about this work at the Southern University/University of West Florida. In 1996, the Southern University/University of West Florida team conducted a 4day group interview with a panel of five experts to discuss failures of the RL10 rocket engine in conjunction with the Centaur launch vehicle. The

  9. Measuring Model Rocket Engine Thrust Curves

    Science.gov (United States)

    Penn, Kim; Slaton, William V.

    2010-01-01

    This paper describes a method and setup to quickly and easily measure a model rocket engine's thrust curve using a computer data logger and force probe. Horst describes using Vernier's LabPro and force probe to measure the rocket engine's thrust curve; however, the method of attaching the rocket to the force probe is not discussed. We show how a…

  10. Metal Matrix Composites for Liquid Rocket Engines

    National Research Council Canada - National Science Library

    Shelley, J

    2001-01-01

    ...) technologies being developed for application to Liquid Rocket Engines (LIRE). Developments in LRE technology for the US Air Force are being tracked and planned through the Integrated High Payoff Rocket Propulsion Technologies Program (IHPRPT...

  11. Nuclear Thermal Rocket Simulation in NPSS

    Science.gov (United States)

    Belair, Michael L.; Sarmiento, Charles J.; Lavelle, Thomas M.

    2013-01-01

    Four nuclear thermal rocket (NTR) models have been created in the Numerical Propulsion System Simulation (NPSS) framework. The models are divided into two categories. One set is based upon the ZrC-graphite composite fuel element and tie tube-style reactor developed during the Nuclear Engine for Rocket Vehicle Application (NERVA) project in the late 1960s and early 1970s. The other reactor set is based upon a W-UO2 ceramic-metallic (CERMET) fuel element. Within each category, a small and a large thrust engine are modeled. The small engine models utilize RL-10 turbomachinery performance maps and have a thrust of approximately 33.4 kN (7,500 lbf ). The large engine models utilize scaled RL-60 turbomachinery performance maps and have a thrust of approximately 111.2 kN (25,000 lbf ). Power deposition profiles for each reactor were obtained from a detailed Monte Carlo N-Particle (MCNP5) model of the reactor cores. Performance factors such as thermodynamic state points, thrust, specific impulse, reactor power level, and maximum fuel temperature are analyzed for each engine design.

  12. Nuclear Thermal Rocket Element Environmental Simulator (NTREES)

    Science.gov (United States)

    Emrich, William J.

    2008-01-01

    To support a potential future development of a nuclear thermal rocket engine, a state-of-the-art non nuclear experimental test setup has been constructed to evaluate the performance characteristics of candidate fuel element materials and geometries in representative environments. The test device simulates the environmental conditions (minus the radiation) to which nuclear rocket fuel components could be subjected during reactor operation. Test articles mounted in the simulator are inductively heated in such a manner as to accurately reproduce the temperatures and heat fluxes normally expected to occur as a result of nuclear fission while at the same time being exposed to flowing hydrogen. This project is referred to as the Nuclear Thermal Rocket Element Environment Simulator or NTREES. The NTREES device is located at the Marshall Space flight Center in a laboratory which has been modified to accommodate the high powers required to heat the test articles to the required temperatures and to handle the gaseous hydrogen flow required for the tests. Other modifications to the laboratory include the installation of a nitrogen gas supply system and a cooling water supply system. During the design and construction of the facility, every effort was made to comply with all pertinent regulations to provide assurance that the facility could be operated in a safe and efficient manner. The NTREES system can currently supply up to 50 kW of inductive heating to the fuel test articles, although the facility has been sized to eventually allow test article heating levels of up to several megawatts.

  13. Pressure And Thermal Modeling Of Rocket Launches

    Science.gov (United States)

    Smith, Sheldon D.; Myruski, Brian L.; Farmer, Richard C.; Freeman, Jon A.

    1995-01-01

    Report presents mathematical model for use in designing rocket-launching stand. Predicts pressure and thermal environment, as well as thermal responses of structures to impinging rocket-exhaust plumes. Enables relatively inexperienced analyst to determine time-varying distributions and absolute levels of pressure and heat loads on structures.

  14. High Thrust & High ISP Nuclear Thermal Rocket (NTR) Grooved Ring Fuel Element (GRFE)

    Data.gov (United States)

    National Aeronautics and Space Administration — Missions to Mars will benefit from propulsion systems with performance levels exceeding that of today's best chemical engines. Nuclear Thermal Rocket (NTR)...

  15. Scale-Up of GRCop: From Laboratory to Rocket Engines

    Science.gov (United States)

    Ellis, David L.

    2016-01-01

    GRCop is a high temperature, high thermal conductivity copper-based series of alloys designed primarily for use in regeneratively cooled rocket engine liners. It began with laboratory-level production of a few grams of ribbon produced by chill block melt spinning and has grown to commercial-scale production of large-scale rocket engine liners. Along the way, a variety of methods of consolidating and working the alloy were examined, a database of properties was developed and a variety of commercial and government applications were considered. This talk will briefly address the basic material properties used for selection of compositions to scale up, the methods used to go from simple ribbon to rocket engines, the need to develop a suitable database, and the issues related to getting the alloy into a rocket engine or other application.

  16. Program For Optimization Of Nuclear Rocket Engines

    Science.gov (United States)

    Plebuch, R. K.; Mcdougall, J. K.; Ridolphi, F.; Walton, James T.

    1994-01-01

    NOP is versatile digital-computer program devoloped for parametric analysis of beryllium-reflected, graphite-moderated nuclear rocket engines. Facilitates analysis of performance of engine with respect to such considerations as specific impulse, engine power, type of engine cycle, and engine-design constraints arising from complications of fuel loading and internal gradients of temperature. Predicts minimum weight for specified performance.

  17. Alternate Propellant Thermal Rocket, Phase I

    Data.gov (United States)

    National Aeronautics and Space Administration — The Alternate Propellant Thermal Rocket (APTR) is a novel concept for propulsion of space exploration or orbit transfer vehicles. APTR propulsion is provided by...

  18. Turbopump options for nuclear thermal rockets

    International Nuclear Information System (INIS)

    Bissell, W.R.; Gunn, S.V.

    1992-07-01

    Several turbopump options for delivering liquid nitrogen to nuclear thermal rocket (NTR) engines were evaluated and compared. Axial and centrifugal flow pumps were optimized, with and without boost pumps, utilizing current design criteria within the latest turbopump technology limits. Two possible NTR design points were used, a modest pump pressure rise of 1,743 psia and a relatively higher pump pressure rise of 4,480 psia. Both engines utilized the expander cycle to maximize engine performance for the long duration mission. Pump suction performance was evaluated. Turbopumps with conventional cavitating inducers were compared with zero NPSH (saturated liquid in the tanks) pumps over a range of tank saturation pressures, with and without boost pumps. Results indicate that zero NSPH pumps at high tank vapor pressures, 60 psia, are very similar to those with the finite NPSHs. At low vapor pressures efficiencies fall and turbine pressure ratios increase leading to decreased engine chamber pressures and or increased pump pressure discharges and attendant high-pressure component weights. It may be concluded that zero tank NSPH capabilities can be obtained with little penalty to the engine systems but boost pumps are needed if tank vapor pressure drops below 30 psia. Axial pumps have slight advantages in weight and chamber pressure capability while centrifugal pumps have a greater operating range. 10 refs

  19. Vacuum plasma spray applications on liquid fuel rocket engines

    Science.gov (United States)

    McKechnie, T. N.; Zimmerman, F. R.; Bryant, M. A.

    1992-07-01

    The vacuum plasma spray process (VPS) has been developed by NASA and Rocketdyne for a variety of applications on liquid fuel rocket engines, including the Space Shuttle Main Engine. These applications encompass thermal barrier coatings which are thermal shock resistant for turbopump blades and nozzles; bond coatings for cryogenic titanium components; wear resistant coatings and materials; high conductivity copper, NaRloy-Z, combustion chamber liners, and structural nickel base material, Inconel 718, for nozzle and combustion chamber support jackets.

  20. Nitrous Oxide/Paraffin Hybrid Rocket Engines

    Science.gov (United States)

    Zubrin, Robert; Snyder, Gary

    2010-01-01

    Nitrous oxide/paraffin (N2OP) hybrid rocket engines have been invented as alternatives to other rocket engines especially those that burn granular, rubbery solid fuels consisting largely of hydroxyl- terminated polybutadiene (HTPB). Originally intended for use in launching spacecraft, these engines would also be suitable for terrestrial use in rocket-assisted takeoff of small airplanes. The main novel features of these engines are (1) the use of reinforced paraffin as the fuel and (2) the use of nitrous oxide as the oxidizer. Hybrid (solid-fuel/fluid-oxidizer) rocket engines offer advantages of safety and simplicity over fluid-bipropellant (fluid-fuel/fluid-oxidizer) rocket en - gines, but the thrusts of HTPB-based hybrid rocket engines are limited by the low regression rates of the fuel grains. Paraffin used as a solid fuel has a regression rate about 4 times that of HTPB, but pure paraffin fuel grains soften when heated; hence, paraffin fuel grains can, potentially, slump during firing. In a hybrid engine of the present type, the paraffin is molded into a 3-volume-percent graphite sponge or similar carbon matrix, which supports the paraffin against slumping during firing. In addition, because the carbon matrix material burns along with the paraffin, engine performance is not appreciably degraded by use of the matrix.

  1. Yes--This is Rocket Science: MMCs for Liquid Rocket Engines

    National Research Council Canada - National Science Library

    Shelley, J

    2001-01-01

    The Air Force's Integrated High-Payoff Rocket Propulsion Technologies (IHPRPT) Program has established aggressive goals for both improved performance and reduced cost of rocket engines and components...

  2. The Thermal State Computational Research of the Low-Thrust Oxygen-Methane Gaseous-Propellant Rocket Engine in the Pulse Mode of Operation

    Directory of Open Access Journals (Sweden)

    O. A. Vorozheeva

    2014-01-01

    Full Text Available Currently promising development direction of space propulsion engineering is to use, as spacecraft controls, low-thrust rocket engines (RDTM on clean fuels, such as oxygen-methane. Modern RDTM are characterized by a lack regenerative cooling and pulse mode of operation, during which there is accumulation of heat energy to lead to the high thermal stress of RDTM structural elements. To get an idea about the thermal state of its elements, which further will reduce the number of fire tests is therefore necessary in the development phase of a new product. Accordingly, the aim of this work is the mathematical modeling and computational study of the thermal state of gaseous oxygen-methane propellant RDMT operating in pulse mode.In this paper we consider a model RDTM working on gaseous propellants oxygen-methane in pulse mode.To calculate the temperature field of the chamber wall of model RDMT under consideration is used the mathematical model of non-stationary heat conduction in a two-dimensional axisymmetric formulation that takes into account both the axial heat leakages and the nonstationary processes occurring inside the chamber during pulse operation of RDMT.As a result of numerical study of the thermal state of model RDMT, are obtained the temperature fields during engine operation based on convective, conductive, and radiative mechanisms of heat transfer from the combustion products to the wall.It is shown that the elements of flanges of combustion chamber of model RDMT act as heat sinks structural elements. Temperatures in the wall of the combustion chamber during the engine mode of operation are considered relatively low.Raised temperatures can also occur in the mixing head in the feeding area of the oxidant into the combustion chamber.During engine operation in the area forming the critical section, there is an intensive heating of a wall, which can result in its melting, which in turn will increase the minimum nozzle throat area and hence

  3. Development of Kabila rocket: A radioisotope heated thermionic plasma rocket engine

    Directory of Open Access Journals (Sweden)

    Kalomba Mboyi

    2015-04-01

    Full Text Available A new type of plasma rocket engine, the Kabila rocket, using a radioisotope heated thermionic heating chamber instead of a conventional combustion chamber or catalyst bed is introduced and it achieves specific impulses similar to the ones of conventional solid and bipropellant rockets. Curium-244 is chosen as a radioisotope heat source and a thermal reductive layer is also used to obtain precise thermionic emissions. The self-sufficiency principle is applied by simultaneously heating up the emitting material with the radioisotope decay heat and by powering the different valves of the plasma rocket engine with the same radioisotope decay heat using a radioisotope thermoelectric generator. This rocket engine is then benchmarked against a 1 N hydrazine thruster configuration operated on one of the Pleiades-HR-1 constellation spacecraft. A maximal specific impulse and power saving of respectively 529 s and 32% are achieved with helium as propellant. Its advantages are its power saving capability, high specific impulses and simultaneous ease of storage and restart. It can however be extremely voluminous and potentially hazardous. The Kabila rocket is found to bring great benefits to the existing spacecraft and further research should optimize its geometric characteristics and investigate the physical principals of its operation.

  4. Additive Manufacturing for Affordable Rocket Engines

    Science.gov (United States)

    West, Brian; Robertson, Elizabeth; Osborne, Robin; Calvert, Marty

    2016-01-01

    Additive manufacturing (also known as 3D printing) technology has the potential to drastically reduce costs and lead times associated with the development of complex liquid rocket engine systems. NASA is using 3D printing to manufacture rocket engine components including augmented spark igniters, injectors, turbopumps, and valves. NASA is advancing the process to certify these components for flight. Success Story: MSFC has been developing rocket 3D-printing technology using the Selective Laser Melting (SLM) process. Over the last several years, NASA has built and tested several injectors and combustion chambers. Recently, MSFC has 3D printed an augmented spark igniter for potential use the RS-25 engines that will be used on the Space Launch System. The new design is expected to reduce the cost of the igniter by a factor of four. MSFC has also 3D printed and tested a liquid hydrogen turbopump for potential use on an Upper Stage Engine. Additive manufacturing of the turbopump resulted in a 45% part count reduction. To understanding how the 3D printed parts perform and to certify them for flight, MSFC built a breadboard liquid rocket engine using additive manufactured components including injectors, turbomachinery, and valves. The liquid rocket engine was tested seven times in 2016 using liquid oxygen and liquid hydrogen. In addition to exposing the hardware to harsh environments, engineers learned to design for the new manufacturing technique, taking advantage of its capabilities and gaining awareness of its limitations. Benefit: The 3D-printing technology promises reduced cost and schedule for rocket engines. Cost is a function of complexity, and the most complicated features provide the largest opportunities for cost reductions. This is especially true where brazes or welds can be eliminated. The drastic reduction in part count achievable with 3D printing creates a waterfall effect that reduces the number of processes and drawings, decreases the amount of touch

  5. Solar-thermal engine testing

    Science.gov (United States)

    Tucker, Stephen; Salvail, Pat

    2002-01-01

    A solar-thermal engine serves as a high-temperature solar-radiation absorber, heat exchanger, and rocket nozzle, collecting concentrated solar radiation into an absorber cavity and transferring this energy to a propellant as heat. Propellant gas can be heated to temperatures approaching 4,500 °F and expanded in a rocket nozzle, creating low thrust with a high specific impulse (Isp). The Shooting Star Experiment (SSE) solar-thermal engine is made of 100 percent chemically vapor deposited (CVD) rhenium. The engine ``module'' consists of an engine assembly, propellant feedline, engine support structure, thermal insulation, and instrumentation. Engine thermal performance tests consist of a series of high-temperature thermal cycles intended to characterize the propulsive performance of the engines and the thermal effectiveness of the engine support structure and insulation system. A silicone-carbide electrical resistance heater, placed inside the inner shell, substitutes for solar radiation and heats the engine. Although the preferred propellant is hydrogen, the propellant used in these tests is gaseous nitrogen. Because rhenium oxidizes at elevated temperatures, the tests are performed in a vacuum chamber. Test data will include transient and steady state temperatures on selected engine surfaces, propellant pressures and flow rates, and engine thrust levels. The engine propellant-feed system is designed to supply GN2 to the engine at a constant inlet pressure of 60 psia, producing a near-constant thrust of 1.0 lb. Gaseous hydrogen will be used in subsequent tests. The propellant flow rate decreases with increasing propellant temperature, while maintaining constant thrust, increasing engine Isp. In conjunction with analytical models of the heat exchanger, the temperature data will provide insight into the effectiveness of the insulation system, the structural support system, and the overall engine performance. These tests also provide experience on operational aspects

  6. Hydrocarbon Rocket Engine Plume Imaging with Laser Induced Incandescence Project

    Data.gov (United States)

    National Aeronautics and Space Administration — NASA/ Marshall Space Flight Center (MSFC) needs sensors that can be operated on rocket engine plume environments to improve NASA/SSC rocket engine performance. In...

  7. Structurally compliant rocket engine combustion chamber: Experimental and analytical validation

    Science.gov (United States)

    Jankovsky, Robert S.; Arya, Vinod K.; Kazaroff, John M.; Halford, Gary R.

    1994-03-01

    A new, structurally compliant rocket engine combustion chamber design has been validated through analysis and experiment. Subscale, tubular channel chambers have been cyclically tested and analytically evaluated. Cyclic lives were determined to have a potential for 1000 percent increase over those of rectangular channel designs, the current state of the art. Greater structural compliance in the circumferential direction gave rise to lower thermal strains during hot firing, resulting in lower thermal strain ratcheting and longer predicted fatigue lives. Thermal, structural, and durability analyses of the combustion chamber design, involving cyclic temperatures, strains, and low-cycle fatigue lives, have corroborated the experimental observations.

  8. Thermal Barrier Coatings on Copper Substrates for Rocket Applications

    Science.gov (United States)

    Schloesser, Jana; Fedorova, Tatiana; Bäker, Martin; Rösler, Joachim

    Currently a new generation of relaunchable space transportation system using liquid hydrogen/ liquid oxygen rocket engines is under development. The inner combustion chamber is exposed to extreme thermal loads and environmental attack during starts. To prevent failure of the cooling channels, a thermal barrier coating to provide thermal and oxidation protection could be applied. Thermal barrier coatings are state of the art for gas turbines and this concept should be transferred to copper substrates in rocket engine applications. The thermomechanical loading conditions are quite different from the gas turbine applications as heat fluxes and temperature gradients are much higher while overall service time is much shorter. As a start for optimization of a suitable coating, a material system known for gas turbines is employed. In this work a thermal barrier coating system is applied by atmospheric plasma spraying to the copper-based high strength alloy Cu-1%Cr-0.3%Zr. The bond coat consists of a NiCrAlY alloy, while partially stabilized zirconia is used as a top coat. Spraying parameter optimization for the new substrate is described. The reached coating system is tested in thermal cycling experiments, where no failure of the coating could be detected. In oxidation experiments good environmental protection of the coating is shown.

  9. AJ26 rocket engine testing news briefing

    Science.gov (United States)

    2010-01-01

    NASA's John C. Stennis Space Center Director Gene Goldman (center) stands in front of a 'pathfinder' rocket engine with Orbital Sciences Corp. President and Chief Operating Officer J.R. Thompson (left) and Aerojet President Scott Seymour during a Feb. 24 news briefing at the south Mississippi facility. The leaders appeared together to announce a partnership for testing Aerojet AJ26 rocket engines at Stennis. The engines will be used to power Orbital's Taurus II space vehicles to provide commercial cargo transportation missions to the International Space Station for NASA. During the event, the Stennis partnership with Orbital was cited as an example of the new direction of NASA to work with commercial interests for space travel and transport.

  10. Closed-cycle liquid propellant rocket engines

    Science.gov (United States)

    Kuznetsov, N. D.

    1993-06-01

    The paper presents experience gained by SSSPE TRUD in development of NK-33, NK-43, NK-39, and NK-31 liquid propellant rocket engines, which are reusable, closed-cycle type, working on liquid oxygen and kerosene. Results are presented showing the engine structure efficiency, configuration rationality, and optimal thrust values which provide the following specific parameters: specific vacuum impulses in the range 331-353 s (for NK-33 and NK-31 engines, respectively) and specific weight of about 8 kg/tf (NK-33 and NK-43 engines). The problems which occurred during engine development and the study of the main components of these engines are discussed. The important technical data, materials, methodology, and bench development data are presented for the gas generator, turbopump assembly, combustion chamber and full-scale engines.

  11. Software for Collaborative Engineering of Launch Rockets

    Science.gov (United States)

    Stanley, Thomas Troy

    2003-01-01

    The Rocket Evaluation and Cost Integration for Propulsion and Engineering software enables collaborative computing with automated exchange of information in the design and analysis of launch rockets and other complex systems. RECIPE can interact with and incorporate a variety of programs, including legacy codes, that model aspects of a system from the perspectives of different technological disciplines (e.g., aerodynamics, structures, propulsion, trajectory, aeroheating, controls, and operations) and that are used by different engineers on different computers running different operating systems. RECIPE consists mainly of (1) ISCRM a file-transfer subprogram that makes it possible for legacy codes executed in their original operating systems on their original computers to exchange data and (2) CONES an easy-to-use filewrapper subprogram that enables the integration of legacy codes. RECIPE provides a tightly integrated conceptual framework that emphasizes connectivity among the programs used by the collaborators, linking these programs in a manner that provides some configuration control while facilitating collaborative engineering tradeoff studies, including design to cost studies. In comparison with prior collaborative-engineering schemes, one based on the use of RECIPE enables fewer engineers to do more in less time.

  12. Oxidation of Copper Alloy Candidates for Rocket Engine Applications

    Science.gov (United States)

    Ogbuji, Linus U. Thomas; Humphrey, Donald L.

    2002-01-01

    The gateway to affordable and reliable space transportation in the near future remains long-lived rocket-based propulsion systems; and because of their high conductivities, copper alloys remain the best materials for lining rocket engines and dissipating their enormous thermal loads. However, Cu and its alloys are prone to oxidative degradation -- especially via the ratcheting phenomenon of blanching, which occurs in situations where the local ambient can oscillate between oxidation and reduction, as it does in a H2/02- fuelled rocket engine. Accordingly, resistance to blanching degradation is one of the key requirements for the next generation of reusable launch vehicle (RLV) liner materials. Candidate copper alloys have been studied with a view to comparing their oxidation behavior, and hence resistance to blanching, in ambients corresponding to conditions expected in rocket engine service. These candidate materials include GRCop-84 and GRCop-42 (Cu - Cr-8 - Nb-4 and Cu - Cr-4 - Nb-2 respectively); NARloy-Z (Cu-3%Ag-0.5%Y), and GlidCop (Cu-O.l5%Al2O3 ODS alloy); they represent different approaches to improving the mechanical properties of Cu without incurring a large drop in thermal conductivity. Pure Cu (OFHC-Cu) was included in the study to provide a baseline for comparison. The samples were exposed for 10 hours in the TGA to oxygen partial pressures ranging from 322 ppm to 1.0 atmosphere and at temperatures of up to 700 C, and examined by SEM-EDS and other techniques of metallography. This paper will summarize the results obtained.

  13. Cycle Trades for Nuclear Thermal Rocket Propulsion Systems

    Science.gov (United States)

    White, C.; Guidos, M.; Greene, W.

    2003-01-01

    Nuclear fission has been used as a reliable source for utility power in the United States for decades. Even in the 1940's, long before the United States had a viable space program, the theoretical benefits of nuclear power as applied to space travel were being explored. These benefits include long-life operation and high performance, particularly in the form of vehicle power density, enabling longer-lasting space missions. The configurations for nuclear rocket systems and chemical rocket systems are similar except that a nuclear rocket utilizes a fission reactor as its heat source. This thermal energy can be utilized directly to heat propellants that are then accelerated through a nozzle to generate thrust or it can be used as part of an electricity generation system. The former approach is Nuclear Thermal Propulsion (NTP) and the latter is Nuclear Electric Propulsion (NEP), which is then used to power thruster technologies such as ion thrusters. This paper will explore a number of indirect-NTP engine cycle configurations using assumed performance constraints and requirements, discuss the advantages and disadvantages of each cycle configuration, and present preliminary performance and size results. This paper is intended to lay the groundwork for future efforts in the development of a practical NTP system or a combined NTP/NEP hybrid system.

  14. Rocket Engine Innovations Advance Clean Energy

    Science.gov (United States)

    2012-01-01

    During launch countdown, at approximately T-7 seconds, the Space Shuttle Main Engines (SSMEs) roar to life. When the controllers indicate normal operation, the solid rocket boosters ignite and the shuttle blasts off. Initially, the SSMEs throttle down to reduce stress during the period of maximum dynamic pressure, but soon after, they throttle up to propel the orbiter to 17,500 miles per hour. In just under 9 minutes, the three SSMEs burn over 1.6 million pounds of propellant, and temperatures inside the main combustion chamber reach 6,000 F. To cool the engines, liquid hydrogen circulates through miles of tubing at -423 F. From 1981to 2011, the Space Shuttle fleet carried crew and cargo into orbit to perform a myriad of unprecedented tasks. After 30 years and 135 missions, the feat of engineering known as the SSME boasted a 100-percent flight success rate.

  15. Rocket engine control and monitoring expert system

    Science.gov (United States)

    Ali, Moonis; Crawford, Roger

    1988-01-01

    This paper focuses on the application of expert systems technology to the automatic detection, verification and correction of anomalous rocket engine operations through interfacing with an intelligent adaptive control system. The design of a reliable and intelligent propulsion control and monitoring system is outlined which includes the architecture of an Integrated Expert System (IES) serving as the core component. The IES functions include automatic knowledge acquisition, integrated knowledge base, and fault diagnosis and prediction methodology. The results of fault analysis and diagnostic techniques are presented for an example fault in the SSME main combustion chamber injectors.

  16. Numerical investigations of hybrid rocket engines

    Science.gov (United States)

    Betelin, V. B.; Kushnirenko, A. G.; Smirnov, N. N.; Nikitin, V. F.; Tyurenkova, V. V.; Stamov, L. I.

    2018-03-01

    Paper presents the results of numerical studies of hybrid rocket engines operating cycle including unsteady-state transition stage. A mathematical model is developed accounting for the peculiarities of diffusion combustion of fuel in the flow of oxidant, which is composed of oxygen-nitrogen mixture. Three dimensional unsteady-state simulations of chemically reacting gas mixture above thermochemically destructing surface are performed. The results show that the diffusion combustion brings to strongly non-uniform fuel mass regression rate in the flow direction. Diffusive deceleration of chemical reaction brings to the decrease of fuel regression rate in the longitudinal direction.

  17. Liquid fuel injection elements for rocket engines

    Science.gov (United States)

    Cox, George B., Jr. (Inventor)

    1993-01-01

    Thrust chambers for liquid propellant rocket engines include three principal components. One of these components is an injector which contains a plurality of injection elements to meter the flow of propellants at a predetermined rate, and fuel to oxidizer mixture ratio, to introduce the mixture into the combustion chamber, and to cause them to be atomized within the combustion chamber so that even combustion takes place. Evolving from these injectors are tube injectors. These tube injectors have injection elements for injecting the oxidizer into the combustion chamber. The oxidizer and fuel must be metered at predetermined rates and mixture ratios in order to mix them within the combustion chamber so that combustion takes place smoothly and completely. Hence tube injectors are subject to improvement. An injection element for a liquid propellant rocket engine of the bipropellant type is provided which includes tangential fuel metering orifices, and a plurality of oxidizer tube injection elements whose injection tubes are also provided with tangential oxidizer entry slots and internal reed valves.

  18. Nuclear thermal rocket propulsion application to Mars missions

    International Nuclear Information System (INIS)

    Emrich, W.J. Jr.; Young, A.C.; Mulqueen, J.A.

    1991-01-01

    Options for vehicle configurations are reviewed in which nuclear thermal rocket (NTR) propulsion is used for a reference mission to Mars. The scenario assumes an opposition-class Mars transfer trajectory, a 435-day mission, and the use of a single nuclear engine with 75,000 lbs of thrust. Engine parameters are examined by calculating mission variables for a range of specific impulses and thrust/weight ratios. The reference mission is found to have optimal values of 925 s for the specific impulse and thrust/weight ratios of 4.0 and 0.06 for the engine and total stage ratios respectively. When the engine thrust/weight ratio is at least 4/1 the most critical engine parameter is engine specific impulse for reducing overall stage weight. In the context of this trans-Mars three-burn maneuver the NTR engine with an expander engine cycle is considered a more effective alternative than chemical/aerobrake and other propulsion options

  19. Enrichment Zoning Options for the Small Nuclear Rocket Engine (SNRE)

    Energy Technology Data Exchange (ETDEWEB)

    Bruce G. Schnitzler; Stanley K. Borowski

    2010-07-01

    Advancement of U.S. scientific, security, and economic interests through a robust space exploration program requires high performance propulsion systems to support a variety of robotic and crewed missions beyond low Earth orbit. In NASA’s recent Mars Design Reference Architecture (DRA) 5.0 study (NASA-SP-2009-566, July 2009), nuclear thermal propulsion (NTP) was again selected over chemical propulsion as the preferred in-space transportation system option because of its high thrust and high specific impulse (-900 s) capability, increased tolerance to payload mass growth and architecture changes, and lower total initial mass in low Earth orbit. An extensive nuclear thermal rocket technology development effort was conducted from 1955-1973 under the Rover/NERVA Program. The Small Nuclear Rocket Engine (SNRE) was the last engine design studied by the Los Alamos National Laboratory during the program. At the time, this engine was a state-of-the-art design incorporating lessons learned from the very successful technology development program. Past activities at the NASA Glenn Research Center have included development of highly detailed MCNP Monte Carlo transport models of the SNRE and other small engine designs. Preliminary core configurations typically employ fuel elements with fixed fuel composition and fissile material enrichment. Uniform fuel loadings result in undesirable radial power and temperature profiles in the engines. Engine performance can be improved by some combination of propellant flow control at the fuel element level and by varying the fuel composition. Enrichment zoning at the fuel element level with lower enrichments in the higher power elements at the core center and on the core periphery is particularly effective. Power flattening by enrichment zoning typically results in more uniform propellant exit temperatures and improved engine performance. For the SNRE, element enrichment zoning provided very flat radial power profiles with 551 of the 564

  20. Investigation of Cooling Water Injection into Supersonic Rocket Engine Exhaust

    Science.gov (United States)

    Jones, Hansen; Jeansonne, Christopher; Menon, Shyam

    2017-11-01

    Water spray cooling of the exhaust plume from a rocket undergoing static testing is critical in preventing thermal wear of the test stand structure, and suppressing the acoustic noise signature. A scaled test facility has been developed that utilizes non-intrusive diagnostic techniques including Focusing Color Schlieren (FCS) and Phase Doppler Particle Anemometry (PDPA) to examine the interaction of a pressure-fed water jet with a supersonic flow of compressed air. FCS is used to visually assess the interaction of the water jet with the strong density gradients in the supersonic air flow. PDPA is used in conjunction to gain statistical information regarding water droplet size and velocity as the jet is broken up. Measurement results, along with numerical simulations and jet penetration models are used to explain the observed phenomena. Following the cold flow testing campaign a scaled hybrid rocket engine will be constructed to continue tests in a combusting flow environment similar to that generated by the rocket engines tested at NASA facilities. LaSPACE.

  1. Investigation of the cooling film distribution in liquid rocket engine

    Directory of Open Access Journals (Sweden)

    Luís Antonio Silva

    2011-05-01

    Full Text Available This study presents the results of the investigation of a cooling method widely used in the combustion chambers, which is called cooling film, and it is applied to a liquid rocket engine that uses as propellants liquid oxygen and kerosene. Starting from an engine cooling, whose film is formed through the fuel spray guns positioned on the periphery of the injection system, the film was experimentally examined, it is formed by liquid that seeped through the inner wall of the combustion chamber. The parameter used for validation and refinement of the theoretical penetration of the film was cooling, as this parameter is of paramount importance to obtain an efficient thermal protection inside the combustion chamber. Cold tests confirmed a penetrating cold enough cooling of the film for the length of the combustion chamber of the studied engine.

  2. Developments in REDES: The Rocket Engine Design Expert System

    Science.gov (United States)

    Davidian, Kenneth O.

    1990-01-01

    The Rocket Engine Design Expert System (REDES) was developed at NASA-Lewis to collect, automate, and perpetuate the existing expertise of performing a comprehensive rocket engine analysis and design. Currently, REDES uses the rigorous JANNAF methodology to analyze the performance of the thrust chamber and perform computational studies of liquid rocket engine problems. The following computer codes were included in REDES: a gas properties program named GASP; a nozzle design program named RAO; a regenerative cooling channel performance evaluation code named RTE; and the JANNAF standard liquid rocket engine performance prediction code TDK (including performance evaluation modules ODE, ODK, TDE, TDK, and BLM). Computational analyses are being conducted by REDES to provide solutions to liquid rocket engine thrust chamber problems. REDES was built in the Knowledge Engineering Environment (KEE) expert system shell and runs on a Sun 4/110 computer.

  3. Advances for laser ignition of internal combustion and rocket engines

    International Nuclear Information System (INIS)

    Schwarz, E.

    2011-01-01

    The scope of the PhD thesis presented here is the investigation of theoretical and practical aspects of laser-induced spark ignition and laser thermal ignition. Laser ignition systems are currently undergoing a rapidly development with growing intensity involving more and more research groups who mainly concentrate on the field of car and large combustion engines. This research is primarily driven by the engagement to meet the increasingly strict emission limits and by the intention to use the limited energy reserves more efficiently. For internal combustion engines, laser plasma-induced ignition will allow to combine the goals for legally required reductions of pollutant emissions and higher engine efficiencies. Also for rocket engines laser ignition turns out to be very attractive. A highly reliable ignition system like laser ignition would represent an option for introducing non-toxic propellants in order to replace highly toxic and carcinogenic hydrazine-based propellants commonly used in launch vehicle upper stages and satellites. The most important results on laser ignition and laser plasma generation, accomplished by the author and, in some respects, enriched by cooperation with colleagues are presented in the following. The emphasis of this thesis is placed on the following issues: - Two-color effects on laser plasma generation - Theoretical considerations about the focal volume concerning plasma generation - Plasma transmission experiments - Ignition experiments on laser-induced ignition - Ignition experiments on thermally-induced ignition - Feasibility study on laser ignition of rocket engines The purpose of the two-color laser plasma experiments is to investigate possible constructive interference effects of driving fields that are not monochromatic, but contain (second) harmonic radiation with respect to the goal of lowering the plasma generation threshold. Such effects have been found in a number of related processes, such as laser ablation or high

  4. Software Estimates Costs of Testing Rocket Engines

    Science.gov (United States)

    Smith, C. L.

    2003-01-01

    Simulation-Based Cost Model (SiCM), a discrete event simulation developed in Extend , simulates pertinent aspects of the testing of rocket propulsion test articles for the purpose of estimating the costs of such testing during time intervals specified by its users. A user enters input data for control of simulations; information on the nature of, and activity in, a given testing project; and information on resources. Simulation objects are created on the basis of this input. Costs of the engineering-design, construction, and testing phases of a given project are estimated from numbers and labor rates of engineers and technicians employed in each phase, the duration of each phase; costs of materials used in each phase; and, for the testing phase, the rate of maintenance of the testing facility. The three main outputs of SiCM are (1) a curve, updated at each iteration of the simulation, that shows overall expenditures vs. time during the interval specified by the user; (2) a histogram of the total costs from all iterations of the simulation; and (3) table displaying means and variances of cumulative costs for each phase from all iterations. Other outputs include spending curves for each phase.

  5. Telemetry Boards Interpret Rocket, Airplane Engine Data

    Science.gov (United States)

    2009-01-01

    For all the data gathered by the space shuttle while in orbit, NASA engineers are just as concerned about the information it generates on the ground. From the moment the shuttle s wheels touch the runway to the break of its electrical umbilical cord at 0.4 seconds before its next launch, sensors feed streams of data about the status of the vehicle and its various systems to Kennedy Space Center s shuttle crews. Even while the shuttle orbiter is refitted in Kennedy s orbiter processing facility, engineers constantly monitor everything from power levels to the testing of the mechanical arm in the orbiter s payload bay. On the launch pad and up until liftoff, the Launch Control Center, attached to the large Vehicle Assembly Building, screens all of the shuttle s vital data. (Once the shuttle clears its launch tower, this responsibility shifts to Mission Control at Johnson Space Center, with Kennedy in a backup role.) Ground systems for satellite launches also generate significant amounts of data. At Cape Canaveral Air Force Station, across the Banana River from Kennedy s location on Merritt Island, Florida, NASA rockets carrying precious satellite payloads into space flood the Launch Vehicle Data Center with sensor information on temperature, speed, trajectory, and vibration. The remote measurement and transmission of systems data called telemetry is essential to ensuring the safe and successful launch of the Agency s space missions. When a launch is unsuccessful, as it was for this year s Orbiting Carbon Observatory satellite, telemetry data also provides valuable clues as to what went wrong and how to remedy any problems for future attempts. All of this information is streamed from sensors in the form of binary code: strings of ones and zeros. One small company has partnered with NASA to provide technology that renders raw telemetry data intelligible not only for Agency engineers, but also for those in the private sector.

  6. Analytical study of nozzle performance for nuclear thermal rockets

    International Nuclear Information System (INIS)

    Davidian, K.O.; Kacynski, K.J.

    1991-01-01

    Nuclear propulsion has been identified as one of the key technologies needed for human exploration of the Moon and Mars. The Nuclear Thermal Rocket (NTR) uses a nuclear reactor to heat hydrogen to a high temperature followed by expansion through a conventional convergent-divergent nozzle. A parametric study of NTR nozzles was performed using the Rocket Engine Design Expert System (REDES) at the NASA Lewis Research Center. The REDES used the JANNAF standard rigorous methodology to determine nozzle performance over a range of chamber temperatures, chamber pressures, thrust levels, and different nozzle configurations. A design condition was set by fixing the propulsion system exit radius at five meters and throat radius was varied to achieve a target thrust level. An adiabatic wall was assumed for the nozzle, and its length was assumed to be 80 percent of a 15 degree cone. The results conclude that although the performance of the NTR, based on infinite reaction rates, looks promising at low chamber pressures, finite rate chemical reactions will cause the actual performance to be considerably lower. Parameters which have a major influence on the delivered specific impulse value include the chamber temperature and the chamber pressures in the high thrust domain. Other parameters, such as 2-D and boundary layer effects, kinetic rates, and number of nozzles, affect the deliverable performance of an NTR nozzle to a lesser degree. For a single nozzle, maximum performance of 930 seconds and 1030 seconds occur at chamber temperatures of 2700 and 3100 K, respectively

  7. Distributed Rocket Engine Testing Health Monitoring System, Phase I

    Data.gov (United States)

    National Aeronautics and Space Administration — The on-ground and Distributed Rocket Engine Testing Health Monitoring System (DiRETHMS) provides a system architecture and software tools for performing diagnostics...

  8. Distributed Rocket Engine Testing Health Monitoring System, Phase II

    Data.gov (United States)

    National Aeronautics and Space Administration — Leveraging the Phase I achievements of the Distributed Rocket Engine Testing Health Monitoring System (DiRETHMS) including its software toolsets and system building...

  9. Advanced Vortex Hybrid Rocket Engine (AVHRE), Phase II

    Data.gov (United States)

    National Aeronautics and Space Administration — ORBITEC proposes to develop a unique Advanced Vortex Hybrid Rocket Engine (AVHRE) to achieve a safe, highly-reliable, low-cost and uniquely versatile propulsion...

  10. Distributed Rocket Engine Testing Health Monitoring System Project

    Data.gov (United States)

    National Aeronautics and Space Administration — Leveraging the Phase I achievements of the Distributed Rocket Engine Testing Health Monitoring System (DiRETHMS) including its software toolsets and system building...

  11. Distributed Rocket Engine Testing Health Monitoring System Project

    Data.gov (United States)

    National Aeronautics and Space Administration — The on-ground and Distributed Rocket Engine Testing Health Monitoring System (DiRETHMS) provides a system architecture and software tools for performing diagnostics...

  12. Propellant Flow Actuated Piezoelectric Rocket Engine Igniter, Phase I

    Data.gov (United States)

    National Aeronautics and Space Administration — Spark ignition of a bi-propellant rocket engine is a classic, proven, and generally reliable process. However, timing can be critical, and the control logic,...

  13. Advanced Vortex Hybrid Rocket Engine (AVHRE), Phase I

    Data.gov (United States)

    National Aeronautics and Space Administration — Orbital Technologies Corporation (ORBITEC) proposes to develop a unique Advanced Vortex Hybrid Rocket Engine (AVHRE) to achieve a highly-reliable, low-cost and...

  14. Propellant-Flow-Actuated Rocket Engine Igniter

    Science.gov (United States)

    Wollen, Mark

    2013-01-01

    A rocket engine igniter has been created that uses a pneumatically driven hammer that, by specialized geometry, is induced into an oscillatory state that can be used to either repeatedly impact a piezoelectric crystal with sufficient force to generate a spark capable of initiating combustion, or can be used with any other system capable of generating a spark from direct oscillatory motion. This innovation uses the energy of flowing gaseous propellant, which by means of pressure differentials and kinetic motion, causes a hammer object to oscillate. The concept works by mass flows being induced through orifices on both sides of a cylindrical tube with one or more vent paths. As the mass flow enters the chamber, the pressure differential is caused because the hammer object is supplied with flow on one side and the other side is opened with access to the vent path. The object then crosses the vent opening and begins to slow because the pressure differential across the ball reverses due to the geometry in the tube. Eventually, the object stops because of the increasing pressure differential on the object until all of the kinetic energy has been transferred to the gas via compression. This is the point where the object reverses direction because of the pressure differential. This behavior excites a piezoelectric crystal via direct impact from the hammer object. The hammer strikes a piezoelectric crystal, then reverses direction, and the resultant high voltage created from the crystal is transferred via an electrode to a spark gap in the ignition zone, thereby providing a spark to ignite the engine. Magnets, or other retention methods, might be employed to favorably position the hammer object prior to start, but are not necessary to maintain the oscillatory behavior. Various manifestations of the igniter have been developed and tested to improve device efficiency, and some improved designs are capable of operation at gas flow rates of a fraction of a gram per second (0

  15. Nuclear Thermal Rocket (NTR) Development Risk Communication

    Science.gov (United States)

    Kim, Tony

    2014-01-01

    There are clear advantages of development of a Nuclear Thermal Rocket (NTR) for a crewed mission to Mars. NTR for in-space propulsion enables more ambitious space missions by providing high thrust at high specific impulse (approximately 900 sec) that is 2 times the best theoretical performance possible for chemical rockets. Missions can be optimized for maximum payload capability to take more payload with reduced total mass to orbit; saving cost on reduction of the number of launch vehicles needed. Or missions can be optimized to minimize trip time significantly to reduce the deep space radiation exposure to the crew. NTR propulsion technology is a game changer for space exploration. However, "NUCLEAR" is a word that is feared and vilified by some groups and the hostility towards development of any nuclear systems can meet great opposition by the public as well as from national leaders and people in authority. Communication of nuclear safety will be critical to the success of the development of the NTR. Why is there a fear of nuclear? A bomb that can level a city is a scary weapon. The first and only times the Nuclear Bomb was used in a war was on Hiroshima and Nagasaki during World War 2. The "Little Boy" atomic bomb was dropped on Hiroshima on August 6, 1945 and the "Fat Man" on Nagasaki 3 days later on August 9th. Within the first 4 months of bombings, 90- 166 thousand people died in Hiroshima and 60-80 thousand died in Nagasaki. It is important to note for comparison that over 500 thousand people died and 5 million made homeless due to strategic bombing (approximately 150 thousand tons) of Japanese cities and war assets with conventional non-nuclear weapons between 1942- 1945. A major bombing campaign of "firebombing" of Tokyo called "Operation Meetinghouse" on March 9 and 10 consisting of 334 B-29's dropped approximately1,700 tons of bombs around 16 square mile area and over 100 thousand people have been estimated to have died. The declaration of death is very

  16. Nuclear thermal rocket nozzle testing and evaluation program

    Science.gov (United States)

    Davidian, Kenneth O.; Kacynski, Kenneth J.

    1993-01-01

    Performance characteristics of the Nuclear Thermal Rocket can be enhanced through the use of unconventional nozzles as part of the propulsion system. The Nuclear Thermal Rocket nozzle testing and evaluation program being conducted at the NASA Lewis is outlined and the advantages of a plug nozzle are described. A facility description, experimental designs and schematics are given. Results of pretest performance analyses show that high nozzle performance can be attained despite substantial nozzle length reduction through the use of plug nozzles as compared to a convergent-divergent nozzle. Pretest measurement uncertainty analyses indicate that specific impulse values are expected to be within + or - 1.17 pct.

  17. Initial Operation of the Nuclear Thermal Rocket Element Environmental Simulator

    Science.gov (United States)

    Emrich, William J., Jr.; Pearson, J. Boise; Schoenfeld, Michael P.

    2015-01-01

    The Nuclear Thermal Rocket Element Environmental Simulator (NTREES) facility is designed to perform realistic non-nuclear testing of nuclear thermal rocket (NTR) fuel elements and fuel materials. Although the NTREES facility cannot mimic the neutron and gamma environment of an operating NTR, it can simulate the thermal hydraulic environment within an NTR fuel element to provide critical information on material performance and compatibility. The NTREES facility has recently been upgraded such that the power capabilities of the facility have been increased significantly. At its present 1.2 MW power level, more prototypical fuel element temperatures nay now be reached. The new 1.2 MW induction heater consists of three physical units consisting of a transformer, rectifier, and inverter. This multiunit arrangement facilitated increasing the flexibility of the induction heater by more easily allowing variable frequency operation. Frequency ranges between 20 and 60 kHz can accommodated in the new induction heater allowing more representative power distributions to be generated within the test elements. The water cooling system was also upgraded to so as to be capable of removing 100% of the heat generated during testing In this new higher power configuration, NTREES will be capable of testing fuel elements and fuel materials at near-prototypic power densities. As checkout testing progressed and as higher power levels were achieved, several design deficiencies were discovered and fixed. Most of these design deficiencies were related to stray RF energy causing various components to encounter unexpected heating. Copper shielding around these components largely eliminated these problems. Other problems encountered involved unexpected movement in the coil due to electromagnetic forces and electrical arcing between the coil and a dummy test article. The coil movement and arcing which were encountered during the checkout testing effectively destroyed the induction coil in use at

  18. A Historical Systems Study of Liquid Rocket Engine Throttling Capabilities

    Science.gov (United States)

    Betts, Erin M.; Frederick, Robert A., Jr.

    2010-01-01

    This is a comprehensive systems study to examine and evaluate throttling capabilities of liquid rocket engines. The focus of this study is on engine components, and how the interactions of these components are considered for throttling applications. First, an assessment of space mission requirements is performed to determine what applications require engine throttling. A background on liquid rocket engine throttling is provided, along with the basic equations that are used to predict performance. Three engines are discussed that have successfully demonstrated throttling. Next, the engine system is broken down into components to discuss special considerations that need to be made for engine throttling. This study focuses on liquid rocket engines that have demonstrated operational capability on American space launch vehicles, starting with the Apollo vehicle engines and ending with current technology demonstrations. Both deep throttling and shallow throttling engines are discussed. Boost and sustainer engines have demonstrated throttling from 17% to 100% thrust, while upper stage and lunar lander engines have demonstrated throttling in excess of 10% to 100% thrust. The key difficulty in throttling liquid rocket engines is maintaining an adequate pressure drop across the injector, which is necessary to provide propellant atomization and mixing. For the combustion chamber, cooling can be an issue at low thrust levels. For turbomachinery, the primary considerations are to avoid cavitation, stall, surge, and to consider bearing leakage flows, rotordynamics, and structural dynamics. For valves, it is necessary to design valves and actuators that can achieve accurate flow control at all thrust levels. It is also important to assess the amount of nozzle flow separation that can be tolerated at low thrust levels for ground testing.

  19. Use of Soft Computing Technologies For Rocket Engine Control

    Science.gov (United States)

    Trevino, Luis C.; Olcmen, Semih; Polites, Michael

    2003-01-01

    The problem to be addressed in this paper is to explore how the use of Soft Computing Technologies (SCT) could be employed to further improve overall engine system reliability and performance. Specifically, this will be presented by enhancing rocket engine control and engine health management (EHM) using SCT coupled with conventional control technologies, and sound software engineering practices used in Marshall s Flight Software Group. The principle goals are to improve software management, software development time and maintenance, processor execution, fault tolerance and mitigation, and nonlinear control in power level transitions. The intent is not to discuss any shortcomings of existing engine control and EHM methodologies, but to provide alternative design choices for control, EHM, implementation, performance, and sustaining engineering. The approaches outlined in this paper will require knowledge in the fields of rocket engine propulsion, software engineering for embedded systems, and soft computing technologies (i.e., neural networks, fuzzy logic, and Bayesian belief networks), much of which is presented in this paper. The first targeted demonstration rocket engine platform is the MC-1 (formerly FASTRAC Engine) which is simulated with hardware and software in the Marshall Avionics & Software Testbed laboratory that

  20. Rocket and Missile Container Engineering Guide

    Science.gov (United States)

    1982-01-01

    complex equations which may, for the sake of expe- diency, be circumvented. To satisfy the intent of this handbook, it will suffice merely to be aware of...tabulates dm given h and Gm. 1-11 RATE OFTRAVEL To attain a required level of protection (G,.- factor), it has been shown that the item to be pro- tected...706-298 This page intentionally left blank. - 2-2 MISSILE OR ROCKET PROFILE SHILLELAGH • REDEYE LAUNCHER (CONTAINS .. MISSILE) M41A2 M41A3

  1. Bimodal Nuclear Thermal Rocket Analysis Developments

    Science.gov (United States)

    Belair, Michael; Lavelle, Thomas; Saimento, Charles; Juhasz, Albert; Stewart, Mark

    2014-01-01

    Nuclear thermal propulsion has long been considered an enabling technology for human missions to Mars and beyond. One concept of operations for these missions utilizes the nuclear reactor to generate electrical power during coast phases, known as bimodal operation. This presentation focuses on the systems modeling and analysis efforts for a NERVA derived concept. The NERVA bimodal operation derives the thermal energy from the core tie tube elements. Recent analysis has shown potential temperature distributions in the tie tube elements that may limit the thermodynamic efficiency of the closed Brayton cycle used to generate electricity with the current design. The results of this analysis are discussed as well as the potential implications to a bimodal NERVA type reactor.

  2. Nuclear thermal rocket clustering: 1, A summary of previous work and relevant issues

    International Nuclear Information System (INIS)

    Buksa, J.J.; Houts, M.G.

    1991-01-01

    A general review of the technical merits of nuclear thermal rocket clustering is presented. A summary of previous analyses performed during the Rover program is presented and used to assess clustering in the context of projected Space Exploration Initiative missions. A number of technical issues are discussed including cluster reliability, engine-out operation, neutronic coupling, shutdown core power generation, shutdown reactivity requirements, reactor kinetics, and radiation shielding. 7 refs., 3 figs., 2 tabs

  3. Additive Manufacturing a Liquid Hydrogen Rocket Engine

    Science.gov (United States)

    Jones, Carl P.; Robertson, Elizabeth H.; Koelbl, Mary Beth; Singer, Chris

    2016-01-01

    Space Propulsion is a 5 day event being held from 2nd May to the 6th May 2016 at the Rome Marriott Park Hotel in Rome, Italy. This event showcases products like Propulsion sub-systems and components, Production and manufacturing issues, Liquid, Solid, Hybrid and Air-breathing Propulsion Systems for Launcher and Upper Stages, Overview of current programmes, AIV issues and tools, Flight testing and experience, Technology building blocks for Future Space Transportation Propulsion Systems : Launchers, Exploration platforms & Space Tourism, Green Propulsion for Space Transportation, New propellants, Rocket propulsion & global environment, Cost related aspects of Space Transportation propulsion, Modelling, Pressure-Thrust oscillations issues, Impact of new requirements and regulations on design etc. in the Automotive, Manufacturing, Fabrication, Repair & Maintenance industries.

  4. Two-step rocket engine bipropellant valve concept

    Science.gov (United States)

    Capps, J. E.; Ferguson, R. E.; Pohl, H. O.

    1969-01-01

    Initiating combustion of altitude control rocket engines in a precombustion chamber of ductile material reduces high pressure surges generated by hypergolic propellants. Two-step bipropellant valve concepts control initial propellant flow into precombustion chamber and subsequent full flow into main chamber.

  5. Liquid rocket engine fluid-cooled combustion chambers

    Science.gov (United States)

    1972-01-01

    A monograph on the design and development of fluid cooled combustion chambers for liquid propellant rocket engines is presented. The subjects discussed are (1) regenerative cooling, (2) transpiration cooling, (3) film cooling, (4) structural analysis, (5) chamber reinforcement, and (6) operational problems.

  6. Analysis of supercritical methane in rocket engine cooling channels

    NARCIS (Netherlands)

    Denies, L.; Zandbergen, B.T.C.; Natale, P.; Ricci, D.; Invigorito, M.

    2016-01-01

    Methane is a promising propellant for liquid rocket engines. As a regenerative coolant, it would be close to its critical point, complicating cooling analysis. This study encompasses the development and validation of a new, open-source computational fluid dynamics (CFD) method for analysis of

  7. Structurally-compliant rocket engine combustion chamber: Experimental/analytical validation

    Science.gov (United States)

    Jankovsky, R. S.; Kazaroff, J. M.; Galford, G. R.; Arya, V. K.

    1993-11-01

    A new, structurally-compliant rocket engine combustion chamber design has been validated through analysis and experiment. Subscale, tubular channel chambers have been cyclically tested, and analytically evaluated. Cyclic lives were determined to have a potential for 1000 percent increase in life over that of rectangular channel designs, the current state-of-the-art. Greater structural compliance in the circumferential direction gives rise to lower thermal strains during hot firing, resulting in lower thermal strain ratcheting and longer predicted fatigue lives. Thermal/durability analyses of the combustion chamber design, involving cyclic temperatures, strains, and low-cycle fatigue lives have corroborated the experimental observations.

  8. Thermohydraulic modeling of nuclear thermal rockets: The KLAXON code

    International Nuclear Information System (INIS)

    Hall, M.L.; Rider, W.J.; Cappiello, M.W.

    1992-01-01

    The hydrogen flow from the storage tanks, through the reactor core, and out the nozzle of a Nuclear Thermal Rocket is an integral design consideration. To provide an analysis and design tool for this phenomenon, the KLAXON code is being developed. A shock-capturing numerical methodology is used to model the gas flow (the Harten, Lax, and van Leer method, as implemented by Einfeldt). Preliminary results of modeling the flow through the reactor core and nozzle are given in this paper

  9. Temperature State of Noncooled Nozzle Adjutage of Liquid Rocket Engine

    Directory of Open Access Journals (Sweden)

    V. S. Zarubin

    2015-01-01

    Full Text Available The increasing specific impulse of the liquid rocket engine (LRE, which is designed to operate in space or in rarefied atmosphere, is directly related to the increasing speed of the combustion gases in the outlet section of the nozzle due to increasing nozzle expansion ratio. An intensity of the convective heat transfer of LRE combustion with the supersonic part of a nozzle shell in the first approximation is inversely proportional to the cross sectional area of gas dynamic path and reduces substantially as approaching to the outlet section of the nozzle.Therefore, in case of large nozzle expansion ratio the use of modern heat-resistant materials allows us to implement its outlet section as a thin-walled uncooled adjutage. This design solution results in reducing total weight of nozzle and decreasing overall preheat of LRE propellant used to cool the engine chamber. For a given diameter of the nozzle outlet section and pressure of combustion gases in this section, to make informed choices of permissible length for uncooled adjutage, it is necessary to have a reliable estimate of its thermal state on the steady-state LRE operation. A mathematical model of the nozzle shell heat transfer with the gas stream taking into account the heat energy transfer by convection and radiation, as well as by heat conduction along the generatrix of the shell enables this estimate.Quantitative analysis of given mathematical model showed that, because of the comparatively low pressure and temperature level of combustion gases, it is acceptable to ignore their own radiation and absorption capacity as compared with the convective heat intensity and the surface nozzle radiation. Thus, re-radiation of its internal surface portions is a factor of importance. Its taking into consideration is the main feature of the developed mathematical model.

  10. Developing Avionics Hardware and Software for Rocket Engine Testing

    Science.gov (United States)

    Aberg, Bryce Robert

    2014-01-01

    My summer was spent working as an intern at Kennedy Space Center in the Propulsion Avionics Branch of the NASA Engineering Directorate Avionics Division. The work that I was involved with was part of Rocket University's Project Neo, a small scale liquid rocket engine test bed. I began by learning about the layout of Neo in order to more fully understand what was required of me. I then developed software in LabView to gather and scale data from two flowmeters and integrated that code into the main control software. Next, I developed more LabView code to control an igniter circuit and integrated that into the main software, as well. Throughout the internship, I performed work that mechanics and technicians would do in order to maintain and assemble the engine.

  11. Reliability Improvements in Liquid Rocket Engine Instrumentation

    Science.gov (United States)

    Hill, A.; Acosta, E.

    2005-01-01

    Instrumentation hardware is often the weak link in advanced liquid fueled propulsion systems. The development of the Space Shuttle Main Engine (SSME) was no exception. By sheer necessity, a reusable, high energy, low weight engine system often relegates the instrumentation hardware to the backseat in the critical hardware development process. This produces less than optimum hardware constraints; including size, location, mounting, redundancy, and signal conditioning. This can negatively affect the development effort and ultimately the system reliability. The challenge was clear, however, the outcome was less certain. Unfortunately, the SSME hardware development culminated in series of measurement failures, most significant of which was the premature engine shutdown during the launch of STS-51F on July 29, 1985. The Return to Flight activities following the Challenger disaster redoubled our efforts to eliminate, once and for all, sensor malfunctions as the determining factor in overall engine reliability. This paper describes each phase of this effort in detail and includes discussion of the tasks related to improving measurement reliability.

  12. Development of Thermal Barriers For Solid Rocket Motor Nozzle Joints

    Science.gov (United States)

    Steinetz, Bruce M.; Dunlap, Patrick H., Jr.

    2000-01-01

    Joints in the Space Shuttle solid rocket motors are sealed by O-rings to contain combustion gases inside the rocket that reach pressures of up to 900 psi and temperatures of up to 5500 F. To provide protection for the O-rings, the motors are insulated with either phenolic or rubber insulation. Gaps in the joints leading up to the O-rings are filled with polysulfide joint-fill compounds as an additional level of protection. The current RSRM nozzle-to-case joint design incorporating primary, secondary, and wiper O-rings experiences gas paths through the joint-fill compound to the innermost wiper O-ring in about one out of every seven motors. Although this does not pose a safety hazard to the motor, it is an undesirable condition that NASA and rocket manufacturer Thiokol want to eliminate. Each nozzle-to-case joint gas path results in extensive reviews and evaluation before flights can be resumed. Thiokol and NASA Marshall are currently working to improve the nozzle-to-case joint design by implementing a more reliable J-leg design that has been used successfully in the field and igniter joint. They are also planning to incorporate the NASA Glenn braided carbon fiber thermal barrier into the joint. The thermal barrier would act as an additional level of protection for the O-rings and allow the elimination of the joint-fill compound from the joint.

  13. Expert System Architecture for Rocket Engine Numerical Simulators: A Vision

    Science.gov (United States)

    Mitra, D.; Babu, U.; Earla, A. K.; Hemminger, Joseph A.

    1998-01-01

    Simulation of any complex physical system like rocket engines involves modeling the behavior of their different components using mostly numerical equations. Typically a simulation package would contain a set of subroutines for these modeling purposes and some other ones for supporting jobs. A user would create an input file configuring a system (part or whole of a rocket engine to be simulated) in appropriate format understandable by the package and run it to create an executable module corresponding to the simulated system. This module would then be run on a given set of input parameters in another file. Simulation jobs are mostly done for performance measurements of a designed system, but could be utilized for failure analysis or a design job such as inverse problems. In order to use any such package the user needs to understand and learn a lot about the software architecture of the package, apart from being knowledgeable in the target domain. We are currently involved in a project in designing an intelligent executive module for the rocket engine simulation packages, which would free any user from this burden of acquiring knowledge on a particular software system. The extended abstract presented here will describe the vision, methodology and the problems encountered in the project. We are employing object-oriented technology in designing the executive module. The problem is connected to the areas like the reverse engineering of any simulation software, and the intelligent systems for simulation.

  14. Nonlinear Control of a Reusable Rocket Engine for Life Extension

    Science.gov (United States)

    Lorenzo, Carl F.; Holmes, Michael S.; Ray, Asok

    1998-01-01

    This paper presents the conceptual development of a life-extending control system where the objective is to achieve high performance and structural durability of the plant. A life-extending controller is designed for a reusable rocket engine via damage mitigation in both the fuel (H2) and oxidizer (O2) turbines while achieving high performance for transient responses of the combustion chamber pressure and the O2/H2 mixture ratio. The design procedure makes use of a combination of linear and nonlinear controller synthesis techniques and also allows adaptation of the life-extending controller module to augment a conventional performance controller of the rocket engine. The nonlinear aspect of the design is achieved using non-linear parameter optimization of a prescribed control structure. Fatigue damage in fuel and oxidizer turbine blades is primarily caused by stress cycling during start-up, shutdown, and transient operations of a rocket engine. Fatigue damage in the turbine blades is one of the most serious causes for engine failure.

  15. Schlieren image velocimetry measurements in a rocket engine exhaust plume

    Science.gov (United States)

    Morales, Rudy; Peguero, Julio; Hargather, Michael

    2017-11-01

    Schlieren image velocimetry (SIV) measures velocity fields by tracking the motion of naturally-occurring turbulent flow features in a compressible flow. Here the technique is applied to measuring the exhaust velocity profile of a liquid rocket engine. The SIV measurements presented include discussion of visibility of structures, image pre-processing for structure visibility, and ability to process resulting images using commercial particle image velocimetry (PIV) codes. The small-scale liquid bipropellant rocket engine operates on nitrous oxide and ethanol as propellants. Predictions of the exhaust velocity are obtained through NASA CEA calculations and simple compressible flow relationships, which are compared against the measured SIV profiles. Analysis of shear layer turbulence along the exhaust plume edge is also presented.

  16. Inspection of a prototype of rocket based combined cycle engine

    Science.gov (United States)

    1998-01-01

    The Direct Gain Solar Thermal Engine was designed with no moving parts. The concept of Solar Thermal Propulsion Research uses focused solar energy from an inflatable concentrator (a giant magnifying glass) to heat a propellant (hydrogen) and allows thermal expansion through the nozzle for low thrust without chemical combustion. Energy limitations and propellant weight associated with traditional combustion engines are non-existant with this concept. The Direct Gain Solar Thermal Engine would be used for moving from a lower orbit to an upper synchronous orbit.

  17. Analysis of supercritical methane in rocket engine cooling channels

    OpenAIRE

    Denies, L.; Zandbergen, B.T.C.; Natale, P.; Ricci, D.; Invigorito, M.

    2016-01-01

    Methane is a promising propellant for liquid rocket engines. As a regenerative coolant, it would be close to its critical point, complicating cooling analysis. This study encompasses the development and validation of a new, open-source computational fluid dynamics (CFD) method for analysis of methane cooling channels. Validation with experimental data has been carried out, showing an accuracy within 20 K for wall temperature and 10% for pressure drop. It is shown that the turbulence model has...

  18. History of the Development of NERVA Nuclear Rocket Engine Technology

    International Nuclear Information System (INIS)

    David L., Black

    2000-01-01

    During the 17 yr between 1955 and 1972, the Atomic Energy Commission (AEC), the U.S. Air Force (USAF), and the National Aeronautics and Space Administration (NASA) collaborated on an effort to develop a nuclear rocket engine. Based on studies conducted in 1946, the concept selected was a fully enriched uranium-filled, graphite-moderated, beryllium-reflected reactor, cooled by a monopropellant, hydrogen. The program, known as Rover, was centered at Los Alamos Scientific Laboratory (LASL), funded jointly by the AEC and the USAF, with the intent of designing a rocket engine for long-range ballistic missiles. Other nuclear rocket concepts were studied during these years, such as cermet and gas cores, but are not reviewed herein. Even thought the program went through the termination phase in a very short time, the technology may still be fully recoverable/retrievable to the state of its prior technological readiness in a reasonably short time. Documents; drawings; and technical, purchasing, manufacturing, and materials specifications were all stored for ease of retrieval. If the U.S. space program were to discover a need/mission for this engine, its 1972 'pencils down' status could be updated for the technology developments of the past 28 yr for flight demonstration in 8 or fewer years. Depending on today's performance requirements, temperatures and pressures could be increased and weight decreased considerably

  19. Contact diagnostics of combustion products of rocket engines, their units, and systems

    Science.gov (United States)

    Ivanov, N. N.; Ivanov, A. N.

    2013-12-01

    This article is devoted to a new block-module device used in the diagnostics of condensed combustion products of rocket engines during research and development with liquid-propellant rocket engines (Glushko NPO Energomash; engines RD-171, RD-180, and RD-191) and solid-propellant rocket motors. Soot samplings from the supersonic high-temperature jet of a high-power liquid-propellant rocket engine were taken by the given device for the first time in practice for closed-exhaust lines. A large quantity of significant results was also obtained during a combustion investigation of solid propellants within solid-propellant rocket motors.

  20. Rover nuclear rocket engine program: Overview of rover engine tests

    Science.gov (United States)

    Finseth, J. L.

    1991-01-01

    The results of nuclear rocket development activities from the inception of the ROVER program in 1955 through the termination of activities on January 5, 1973 are summarized. This report discusses the nuclear reactor test configurations (non cold flow) along with the nuclear furnace demonstrated during this time frame. Included in the report are brief descriptions of the propulsion systems, test objectives, accomplishments, technical issues, and relevant test results for the various reactor tests. Additionally, this document is specifically aimed at reporting performance data and their relationship to fuel element development with little or no emphasis on other (important) items.

  1. Reusable rocket engine preventive maintenance scheduling using genetic algorithm

    International Nuclear Information System (INIS)

    Chen, Tao; Li, Jiawen; Jin, Ping; Cai, Guobiao

    2013-01-01

    This paper deals with the preventive maintenance (PM) scheduling problem of reusable rocket engine (RRE), which is different from the ordinary repairable systems, by genetic algorithm. Three types of PM activities for RRE are considered and modeled by introducing the concept of effective age. The impacts of PM on all subsystems' aging processes are evaluated based on improvement factor model. Then the reliability of engine is formulated by considering the accumulated time effect. After that, optimization model subjected to reliability constraint is developed for RRE PM scheduling at fixed interval. The optimal PM combination is obtained by minimizing the total cost in the whole life cycle for a supposed engine. Numerical investigations indicate that the subsystem's intrinsic reliability characteristic and the improvement factor of maintain operations are the most important parameters in RRE's PM scheduling management

  2. Ablative material testing for low-pressure, low-cost rocket engines

    Science.gov (United States)

    Richter, G. Paul; Smith, Timothy D.

    1995-01-01

    The results of an experimental evaluation of ablative materials suitable for the production of light weight, low cost rocket engine combustion chambers and nozzles are presented. Ten individual specimens of four different compositions of silica cloth-reinforced phenolic resin materials were evaluated for comparative erosion in a subscale rocket engine combustion chamber. Gaseous hydrogen and gaseous oxygen were used as propellants, operating at a nominal chamber pressure of 1138 kPa (165 psi) and a nominal mixture ratio (O/F) of 3.3. These conditions were used to thermally simulate operation with RP-1 and liquid oxygen, and achieved a specimen throat gas temperature of approximately 2456 K (4420 R). Two high-density composition materials exhibited high erosion resistance, while two low-density compositions exhibited approximately 6-75 times lower average erosion resistance. The results compare favorably with previous testing by NASA and provide adequate data for selection of ablatives for low pressure, low cost rocket engines.

  3. Optimal control in thermal engineering

    CERN Document Server

    Badescu, Viorel

    2017-01-01

    This book is the first major work covering applications in thermal engineering and offering a comprehensive introduction to optimal control theory, which has applications in mechanical engineering, particularly aircraft and missile trajectory optimization. The book is organized in three parts: The first part includes a brief presentation of function optimization and variational calculus, while the second part presents a summary of the optimal control theory. Lastly, the third part describes several applications of optimal control theory in solving various thermal engineering problems. These applications are grouped in four sections: heat transfer and thermal energy storage, solar thermal engineering, heat engines and lubrication.Clearly presented and easy-to-use, it is a valuable resource for thermal engineers and thermal-system designers as well as postgraduate students.

  4. Ceramic Matrix Composite Turbine Disk for Rocket Engines

    Science.gov (United States)

    Effinger, Mike; Genge, Gary; Kiser, Doug

    2000-01-01

    NASA has recently completed testing of a ceramic matrix composite (CMC), integrally bladed disk (blisk) for rocket engine turbopumps. The turbopump's main function is to bring propellants from the tank to the combustion chamber at optimal pressures, temperatures, and flow rates. Advantages realized by using CMC blisks are increases in safety by increasing temperature margins and decreasing costs by increasing turbopump performance. A multidisciplinary team, involving materials, design, structural analysis, nondestructive inspection government, academia, and industry experts, was formed to accomplish the 4.5 year effort. This article will review some of the background and accomplishments of the CMC Blisk Program relative to the benefits of this technology.

  5. Effects of rocket engines on laser during lunar landing

    Energy Technology Data Exchange (ETDEWEB)

    Wan, Xiong, E-mail: wanxiong1@126.com [Key Laboratory of Space Active Opto-Electronics Technology, Shanghai Institute of Technical Physics, Chinese Academy of Science, Shanghai 200083 (China); Key Laboratory of Nondestructive Test (Ministry of Education), Nanchang Hangkong University, Nanchang 330063 (China); Shu, Rong; Huang, Genghua [Key Laboratory of Space Active Opto-Electronics Technology, Shanghai Institute of Technical Physics, Chinese Academy of Science, Shanghai 200083 (China)

    2013-11-15

    In the Chinese moon exploration project “ChangE-3”, the laser telemeter and lidar are important equipments on the lunar landing vehicle. A low-thrust vernier rocket engine works during the soft landing, whose plume may influence on the laser equipments. An experiment has first been accomplished to evaluate the influence of the plume on the propagation characteristics of infrared laser under the vacuum condition. Combination with our theoretical analysis has given an appropriate assessment of the plume's effects on the infrared laser hence providing a valuable basis for the design of lunar landing systems.

  6. Low Cost Nuclear Thermal Rocket Cermet Fuel Element Environment Testing

    Science.gov (United States)

    Bradley, David E.; Mireles, Omar R.; Hickman, Robert R.

    2011-01-01

    Deep space missions with large payloads require high specific impulse (Isp) and relatively high thrust in order to achieve mission goals in reasonable time frames. Conventional, storable propellants produce average Isp. Nuclear thermal rockets (NTR) capable of high Isp thrust have been proposed. NTR employs heat produced by fission reaction to heat and therefore accelerate hydrogen which is then forced through a rocket nozzle providing thrust. Fuel element temperatures are very high (up to 3000K) and hydrogen is highly reactive with most materials at high temperatures. Data covering the effects of high temperature hydrogen exposure on fuel elements is limited. The primary concern is the mechanical failure of fuel elements which employ high-melting-point metals, ceramics or a combination (cermet) as a structural matrix into which the nuclear fuel is distributed. It is not necessary to include fissile material in test samples intended to explore high temperature hydrogen exposure of the structural support matrices. A small-scale test bed designed to heat fuel element samples via non-contact RF heating and expose samples to hydrogen is being developed to assist in optimal material and manufacturing process selection without employing fissile material. This paper details the test bed design and results of testing conducted to date.

  7. Nuclear thermal rockets - Key to moon-Mars exploration

    International Nuclear Information System (INIS)

    Borowski, S.K.; Clark, J.S.; Mcilwain, M.C.; Pelaccio, D.G.

    1992-01-01

    The Space Exploration Initiative (SEI) calls for lunar and Martian exploration missions for which solid-core nuclear thermal rockets (NTRs), in virtue of their single-stage, fully-reusable nature, are ideally suited. NTRs promise double the specific impulse of chemical propulsion. A lunar mission employing a reusable NTR is currently being conducted by NASA. The NTR would be assembled in LEO in such a way that it remained 'radioactively cold' during earth-to-orbit deployment by a heavy-lift chemical booster, and therefore presented no radioactive hazard. Also under consideration is a particle-bed reactor in which the hydrogen propulsive fluid directly cools coated-particle fuel spheres

  8. Development and analysis of startup strategies for particle bed nuclear rocket engine

    Science.gov (United States)

    Suzuki, David E.

    1993-06-01

    The particle bed reactor (PBR) nuclear thermal propulsion rocket engine concept is the focus of the Air Force's Space Nuclear Thermal Propulsion program. While much progress has been made in developing the concept, several technical issues remain. Perhaps foremost among these concerns is the issue of flow stability through the porous, heated bed of fuel particles. There are two complementary technical issues associated with this concern: the identification of the flow stability boundary and the design of the engine controller to maintain stable operation. This thesis examines a portion of the latter issue which has yet to be addressed in detail. Specifically, it develops and analyzes general engine system startup strategies which maintain stable flow through the PBR fuel elements while reaching the design conditions as quickly as possible. The PBR engine studies are conducted using a computer model of a representative particle bed reactor and engine system. The computer program utilized is an augmented version of SAFSIM, an existing nuclear thermal propulsion modeling code; the augmentation, dubbed SAFSIM+, was developed by the author and provides a more complete engine system modeling tool.

  9. Distributed Health Monitoring System for Reusable Liquid Rocket Engines

    Science.gov (United States)

    Lin, C. F.; Figueroa, F.; Politopoulos, T.; Oonk, S.

    2009-01-01

    The ability to correctly detect and identify any possible failure in the systems, subsystems, or sensors within a reusable liquid rocket engine is a major goal at NASA John C. Stennis Space Center (SSC). A health management (HM) system is required to provide an on-ground operation crew with an integrated awareness of the condition of every element of interest by determining anomalies, examining their causes, and making predictive statements. However, the complexity associated with relevant systems, and the large amount of data typically necessary for proper interpretation and analysis, presents difficulties in implementing complete failure detection, identification, and prognostics (FDI&P). As such, this paper presents a Distributed Health Monitoring System for Reusable Liquid Rocket Engines as a solution to these problems through the use of highly intelligent algorithms for real-time FDI&P, and efficient and embedded processing at multiple levels. The end result is the ability to successfully incorporate a comprehensive HM platform despite the complexity of the systems under consideration.

  10. Software for Estimating Costs of Testing Rocket Engines

    Science.gov (United States)

    Hines, Merlon M.

    2004-01-01

    A high-level parametric mathematical model for estimating the costs of testing rocket engines and components at Stennis Space Center has been implemented as a Microsoft Excel program that generates multiple spreadsheets. The model and the program are both denoted, simply, the Cost Estimating Model (CEM). The inputs to the CEM are the parameters that describe particular tests, including test types (component or engine test), numbers and duration of tests, thrust levels, and other parameters. The CEM estimates anticipated total project costs for a specific test. Estimates are broken down into testing categories based on a work-breakdown structure and a cost-element structure. A notable historical assumption incorporated into the CEM is that total labor times depend mainly on thrust levels. As a result of a recent modification of the CEM to increase the accuracy of predicted labor times, the dependence of labor time on thrust level is now embodied in third- and fourth-order polynomials.

  11. Rocket-Plume Spectroscopy Simulation for Hydrocarbon-Fueled Rocket Engines

    Science.gov (United States)

    Tejwani, Gopal D.

    2010-01-01

    The UV-Vis spectroscopic system for plume diagnostics monitors rocket engine health by using several analytical tools developed at Stennis Space Center (SSC), including the rocket plume spectroscopy simulation code (RPSSC), to identify and quantify the alloys from the metallic elements observed in engine plumes. Because the hydrocarbon-fueled rocket engine is likely to contain C2, CO, CH, CN, and NO in addition to OH and H2O, the relevant electronic bands of these molecules in the spectral range of 300 to 850 nm in the RPSSC have been included. SSC incorporated several enhancements and modifications to the original line-by-line spectral simulation computer program implemented for plume spectral data analysis and quantification in 1994. These changes made the program applicable to the Space Shuttle Main Engine (SSME) and the Diagnostic Testbed Facility Thruster (DTFT) exhaust plume spectral data. Modifications included updating the molecular and spectral parameters for OH, adding spectral parameter input files optimized for the 10 elements of interest in the spectral range from 320 to 430 nm and linking the output to graphing and analysis packages. Additionally, the ability to handle the non-uniform wavelength interval at which the spectral computations are made was added. This allowed a precise superposition of wavelengths at which the spectral measurements have been made with the wavelengths at which the spectral computations are done by using the line-by-line (LBL) code. To account for hydrocarbon combustion products in the plume, which might interfere with detection and quantification of metallic elements in the spectral region of 300 to 850 nm, the spectroscopic code has been enhanced to include the carbon-based combustion species of C2, CO, and CH. In addition, CN and NO have spectral bands in 300 to 850 nm and, while these molecules are not direct products of hydrocarbon-oxygen combustion systems, they can show up if nitrogen or a nitrogen compound is present

  12. Development Testing of 1-Newton ADN-Based Rocket Engines

    Science.gov (United States)

    Anflo, K.; Gronland, T.-A.; Bergman, G.; Nedar, R.; Thormählen, P.

    2004-10-01

    With the objective to reduce operational hazards and improve specific and density impulse as compared with hydrazine, the Research and Development (R&D) of a new monopropellant for space applications based on AmmoniumDiNitramide (ADN), was first proposed in 1997. This pioneering work has been described in previous papers1,2,3,4 . From the discussion above, it is clear that cost savings as well as risk reduction are the main drivers to develop a new generation of reduced hazard propellants. However, this alone is not enough to convince a spacecraft builder to choose a new technology. Cost, risk and schedule reduction are good incentives, but a spacecraft supplier will ask for evidence that this new propulsion system meets a number of requirements within the following areas: This paper describes the ongoing effort to develop a storable liquid monopropellant blend, based on AND, and its specific rocket engines. After building and testing more than 20 experimental rocket engines, the first Engineering Model (EM-1) has now accumulated more than 1 hour of firing-time. The results from test firings have validated the design. Specific impulse, combustion stability, blow-down capability and short pulse capability are amongst the requirements that have been demonstrated. The LMP-103x propellant candidate has been stored for more than 1 year and initial material compatibility screening and testing has started. 1. Performance &life 2. Impact on spacecraft design &operation 3. Flight heritage Hereafter, the essential requirements for some of these areas are outlined. These issues are discussed in detail in a previous paper1 . The use of "Commercial Of The Shelf" (COTS) propulsion system components as much as possible is essential to minimize the overall cost, risk and schedule. This leads to the conclusion that the Technology Readiness Level (TRL) 5 has been reached for the thruster and propellant. Furthermore, that the concept of ADN-based propulsion is feasible.

  13. The rationale/benefits of nuclear thermal rocket propulsion for NASA's lunar space transportation system

    Science.gov (United States)

    Borowski, Stanley K.

    1994-09-01

    The solid core nuclear thermal rocket (NTR) represents the next major evolutionary step in propulsion technology. With its attractive operating characteristics, which include high specific impulse (approximately 850-1000 s) and engine thrust-to-weight (approximately 4-20), the NTR can form the basis for an efficient lunar space transportation system (LTS) capable of supporting both piloted and cargo missions. Studies conducted at the NASA Lewis Research Center indicate that an NTR-based LTS could transport a fully-fueled, cargo-laden, lunar excursion vehicle to the Moon, and return it to low Earth orbit (LEO) after mission completion, for less initial mass in LEO than an aerobraked chemical system of the type studied by NASA during its '90-Day Study.' The all-propulsive NTR-powered LTS would also be 'fully reusable' and would have a 'return payload' mass fraction of approximately 23 percent--twice that of the 'partially reusable' aerobraked chemical system. Two NTR technology options are examined--one derived from the graphite-moderated reactor concept developed by NASA and the AEC under the Rover/NERVA (Nuclear Engine for Rocket Vehicle Application) programs, and a second concept, the Particle Bed Reactor (PBR). The paper also summarizes NASA's lunar outpost scenario, compares relative performance provided by different LTS concepts, and discusses important operational issues (e.g., reusability, engine 'end-of life' disposal, etc.) associated with using this important propulsion technology.

  14. Software for Preprocessing Data from Rocket-Engine Tests

    Science.gov (United States)

    Cheng, Chiu-Fu

    2004-01-01

    Three computer programs have been written to preprocess digitized outputs of sensors during rocket-engine tests at Stennis Space Center (SSC). The programs apply exclusively to the SSC E test-stand complex and utilize the SSC file format. The programs are the following: Engineering Units Generator (EUGEN) converts sensor-output-measurement data to engineering units. The inputs to EUGEN are raw binary test-data files, which include the voltage data, a list identifying the data channels, and time codes. EUGEN effects conversion by use of a file that contains calibration coefficients for each channel. QUICKLOOK enables immediate viewing of a few selected channels of data, in contradistinction to viewing only after post-test processing (which can take 30 minutes to several hours depending on the number of channels and other test parameters) of data from all channels. QUICKLOOK converts the selected data into a form in which they can be plotted in engineering units by use of Winplot (a free graphing program written by Rick Paris). EUPLOT provides a quick means for looking at data files generated by EUGEN without the necessity of relying on the PV-WAVE based plotting software.

  15. Combustion and Magnetohydrodynamic Processes in Advanced Pulse Detonation Rocket Engines

    Science.gov (United States)

    Cole, Lord Kahil

    A number of promising alternative rocket propulsion concepts have been developed over the past two decades that take advantage of unsteady combustion waves in order to produce thrust. These concepts include the Pulse Detonation Rocket Engine (PDRE), in which repetitive ignition, propagation, and reflection of detonations and shocks can create a high pressure chamber from which gases may be exhausted in a controlled manner. The Pulse Detonation Rocket Induced Magnetohydrodynamic Ejector (PDRIME) is a modification of the basic PDRE concept, developed by Cambier (1998), which has the potential for performance improvements based on magnetohydrodynamic (MHD) thrust augmentation. The PDRIME has the advantage of both low combustion chamber seeding pressure, per the PDRE concept, and efficient energy distribution in the system, per the rocket-induced MHD ejector (RIME) concept of Cole, et al. (1995). In the initial part of this thesis, we explore flow and performance characteristics of different configurations of the PDRIME, assuming quasi-one-dimensional transient flow and global representations of the effects of MHD phenomena on the gas dynamics. By utilizing high-order accurate solvers, we thus are able to investigate the fundamental physical processes associated with the PDRIME and PDRE concepts and identify potentially promising operating regimes. In the second part of this investigation, the detailed coupling of detonations and electric and magnetic fields are explored. First, a one-dimensional spark-ignited detonation with complex reaction kinetics is fully evaluated and the mechanisms for the different instabilities are analyzed. It is found that complex kinetics in addition to sufficient spatial resolution are required to be able to quantify high frequency as well as low frequency detonation instability modes. Armed with this quantitative understanding, we then examine the interaction of a propagating detonation and the applied MHD, both in one-dimensional and two

  16. Analysis of Acoustic Cavitation Surge in a Rocket Engine Turbopump

    Directory of Open Access Journals (Sweden)

    Hideaki Nanri

    2010-01-01

    Full Text Available In a liquid rocket engine, cavitation in an inducer of a turbopump sometimes causes instability phenomena when the inducer is operated at low inlet pressure. Cavitation surge (auto-oscillation, one such instability phenomenon, has been discussed mainly based on an inertia model assuming incompressible flow. When this model is used, the frequency of the cavitation surge decreases continuously as the inlet pressure of the turbopump decreases. However, we obtained an interesting experimental result in which the frequency of cavitation surge varied discontinuously. Therefore, we employed one-dimensional analysis based on an acoustic model in which the fluid is assumed to be compressible. The analytical result qualitatively corresponded with the experimental result.

  17. Critical Performance of Turbopump Mechanical Elements for Rocket Engine

    Science.gov (United States)

    Takada, Satoshi; Kikuchi, Masataka; Sudou, Takayuki; Iwasaki, Fumiya; Watanabe, Yoshiaki; Yoshida, Makoto

    It is generally acknowledged that bearings and axial seals have a tendency to go wrong compared with other rocket engine elements. And when those components have malfunction, missions scarcely succeed. However, fundamental performance (maximum rotational speed, minimum flow rate, power loss, durability, etc.) of those components has not been grasped yet. Purpose of this study is to grasp a critical performance of mechanical seal and hybrid ball bearing of turbopump. In this result, it was found that bearing outer race temperature and bearing coolant outlet temperature changed along saturation line of liquid hydrogen when flow rate was decreased under critical pressure. And normal operation of bearing was possible under conditions of more than 70,000 rpm of rotational speed and more than 0.2 liter/s of coolant flow rate. Though friction coefficient of seal surface increased several times of original value after testing, the seal showed a good performance same as before.

  18. Failure characteristics analysis and fault diagnosis for liquid rocket engines

    CERN Document Server

    Zhang, Wei

    2016-01-01

    This book concentrates on the subject of health monitoring technology of Liquid Rocket Engine (LRE), including its failure analysis, fault diagnosis and fault prediction. Since no similar issue has been published, the failure pattern and mechanism analysis of the LRE from the system stage are of particular interest to the readers. Furthermore, application cases used to validate the efficacy of the fault diagnosis and prediction methods of the LRE are different from the others. The readers can learn the system stage modeling, analyzing and testing methods of the LRE system as well as corresponding fault diagnosis and prediction methods. This book will benefit researchers and students who are pursuing aerospace technology, fault detection, diagnostics and corresponding applications.

  19. Program ELM: A tool for rapid thermal-hydraulic analysis of solid-core nuclear rocket fuel elements

    Science.gov (United States)

    Walton, James T.

    1992-01-01

    This report reviews the state of the art of thermal-hydraulic analysis codes and presents a new code, Program ELM, for analysis of fuel elements. ELM is a concise computational tool for modeling the steady-state thermal-hydraulics of propellant flow through fuel element coolant channels in a nuclear thermal rocket reactor with axial coolant passages. The program was developed as a tool to swiftly evaluate various heat transfer coefficient and friction factor correlations generated for turbulent pipe flow with heat addition which have been used in previous programs. Thus, a consistent comparison of these correlations was performed, as well as a comparison with data from the NRX reactor experiments from the Nuclear Engine for Rocket Vehicle Applications (NERVA) project. This report describes the ELM Program algorithm, input/output, and validation efforts and provides a listing of the code.

  20. Design of a Resistively Heated Thermal Hydraulic Simulator for Nuclear Rocket Reactor Cores

    Science.gov (United States)

    Litchford, Ron J.; Foote, John P.; Ramachandran, Narayanan; Wang, Ten-See; Anghaie, Samim

    2007-01-01

    A preliminary design study is presented for a non-nuclear test facility which uses ohmic heating to replicate the thermal hydraulic characteristics of solid core nuclear reactor fuel element passages. The basis for this testing capability is a recently commissioned nuclear thermal rocket environments simulator, which uses a high-power, multi-gas, wall-stabilized constricted arc-heater to produce high-temperature pressurized hydrogen flows representative of reactor core environments, excepting radiation effects. Initially, the baseline test fixture for this non-nuclear environments simulator was configured for long duration hot hydrogen exposure of small cylindrical material specimens as a low cost means of evaluating material compatibility. It became evident, however, that additional functionality enhancements were needed to permit a critical examination of thermal hydraulic effects in fuel element passages. Thus, a design configuration was conceived whereby a short tubular material specimen, representing a fuel element passage segment, is surrounded by a backside resistive tungsten heater element and mounted within a self-contained module that inserts directly into the baseline test fixture assembly. With this configuration, it becomes possible to create an inward directed radial thermal gradient within the tubular material specimen such that the wall-to-gas heat flux characteristics of a typical fuel element passage are effectively simulated. The results of a preliminary engineering study for this innovative concept are fully summarized, including high-fidelity multi-physics thermal hydraulic simulations and detailed design features.

  1. Nuclear Thermal Rocket Element Environmental Simulator (NTREES) Upgrade Activities

    Science.gov (United States)

    Emrich, William J., Jr.

    2014-01-01

    Over the past year the Nuclear Thermal Rocket Element Environmental Simulator (NTREES) has been undergoing a significant upgrade beyond its initial configuration. The NTREES facility is designed to perform realistic non-nuclear testing of nuclear thermal rocket (NTR) fuel elements and fuel materials. Although the NTREES facility cannot mimic the neutron and gamma environment of an operating NTR, it can simulate the thermal hydraulic environment within an NTR fuel element to provide critical information on material performance and compatibility. The first phase of the upgrade activities which was completed in 2012 in part consisted of an extensive modification to the hydrogen system to permit computer controlled operations outside the building through the use of pneumatically operated variable position valves. This setup also allows the hydrogen flow rate to be increased to over 200 g/sec and reduced the operation complexity of the system. The second stage of modifications to NTREES which has just been completed expands the capabilities of the facility significantly. In particular, the previous 50 kW induction power supply has been replaced with a 1.2 MW unit which should allow more prototypical fuel element temperatures to be reached. The water cooling system was also upgraded to so as to be capable of removing 100% of the heat generated during. This new setup required that the NTREES vessel be raised onto a platform along with most of its associated gas and vent lines. In this arrangement, the induction heater and water systems are now located underneath the platform. In this new configuration, the 1.2 MW NTREES induction heater will be capable of testing fuel elements and fuel materials in flowing hydrogen at pressures up to 1000 psi at temperatures up to and beyond 3000 K and at near-prototypic reactor channel power densities. NTREES is also capable of testing potential fuel elements with a variety of propellants, including hydrogen with additives to inhibit

  2. Solar engineering of thermal processes

    CERN Document Server

    Duffie, John A

    2013-01-01

    The updated fourth edition of the ""bible"" of solar energy theory and applications Over several editions, Solar Engineering of Thermal Processes has become a classic solar engineering text and reference. This revised Fourth Edition offers current coverage of solar energy theory, systems design, and applications in different market sectors along with an emphasis on solar system design and analysis using simulations to help readers translate theory into practice. An important resource for students of solar engineering, solar energy, and alternative energy as well

  3. Performance of an Axisymmetric Rocket Based Combined Cycle Engine During Rocket Only Operation Using Linear Regression Analysis

    Science.gov (United States)

    Smith, Timothy D.; Steffen, Christopher J., Jr.; Yungster, Shaye; Keller, Dennis J.

    1998-01-01

    The all rocket mode of operation is shown to be a critical factor in the overall performance of a rocket based combined cycle (RBCC) vehicle. An axisymmetric RBCC engine was used to determine specific impulse efficiency values based upon both full flow and gas generator configurations. Design of experiments methodology was used to construct a test matrix and multiple linear regression analysis was used to build parametric models. The main parameters investigated in this study were: rocket chamber pressure, rocket exit area ratio, injected secondary flow, mixer-ejector inlet area, mixer-ejector area ratio, and mixer-ejector length-to-inlet diameter ratio. A perfect gas computational fluid dynamics analysis, using both the Spalart-Allmaras and k-omega turbulence models, was performed with the NPARC code to obtain values of vacuum specific impulse. Results from the multiple linear regression analysis showed that for both the full flow and gas generator configurations increasing mixer-ejector area ratio and rocket area ratio increase performance, while increasing mixer-ejector inlet area ratio and mixer-ejector length-to-diameter ratio decrease performance. Increasing injected secondary flow increased performance for the gas generator analysis, but was not statistically significant for the full flow analysis. Chamber pressure was found to be not statistically significant.

  4. Orbital Transfer Rocket Engine Technology. Advanced Engine Study, Task D.6 Final Report

    Science.gov (United States)

    1992-06-01

    development using expert systems and artificial * intelligence techniques. 3.1.2 Power Balance I A rocket engine power balance is an energy and mass balnce com...the whitish, new-copper penny appearance. Blanched surfaces are commonly interconnected by subsurface " wormholing ," severe interconnected porosity, and...In particular, some techniques of artificial intelligence decision making should be adapted to handle this propulsion system. RPT/OOfl?. a/,.o..o 194

  5. Thermal noise engines

    OpenAIRE

    Kish, Laszlo B.

    2010-01-01

    Electrical heat engines driven by the Johnson-Nyquist noise of resistors are introduced. They utilize Coulomb's law and the fluctuation-dissipation theorem of statistical physics that is the reverse phenomenon of heat dissipation in a resistor. No steams, gases, liquids, photons, combustion, phase transition, or exhaust/pollution are present here. In these engines, instead of heat reservoirs, cylinders, pistons and valves, resistors, capacitors and switches are the building elements. For the ...

  6. Innovative concept for an ultra-small nuclear thermal rocket utilizing a new moderated reactor

    Energy Technology Data Exchange (ETDEWEB)

    Nam, Seung Hyun; Venneri, Paolo; Kim, Yong Hee; Lee, Jeong Ik; Chang, Soon Heung; Jeong, Yong Hoon [Dept. of Nuclear and Quantum Engineering, Korea Advanced Institute of Science and Technology, Daejeon (Korea, Republic of)

    2015-10-15

    Although the harsh space environment imposes many severe challenges to space pioneers, space exploration is a realistic and profitable goal for long-term humanity survival. One of the viable and promising options to overcome the harsh environment of space is nuclear propulsion. Particularly, the Nuclear Thermal Rocket (NTR) is a leading candidate for near-term human missions to Mars and beyond due to its relatively high thrust and efficiency. Traditional NTR designs use typically high power reactors with fast or epithermal neutron spectrums to simplify core design and to maximize thrust. In parallel there are a series of new NTR designs with lower thrust and higher efficiency, designed to enhance mission versatility and safety through the use of redundant engines (when used in a clustered engine arrangement) for future commercialization. This paper proposes a new NTR design of the second design philosophy, Korea Advanced NUclear Thermal Engine Rocket (KANUTER), for future space applications. The KANUTER consists of an Extremely High Temperature Gas cooled Reactor (EHTGR) utilizing hydrogen propellant, a propulsion system, and an optional electricity generation system to provide propulsion as well as electricity generation. The innovatively small engine has the characteristics of high efficiency, being compact and lightweight, and bimodal capability. The notable characteristics result from the moderated EHTGR design, uniquely utilizing the integrated fuel element with an ultra heat-resistant carbide fuel, an efficient metal hydride moderator, protectively cooling channels and an individual pressure tube in an all-in-one package. The EHTGR can be bimodally operated in a propulsion mode of 100 MW{sub th} and an electricity generation mode of 100 kW{sub th}, equipped with a dynamic energy conversion system. To investigate the design features of the new reactor and to estimate referential engine performance, a preliminary design study in terms of neutronics and

  7. Innovative concept for an ultra-small nuclear thermal rocket utilizing a new moderated reactor

    Directory of Open Access Journals (Sweden)

    Seung Hyun Nam

    2015-10-01

    Full Text Available Although the harsh space environment imposes many severe challenges to space pioneers, space exploration is a realistic and profitable goal for long-term humanity survival. One of the viable and promising options to overcome the harsh environment of space is nuclear propulsion. Particularly, the Nuclear Thermal Rocket (NTR is a leading candidate for near-term human missions to Mars and beyond due to its relatively high thrust and efficiency. Traditional NTR designs use typically high power reactors with fast or epithermal neutron spectrums to simplify core design and to maximize thrust. In parallel there are a series of new NTR designs with lower thrust and higher efficiency, designed to enhance mission versatility and safety through the use of redundant engines (when used in a clustered engine arrangement for future commercialization. This paper proposes a new NTR design of the second design philosophy, Korea Advanced NUclear Thermal Engine Rocket (KANUTER, for future space applications. The KANUTER consists of an Extremely High Temperature Gas cooled Reactor (EHTGR utilizing hydrogen propellant, a propulsion system, and an optional electricity generation system to provide propulsion as well as electricity generation. The innovatively small engine has the characteristics of high efficiency, being compact and lightweight, and bimodal capability. The notable characteristics result from the moderated EHTGR design, uniquely utilizing the integrated fuel element with an ultra heat-resistant carbide fuel, an efficient metal hydride moderator, protectively cooling channels and an individual pressure tube in an all-in-one package. The EHTGR can be bimodally operated in a propulsion mode of 100 MWth and an electricity generation mode of 100 kWth, equipped with a dynamic energy conversion system. To investigate the design features of the new reactor and to estimate referential engine performance, a preliminary design study in terms of neutronics and

  8. Innovative concept for an ultra-small nuclear thermal rocket utilizing a new moderated reactor

    International Nuclear Information System (INIS)

    Nam, Seung Hyun; Venneri, Paolo; Kim, Yong Hee; Lee, Jeong Ik; Chang, Soon Heung; Jeong, Yong Hoon

    2015-01-01

    Although the harsh space environment imposes many severe challenges to space pioneers, space exploration is a realistic and profitable goal for long-term humanity survival. One of the viable and promising options to overcome the harsh environment of space is nuclear propulsion. Particularly, the Nuclear Thermal Rocket (NTR) is a leading candidate for near-term human missions to Mars and beyond due to its relatively high thrust and efficiency. Traditional NTR designs use typically high power reactors with fast or epithermal neutron spectrums to simplify core design and to maximize thrust. In parallel there are a series of new NTR designs with lower thrust and higher efficiency, designed to enhance mission versatility and safety through the use of redundant engines (when used in a clustered engine arrangement) for future commercialization. This paper proposes a new NTR design of the second design philosophy, Korea Advanced NUclear Thermal Engine Rocket (KANUTER), for future space applications. The KANUTER consists of an Extremely High Temperature Gas cooled Reactor (EHTGR) utilizing hydrogen propellant, a propulsion system, and an optional electricity generation system to provide propulsion as well as electricity generation. The innovatively small engine has the characteristics of high efficiency, being compact and lightweight, and bimodal capability. The notable characteristics result from the moderated EHTGR design, uniquely utilizing the integrated fuel element with an ultra heat-resistant carbide fuel, an efficient metal hydride moderator, protectively cooling channels and an individual pressure tube in an all-in-one package. The EHTGR can be bimodally operated in a propulsion mode of 100 MW th and an electricity generation mode of 100 kW th , equipped with a dynamic energy conversion system. To investigate the design features of the new reactor and to estimate referential engine performance, a preliminary design study in terms of neutronics and thermohydraulics

  9. Thrust stand for low-thrust liquid pulsed rocket engines

    Science.gov (United States)

    Xing, Qin; Zhang, Jun; Qian, Min; Jia, Zhen-yuan; Sun, Bao-yuan

    2010-09-01

    A thrust stand is developed for measuring the pulsed thrust generated by low-thrust liquid pulsed rocket engines. It mainly consists of a thrust dynamometer, a base frame, a connecting frame, and a data acquisition and processing system. The thrust dynamometer assembled with shear mode piezoelectric quartz sensors is developed as the core component of the thrust stand. It adopts integral shell structure. The sensors are inserted into unique double-elastic-half-ring grooves with an interference fit. The thrust is transferred to the sensors by means of static friction forces of fitting surfaces. The sensors could produce an amount of charges which are proportional to the thrust to be measured. The thrust stand is calibrated both statically and dynamically. The in situ static calibration is performed using a standard force sensor. The dynamic calibration is carried out using pendulum-typed steel ball impact technique. Typical thrust pulse is simulated by a trapezoidal impulse force. The results show that the thrust stand has a sensitivity of 25.832 mV/N, a linearity error of 0.24% FSO, and a repeatability error of 0.23% FSO. The first natural frequency of the thrust stand is 1245 Hz. The thrust stand can accurately measure thrust waveform of each firing, which is used for fine control of on-orbit vehicles in the thrust range of 5-20 N with pulse frequency of 50 Hz.

  10. Rocket engine coaxial injector liquid/gas interface flow phenomena

    Science.gov (United States)

    Mayer, Wolfgang; Kruelle, Gerd

    1995-05-01

    Coaxial injectors are used for the injection and mixing of propellants H2/O2 in cryogenic rocket engines. The aim of the theoretical and experimental investigations presented here is to elucidate some of the physical processes in coaxial injector flow with respect to their significance for atomization and mixing. Experiments with the simulation fluids H2O and air were performed under ambient conditions and at elevated counter pressures up to 20 bar. This article reports on phenomenological studies of spray generation under a broad variation of parameters using nanolight photography and high-speed cinematography (up to 3 x 10(exp 4) frames/s). Detailed theoretical and experimental studies of the surface evolution of turbulent jets were performed. Proof was obtained of the impact of internal fluid jet motions on surface deformation. The m = 1 nonaxisymmetric instability of the liquid jet seems to be superimposed onto the small-scale atomization process. A model is presented that calculates droplet atomization quantities as frequency, droplet diameter, and liquid core shape. The overall procedure for implementing this model as a global spray model is also described and an example calculation is presented.

  11. Mars mission opportunity and transit time sensitivity for a nuclear thermal rocket propulsion application

    International Nuclear Information System (INIS)

    Young, A.C.; Mulqueen, J.A.; Nishimuta, E.L.; Emrich, W.J.

    1993-01-01

    President George Bush's 1989 challenge to America to support the Space Exploration Initiative (SEI) of ''Back to the Moon and Human Mission to Mars'' gives the space industry an opportunity to develop effective and efficient space transportation systems. This paper presents stage performance and requirements for a nuclear thermal rocket (NTR) Mars transportation system to support the human Mars mission of the SEI. Two classes of Mars mission profiles are considered in developing the NTR propulsion vehicle performance and requirements. The two Mars mission classes include the opposition class and conjunction class. The opposition class mission is associated with relatively short Mars stay times ranging from 30 to 90 days and total mission duration of 350 to 600 days. The conjunction class mission is associated with much longer Mars stay times ranging from 500 to 600 days and total mission durations of 875 to 1,000 days. Vehicle mass scaling equations are used to determine the NTR stage mass, size, and performance range required for different Mars mission opportunities and for different Mars mission durations. Mission opportunities considered include launch years 2010 to 2018. The 2010 opportunity is the most demanding launch opportunity and the 2018 opportunity is the least demanding opportunity. NTR vehicle mass and size sensitivity to NTR engine thrust level, engine specific impulse, NTR engine thrust-to-weight ratio, and Mars surface payload are presented. NTR propulsion parameter ranges include those associated with NERVA, particle bed reactor (PBR), low-pressure, and ceramic-metal-type engine design

  12. Mars mission opportunity and transit time sensitivity for a nuclear thermal rocket propulsion application

    Science.gov (United States)

    Young, Archie C.; Mulqueen, John A.; Nishimuta, Ena L.; Emrich, William J.

    1993-01-01

    President George Bush's 1989 challenge to America to support the Space Exploration Initiative (SEI) of ``Back to the Moon and Human Mission to Mars'' gives the space industry an opportunity to develop effective and efficient space transportation systems. This paper presents stage performance and requirements for a nuclear thermal rocket (NTR) Mars transportation system to support the human Mars mission of the SEI. Two classes of Mars mission profiles are considered in developing the NTR propulsion vehicle performance and requirements. The two Mars mission classes include the opposition class and conjunction class. The opposition class mission is associated with relatively short Mars stay times ranging from 30 to 90 days and total mission duration of 350 to 600 days. The conjunction class mission is associated with much longer Mars stay times ranging from 500 to 600 days and total mission durations of 875 to 1,000 days. Vehicle mass scaling equations are used to determine the NTR stage mass, size, and performance range required for different Mars mission opportunities and for different Mars mission durations. Mission opportunities considered include launch years 2010 to 2018. The 2010 opportunity is the most demanding launch opportunity and the 2018 opportunity is the least demanding opportunity. NTR vehicle mass and size sensitivity to NTR engine thrust level, engine specific impulse, NTR engine thrust-to-weight ratio, and Mars surface payload are presented. NTR propulsion parameter ranges include those associated with NERVA, particle bed reactor (PBR), low-pressure, and ceramic-metal-type engine design.

  13. Modeling and Testing of Non-Nuclear, Highpower Simulated Nuclear Thermal Rocket Reactor Elements

    Science.gov (United States)

    Kirk, Daniel R.

    2005-01-01

    When the President offered his new vision for space exploration in January of 2004, he said, "Our third goal is to return to the moon by 2020, as the launching point for missions beyond," and, "With the experience and knowledge gained on the moon, we will then be ready to take the next steps of space exploration: human missions to Mars and to worlds beyond." A human mission to Mars implies the need to move large payloads as rapidly as possible, in an efficient and cost-effective manner. Furthermore, with the scientific advancements possible with Project Prometheus and its Jupiter Icy Moons Orbiter (JIMO), (these use electric propulsion), there is a renewed interest in deep space exploration propulsion systems. According to many mission analyses, nuclear thermal propulsion (NTP), with its relatively high thrust and high specific impulse, is a serious candidate for such missions. Nuclear rockets utilize fission energy to heat a reactor core to very high temperatures. Hydrogen gas flowing through the core then becomes superheated and exits the engine at very high exhaust velocities. The combination of temperature and low molecular weight results in an engine with specific impulses above 900 seconds. This is almost twice the performance of the LOX/LH2 space shuttle engines, and the impact of this performance would be to reduce the trip time of a manned Mars mission from the 2.5 years, possible with chemical engines, to about 12-14 months.

  14. Oxidation Behavior of Copper Alloy Candidates for Rocket Engine Applications (Technical Poster)

    Science.gov (United States)

    Ogbuji, Linus U. J.; Humphrey, Donald H.; Barrett, Charles A.; Greenbauer-Seng, Leslie (Technical Monitor); Gray, Hugh R. (Technical Monitor)

    2002-01-01

    A rocket engine's combustion chamber is lined with material that is highly conductive to heat in order to dissipate the huge thermal load (evident in a white-hot exhaust plume). Because of its thermal conductivity copper is the best choice of liner material. However, the mechanical properties of pure copper are inadequate to withstand the high stresses, hence, copper alloys are needed in this application. But copper and its alloys are prone to oxidation and related damage, especially "blanching" (an oxidation-reduction mode of degradation). The space shuttle main engine combustion chamber is lined with a Cu-Ag-Zr alloy, "NARloy-Z", which exhibits blanching. A superior liner is being sought for the next generation of RLVs (Reusable Launch Vehicles) It should have improved mechanical properties and higher resistance to oxidation and blanching, but without substantial penalty in thermal conductivity. GRCop84, a Cu-8Cr-4Nb alloy (Cr2Nb in Cu matrix), developed by NASA Glenn Research Center (GRC) and Case Western Reserve University, is a prime contender for RLV liner material. In this study, the oxidation resistance of GRCop-84 and other related/candidate copper alloys are investigated and compared

  15. A Numerical Study of Combined Convective and Radiative Heat Transfer in a Rocket Engine Combustion Chamber

    National Research Council Canada - National Science Library

    Savur, Mehmet

    2002-01-01

    A numerical study was conducted to predict the combined convective and radiative heat transfer rates on the walls of a small aspect ratio cylinder representative of the scaled model of a rocket engine combustion chamber...

  16. Improved Rocket Test Engine Video Recording with Computational Photography and Computer Vision Techniques

    Data.gov (United States)

    National Aeronautics and Space Administration — High energy processes such as rocket engine flight certification ground testing require high-speed, high dynamic range video imaging in order to capture and record...

  17. Improved Rocket Test Engine Video Recording with Computational Photography and Computer Vision Techniques

    Data.gov (United States)

    National Aeronautics and Space Administration — Rocket engine flight certification ground testing requires high-speed video recording that can capture essential information for NASA. This need is particularly true...

  18. Quasi-2D Unsteady Flow Solver Module for Rocket Engine and Propulsion System Simulations

    National Research Council Canada - National Science Library

    Campell, Bryan T; Davis, Roger L

    2006-01-01

    .... The solver is targeted to the commercial dynamic simulation software package Simulink(Registered) for integration into a larger suite of modules developed for simulating rocket engines and propulsion systems...

  19. Torch-Augmented Spark Igniter for Nanosat Launch Vehicle LOX/Propylene Rocket Engine, Phase I

    Data.gov (United States)

    National Aeronautics and Space Administration — The technical innovation proposed here is the introduction of torch-augmented spark ignition for high performance liquid oxygen (LOX) / propylene rocket engines now...

  20. Code Validation of CFD Heat Transfer Models for Liquid Rocket Engine Combustion Devices

    National Research Council Canada - National Science Library

    Coy, E. B

    2007-01-01

    .... The design of the rig and its capabilities are described. A second objective of the test rig is to provide CFD validation data under conditions relevant to liquid rocket engine thrust chambers...

  1. Regeneratively-Cooled, Turbopump-Fed, Small-Scale Cryogenic Rocket Engines, Phase II

    Data.gov (United States)

    National Aeronautics and Space Administration — To-date, the realization of small-scale, high-performance liquid bipropellant rocket engines has largely been limited by the inability to operate at high chamber...

  2. Multicamera High Dynamic Range High-Speed Video of Rocket Engine Tests and Launches

    Data.gov (United States)

    National Aeronautics and Space Administration — High-speed video recording of rocket engine tests has several challenges. The scenes that are imaged have both bright and dark regions associated with plume emission...

  3. Proposal for a Concept Assessment of a Fission Fragment Rocket Engine (FFRE) Propelled Spacecraft Project

    Data.gov (United States)

    National Aeronautics and Space Administration — A new technology, the Fission Fragment Rocket Engine (FFRE), requires small amounts of readily available, energy dense, long lasting fuel, significant thrust at...

  4. Contamination Control for Thermal Engineers

    Science.gov (United States)

    Rivera, Rachel B.

    2015-01-01

    The presentation will be given at the 26th Annual Thermal Fluids Analysis Workshop (TFAWS 2015) hosted by the Goddard Spaceflight Center (GSFC) Thermal Engineering Branch (Code 545). This course will cover the basics of Contamination Control, including contamination control related failures, the effects of contamination on Flight Hardware, what contamination requirements translate to, design methodology, and implementing contamination control into Integration, Testing and Launch.

  5. Development of test stand for experimental investigation of chemical and physical phenomena in Liquid Rocket Engine

    OpenAIRE

    Emerson Andrade Santos; Wilton Fernandes Alves; André Neves Almeida Prado; Cristiane Aparecida Martins

    2011-01-01

    Abstract The main objective of this work was to present the specification of an experimental firing test stand for liquid rocket engines (LRE) and develop a program for control and acquisition of data. It provides conditions to test rocket engines with thrust from 50 to 100 kgf. A methodology for laboratory work implementation using information technology, which will allow the automatic and remote functioning of the test stand, permits users to input the necessary data to conduct tests safely...

  6. LOX/Methane Regeneratively-Cooled Rocket Engine Development

    Data.gov (United States)

    National Aeronautics and Space Administration — The purpose of this project is to advance the technologies required to build a subcritical regeneratively cooled liquid oxygen/methane rocket combustion chamber for...

  7. Nuclear Thermal Rocket/Vehicle Design Options for Future NASA Missions to the Moon and Mars

    Science.gov (United States)

    Borowski, Stanley K.; Corban, Robert R.; Mcguire, Melissa L.; Beke, Erik G.

    1995-01-01

    The nuclear thermal rocket (NTR) provides a unique propulsion capability to planners/designers of future human exploration missions to the Moon and Mars. In addition to its high specific impulse (approximately 850-1000 s) and engine thrust-to-weight ratio (approximately 3-10), the NTR can also be configured as a 'dual mode' system capable of generating electrical power for spacecraft environmental systems, communications, and enhanced stage operations (e.g., refrigeration for long-term liquid hydrogen storage). At present the Nuclear Propulsion Office (NPO) is examining a variety of mission applications for the NTR ranging from an expendable, single-burn, trans-lunar injection (TLI) stage for NASA's First Lunar Outpost (FLO) mission to all propulsive, multiburn, NTR-powered spacecraft supporting a 'split cargo-piloted sprint' Mars mission architecture. Each application results in a particular set of requirements in areas such as the number of engines and their respective thrust levels, restart capability, fuel operating temperature and lifetime, cryofluid storage, and stage size. Two solid core NTR concepts are examined -- one based on NERVA (Nuclear Engine for Rocket Vehicle Application) derivative reactor (NDR) technology, and a second concept which utilizes a ternary carbide 'twisted ribbon' fuel form developed by the Commonwealth of Independent States (CIS). The NDR and CIS concepts have an established technology database involving significant nuclear testing at or near representative operating conditions. Integrated systems and mission studies indicate that clusters of two to four 15 to 25 klbf NDR or CIS engines are sufficient for most of the lunar and Mars mission scenarios currently under consideration. This paper provides descriptions and performance characteristics for the NDR and CIS concepts, summarizes NASA's First Lunar Outpost and Mars mission scenarios, and describes characteristics for representative cargo and piloted vehicles compatible with a

  8. Some Calculated Research Results of the Working Process Parameters of the Low Thrust Rocket Engine Operating on Gaseous Oxygen-Hydrogen Fuel

    Science.gov (United States)

    Ryzhkov, V.; Morozov, I.

    2018-01-01

    The paper presents the calculating results of the combustion products parameters in the tract of the low thrust rocket engine with thrust P ∼ 100 N. The article contains the following data: streamlines, distribution of total temperature parameter in the longitudinal section of the engine chamber, static temperature distribution in the cross section of the engine chamber, velocity distribution of the combustion products in the outlet section of the engine nozzle, static temperature near the inner wall of the engine. The presented parameters allow to estimate the efficiency of the mixture formation processes, flow of combustion products in the engine chamber and to estimate the thermal state of the structure.

  9. Copper Multiwall Carbon Nanotubes and Copper-Diamond Composites for Advanced Rocket Engines

    Science.gov (United States)

    Bhat, Biliyar N.; Ellis, Dave L.; Smelyanskiy, Vadim; Foygel, Michael; Singh, Jogender; Rape, Aaron; Vohra, Yogesh; Thomas, Vinoy; Li, Deyu; Otte, Kyle

    2013-01-01

    This paper reports on the research effort to improve the thermal conductivity of the copper-based alloy NARloy-Z (Cu-3 wt.%Ag-0.5 wt.% Zr), the state-of-the-art alloy used to make combustion chamber liners in regeneratively-cooled liquid rocket engines, using nanotechnology. The approach was to embed high thermal conductivity multiwall carbon nanotubes (MWCNTs) and diamond (D) particles in the NARloy-Z matrix using powder metallurgy techniques. The thermal conductivity of MWCNTs and D have been reported to be 5 to 10 times that of NARloy-Z. Hence, 10 to 20 vol. % MWCNT finely dispersed in NARloy-Z matrix could nearly double the thermal conductivity, provided there is a good thermal bond between MWCNTs and copper matrix. Quantum mechanics-based modeling showed that zirconium (Zr) in NARloy-Z should form ZrC at the MWCNT-Cu interface and provide a good thermal bond. In this study, NARloy-Z powder was blended with MWCNTs in a ball mill, and the resulting mixture was consolidated under high pressure and temperature using Field Assisted Sintering Technology (FAST). Microstructural analysis showed that the MWCNTs, which were provided as tangles of MWCNTs by the manufacturer, did not detangle well during blending and formed clumps at the prior particle boundaries. The composites made form these powders showed lower thermal conductivity than the base NARloy-Z. To eliminate the observed physical agglomeration, tangled multiwall MWCNTs were separated by acid treatment and electroless plated with a thin layer of chromium to keep them separated during further processing. Separately, the thermal conductivities of MWCNTs used in this work were measured, and the results showed very low values, a major factor in the low thermal conductivity of the composite. On the other hand, D particles embedded in NARloy-Z matrix showed much improved thermal conductivity. Elemental analysis showed migration of Zr to the NARloy-Z-D interface to form ZrC, which appeared to provide a low contact

  10. Copper-Multiwall Carbon Nanotubes and Copper-Diamond Composites for Advanced Rocket Engines

    Science.gov (United States)

    Bhat, Biliyar N.; Ellis, Dave L.; Smelyanskiy, Vadim; Foygel, Michael; Rape, Aaron; Singh, Jogender; Vohra, Yogesh K.; Thomas, Vinoy; Otte, Kyle G.; Li, Deyu

    2013-01-01

    This paper reports on the research effort to improve the thermal conductivity of the copper-based alloy NARloy-Z (Cu-3 wt.%Ag-0.5 wt.% Zr), the state-of-the-art alloy used to make combustion chamber liners in regeneratively-cooled liquid rocket engines, using nanotechnology. The approach was to embed high thermal conductivity multiwall carbon nanotubes (MWCNTs) and diamond (D) particles in the NARloy-Z matrix using powder metallurgy techniques. The thermal conductivity of MWCNTs and D have been reported to be 5 to 10 times that of NARloy-Z. Hence, 10 to 20 vol. % MWCNT finely dispersed in NARloy-Z matrix could nearly double the thermal conductivity, provided there is a good thermal bond between MWCNTs and copper matrix. Quantum mechanics-based modeling showed that zirconium (Zr) in NARloy-Z should form ZrC at the MWCNT-Cu interface and provide a good thermal bond. In this study, NARloy-Z powder was blended with MWCNTs in a ball mill, and the resulting mixture was consolidated under high pressure and temperature using Field Assisted Sintering Technology (FAST). Microstructural analysis showed that the MWCNTs, which were provided as tangles of MWCNTs by the manufacturer, did not detangle well during blending and formed clumps at the prior particle boundaries. The composites made form these powders showed lower thermal conductivity than the base NARloy-Z. To eliminate the observed physical agglomeration, tangled multiwall MWCNTs were separated by acid treatment and electroless plated with a thin layer of chromium to keep them separated during further processing. Separately, the thermal conductivities of MWCNTs used in this work were measured, and the results showed very low values, a major factor in the low thermal conductivity of the composite. On the other hand, D particles embedded in NARloy-Z matrix showed much improved thermal conductivity. Elemental analysis showed migration of Zr to the NARloy-Z-D interface to form ZrC, which appeared to provide a low contact

  11. An historical perspective of the NERVA nuclear rocket engine technology program. Final Report

    International Nuclear Information System (INIS)

    Robbins, W.H.; Finger, H.B.

    1991-07-01

    Nuclear rocket research and development was initiated in the United States in 1955 and is still being pursued to a limited extent. The major technology emphasis occurred in the decade of the 1960s and was primarily associated with the Rover/NERVA Program where the technology for a nuclear rocket engine system for space application was developed and demonstrated. The NERVA (Nuclear Engine for Rocket Vehicle Application) technology developed twenty years ago provides a comprehensive and viable propulsion technology base that can be applied and will prove to be valuable for application to the NASA Space Exploration Initiative (SEI). This paper, which is historical in scope, provides an overview of the conduct of the NERVA Engine Program, its organization and management, development philosophy, the engine configuration, and significant accomplishments

  12. An historical perspective of the NERVA nuclear rocket engine technology program

    Science.gov (United States)

    Robbins, W. H.; Finger, H. B.

    1991-01-01

    Nuclear rocket research and development was initiated in the United States in 1955 and is still being pursued to a limited extent. The major technology emphasis occurred in the decade of the 1960s and was primarily associated with the Rover/NERVA Program where the technology for a nuclear rocket engine system for space application was developed and demonstrated. The NERVA (Nuclear Engine for Rocket Vehicle Application) technology developed twenty years ago provides a comprehensive and viable propulsion technology base that can be applied and will prove to be valuable for application to the NASA Space Exploration Initiative (SEI). This paper, which is historical in scope, provides an overview of the conduct of the NERVA Engine Program, its organization and management, development philosophy, the engine configuration, and significant accomplishments.

  13. Theoretical Acoustic Absorber Design Approach for LOX/LCH4 Pintle Injector Rocket Engines

    Science.gov (United States)

    Candelaria, Jonathan

    Liquid rocket engines, or LREs, have served a key role in space exploration efforts. One current effort involves the utilization of liquid oxygen (LOX) and liquid methane (LCH4) LREs to explore Mars with in-situ resource utilization for propellant production. This on-site production of propellant will allow for greater payload allocation instead of fuel to travel to the Mars surface, and refueling of propellants to travel back to Earth. More useable mass yields a greater benefit to cost ratio. The University of Texas at El Paso's (UTEP) Center for Space Exploration and Technology Research Center (cSETR) aims to further advance these methane propulsion systems with the development of two liquid methane - liquid oxygen propellant combination rocket engines. The design of rocket engines, specifically liquid rocket engines, is complex in that many variables are present that must be taken into consideration in the design. A problem that occurs in almost every rocket engine development program is combustion instability, or oscillatory combustion. It can result in the destruction of the rocket, subsequent destruction of the vehicle and compromise the mission. These combustion oscillations can vary in frequency from 100 to 20,000 Hz or more, with varying effects, and occur from different coupling phenomena. It is important to understand the effects of combustion instability, its physical manifestations, how to identify the instabilities, and how to mitigate or dampen them. Linear theory methods have been developed to provide a mathematical understanding of the low- to mid-range instabilities. Nonlinear theory is more complex and difficult to analyze mathematically, therefore no general analytical method that yields a solution exists. With limited resources, time, and the advice of our NASA mentors, a data driven experimental approach utilizing quarter wave acoustic dampener cavities was designed. This thesis outlines the methodology behind the design of an acoustic

  14. RECENT ACTIVITIES AT THE CENTER FOR SPACE NUCLEAR RESEARCH FOR DEVELOPING NUCLEAR THERMAL ROCKETS

    Energy Technology Data Exchange (ETDEWEB)

    Robert C. O' Brien

    2001-09-01

    Nuclear power has been considered for space applications since the 1960s. Between 1955 and 1972 the US built and tested over twenty nuclear reactors/ rocket-engines in the Rover/NERVA programs. However, changes in environmental laws may make the redevelopment of the nuclear rocket more difficult. Recent advances in fuel fabrication and testing options indicate that a nuclear rocket with a fuel form significantly different from NERVA may be needed to ensure public support. The Center for Space Nuclear Research (CSNR) is pursuing development of tungsten based fuels for use in a NTR, for a surface power reactor, and to encapsulate radioisotope power sources. The CSNR Summer Fellows program has investigated the feasibility of several missions enabled by the NTR. The potential mission benefits of a nuclear rocket, historical achievements of the previous programs, and recent investigations into alternatives in design and materials for future systems will be discussed.

  15. Dual Regenerative Cooling Circuits for Liquid Rocket Engines (POSTPRINT)

    National Research Council Canada - National Science Library

    Naraghi, N. H; Dunn, S; Coats, D

    2006-01-01

    .... Two engines, the SSME and a RP1-LOX engine, are retrofitted with dual-circuits. It is shown that the maximum wall temperatures for both engines are substantially reduced while also lowering coolant pumping power...

  16. An Approximate Analysis of the Inner Wall Loading of a Bimetallic Camera Shell of Reusable Rocket Engine

    OpenAIRE

    V. S. Zarubin; V. N. Zimin; G. N. Kuvyrkin

    2016-01-01

    Various technical devices quite widely use bimetallic shells as the structural elements. A chamber combustion design of the liquid rocket engine (LRE) is a typical use of the bimetallic shells.In LRE operation a combustion chamber shell is subject to intense thermal and mechanical effects, which necessitates cooling. A cooling shell path is formed by a gap between its inner and outer walls connected to each other by milled or grooved spacer ribs. The outer wall of the shell serves as a load-b...

  17. Investigation of low cost material processes for liquid rocket engines

    Science.gov (United States)

    Nguyentat, Thinh; Kawashige, Chester M.; Scala, James G.; Horn, Ronald M.

    1993-01-01

    The development of low cost material processes is essential to the achievement of economical liquid rocket propulsion systems in the next century. This paper will present the results of the evaluation of some promising material processes including powder metallurgy, vacuum plasma spray, metal spray forming, and bulge forming. The physical and mechanical test results from the samples and subscale hardware fabricated from high strength copper alloys and superalloys will be discussed.

  18. Computer Design Technology of the Small Thrust Rocket Engines Using CAE / CAD Systems

    Science.gov (United States)

    Ryzhkov, V.; Lapshin, E.

    2018-01-01

    The paper presents an algorithm for designing liquid small thrust rocket engine, the process of which consists of five aggregated stages with feedback. Three stages of the algorithm provide engineering support for design, and two stages - the actual engine design. A distinctive feature of the proposed approach is a deep study of the main technical solutions at the stage of engineering analysis and interaction with the created knowledge (data) base, which accelerates the process and provides enhanced design quality. The using multifunctional graphic package Siemens NX allows to obtain the final product -rocket engine and a set of design documentation in a fairly short time; the engine design does not require a long experimental development.

  19. Combustion oscillation study in a kerosene fueled rocket-based combined-cycle engine combustor

    Science.gov (United States)

    Huang, Zhi-Wei; He, Guo-Qiang; Qin, Fei; Xue, Rui; Wei, Xiang-Geng; Shi, Lei

    2016-12-01

    This study reports the combustion oscillation features in a three-dimensional (3D) rocket-based combined-cycle (RBCC) engine combustor under flight Mach number (Mflight) 3.0 conditions both experimentally and numerically. Experiment is performed on a direct-connect ground test facility, which measures the wall pressure along the flow-path. High-speed imaging of the flame luminosity and schlieren is carried out at exit of the primary rocket. Compressible reactive large eddy simulation (LES) with reduced chemical kinetics of a surrogate model for kerosene is performed to further understand the combustion oscillation mechanisms in the combustor. LES results are validated with experimental data by the time-averaged and root mean square (RMS) pressure values, and show acceptable agreement. Effects of the primary rocket jet on pressure oscillation in the combustor are analyzed. Relation of the high speed rocket jet oscillation, which is thought to among the most probable sources of combustion oscillation, with the RBCC combustor is recognized. Results reveal that the unsteady over-expanded rocket jet has significant impacts on the combustion oscillation feature of the RBCC combustor, which is different from a thermo-acoustics type oscillation. The rocket jet/air inflow physical interactions under different rocket jet expansion degrees are experimentally studied.

  20. Rocket Science: The Shuttle's Main Engines, though Old, Are not Forgotten in the New Exploration Initiative

    Science.gov (United States)

    Covault, Craig

    2005-01-01

    The Space Shuttle Main Engine (SSME), developed 30 years ago, remains a strong candidate for use in the new Exploration Initiative as part of a shuttle-derived heavy-lift expendable booster. This is because the Boeing-Rocket- dyne man-rated SSME remains the most highly efficient liquid rocket engine ever developed. There are only enough parts for 12-15 existing SSMEs, however, so one NASA option is to reinitiate SSME production to use it as a throw-away, as opposed to a reusable, powerplant for NASA s new heavy-lift booster.

  1. LOX/Methane Regeneratively-Cooled Rocket Engine Development Project

    Data.gov (United States)

    National Aeronautics and Space Administration — Design, build, and test a 5,000 lbf thrust regeneratively cooled combustion chamber at JSC for a low pressure liquid oxygen/methane engine. The engine demonstrates...

  2. Laser Ignition Technology for Bi-Propellant Rocket Engine Applications

    Science.gov (United States)

    Thomas, Matthew E.; Bossard, John A.; Early, Jim; Trinh, Huu; Dennis, Jay; Turner, James (Technical Monitor)

    2001-01-01

    The fiber optically coupled laser ignition approach summarized is under consideration for use in igniting bi-propellant rocket thrust chambers. This laser ignition approach is based on a novel dual pulse format capable of effectively increasing laser generated plasma life times up to 1000 % over conventional laser ignition methods. In the dual-pulse format tinder consideration here an initial laser pulse is used to generate a small plasma kernel. A second laser pulse that effectively irradiates the plasma kernel follows this pulse. Energy transfer into the kernel is much more efficient because of its absorption characteristics thereby allowing the kernel to develop into a much more effective ignition source for subsequent combustion processes. In this research effort both single and dual-pulse formats were evaluated in a small testbed rocket thrust chamber. The rocket chamber was designed to evaluate several bipropellant combinations. Optical access to the chamber was provided through small sapphire windows. Test results from gaseous oxygen (GOx) and RP-1 propellants are presented here. Several variables were evaluated during the test program, including spark location, pulse timing, and relative pulse energy. These variables were evaluated in an effort to identify the conditions in which laser ignition of bi-propellants is feasible. Preliminary results and analysis indicate that this laser ignition approach may provide superior ignition performance relative to squib and torch igniters, while simultaneously eliminating some of the logistical issues associated with these systems. Further research focused on enhancing the system robustness, multiplexing, and window durability/cleaning and fiber optic enhancements is in progress.

  3. Fiber-reinforced ceramic composites for Earth-to-orbit rocket engine turbines

    Science.gov (United States)

    Brockmeyer, Jerry W.; Schnittgrund, Gary D.

    1990-01-01

    Fiber reinforced ceramic matrix composites (FRCMC) are emerging materials systems that offer potential for use in liquid rocket engines. Advantages of these materials in rocket engine turbomachinery include performance gain due to higher turbine inlet temperature, reduced launch costs, reduced maintenance with associated cost benefits, and reduced weight. This program was initiated to assess the state of FRCMC development and to propose a plan for their implementation into liquid rocket engine turbomachinery. A complete range of FRCMC materials was investigated relative to their development status and feasibility for use in the hot gas path of earth-to-orbit rocket engine turbomachinery. Of the candidate systems, carbon fiber-reinforced silicon carbide (C/SiC) offers the greatest near-term potential. Critical hot gas path components were identified, and the first stage inlet nozzle and turbine rotor of the fuel turbopump for the liquid oxygen/hydrogen Space Transportation Main Engine (STME) were selected for conceptual design and analysis. The critical issues associated with the use of FRCMC were identified. Turbine blades were designed, analyzed and fabricated. The Technology Development Plan, completed as Task 5 of this program, provides a course of action for resolution of these issues.

  4. Development and Analysis of Startup Strategies for Particle Bed Nuclear Rocket Engine

    Science.gov (United States)

    1993-06-01

    M. D. Hoover, eds. American Institute of Physics, New York, 1993. [L-3] Ludewig , Hans. "Particle Bed Reactor Nuclear Rocket Concept." Nuclear Thermal...July 1988. [S-I] Stafford, Thomas (Chairman). America’s Space Exploration Initiative: America at the Threshold. Report of the Stafford Committee to

  5. Verification of Model of Calculation of Intra-Chamber Parameters In Hybrid Solid-Propellant Rocket Engines

    Directory of Open Access Journals (Sweden)

    Zhukov Ilya S.

    2016-01-01

    Full Text Available On the basis of obtained analytical estimate of characteristics of hybrid solid-propellant rocket engine verification of earlier developed physical and mathematical model of processes in a hybrid solid-propellant rocket engine for quasi-steady-state flow regime was performed. Comparative analysis of calculated and analytical data indicated satisfactory comparability of simulation results.

  6. JANNAF "Test and Evaluation Guidelines for Liquid Rocket Engines": Status and Application

    Science.gov (United States)

    Parkinson, Douglas; VanLerberghe, Wayne M.; Rahman, Shamim A.

    2017-01-01

    For many decades, the U.S. rocket propulsion industrial base has performed remarkably in developing complex liquid rocket engines that can propel critical payloads into service for the nation, as well as transport people and hardware for missions that open the frontiers of space exploration for humanity. This has been possible only at considerable expense given the lack of detailed guidance that captures the essence of successful practices and knowledge accumulated over five decades of liquid rocket engine development. In an effort to provide benchmarks and guidance for the next generation of rocket engineers, the Joint Army Navy NASA Air Force (JANNAF) Interagency Propulsion Committee published a liquid rocket engine (LRE) test and evaluation (T&E) guideline document in 2012 focusing on the development challenges and test verification considerations for liquid rocket engine systems. This document has been well received and applied by many current LRE developers as a benchmark and guidance tool, both for government-driven applications as well as for fully commercial ventures. The USAF Space and Missile Systems Center (SMC) has taken an additional near-term step and is directing activity to adapt and augment the content from the JANNAF LRE T&E guideline into a standard for potential application to future USAF requests for proposals for LRE development initiatives and launch vehicles for national security missions. A draft of this standard was already sent out for review and comment, and is intended to be formally approved and released towards the end of 2017. The acceptance and use of the LRE T&E guideline is possible through broad government and industry participation in the JANNAF liquid propulsion committee and associated panels. The sponsoring JANNAF community is expanding upon this initial baseline version and delving into further critical development aspects of liquid rocket propulsion testing at the integrated stage level as well as engine component level, in

  7. Technology developments for thrust chambers of future launch vehicle liquid rocket engines

    Science.gov (United States)

    Immich, H.; Alting, J.; Kretschmer, J.; Preclik, D.

    2003-08-01

    In this paper an overview of recent technology developments for thrust chambers of future launch vehicle liquid rocket engines at Astrium, Space Infrastructure Division (SI), is shown. The main technology. developments shown in this paper are: Technologies Technologies for enhanced heat transfer to the coolant for expander cycle engines Advanced injector head technologies Advanced combustion chamber manufacturing technologies. The main technologies for enhanced heat transfer investigated by subscale chamber hot-firing tests are: Increase of chamber length Hot gas side ribs in the chamber Artificially increased surface roughness. The developments for advanced injector head technologies were focused on the design of a new modular subscale chamber injector head. This injector head allows for an easy exchange of different injection elements: By this, cost effective hot-fire tests with different injection element concepts can be performed. The developments for advanced combustion chamber manufacturing technologies are based on subscale chamber tests with a new design of the Astrium subscale chamber. The subscale chamber has been modified by introduction of a segmented cooled cylindrical section which gives the possibility to test different manufacturing concepts for cooled chamber technologies by exchanging the individual segments. The main technology efforts versus advanced manufacturing technologies shown in this paper are: Soldering techniques Thermal barrier coatings for increased chamber life. A new technology effort is dedicated especially to LOX/Hydrocarbon propellant combinations. Recent hot fire tests on the subscale chamber with Kerosene and Methane as fuel have already been performed. A comprehensive engine system trade-off between the both propellant combinations (Kerosene vs. Methane) is presently under preparation.

  8. Preparation and Ablating Behavior of FGM used in a Heat Flux Rocket Engine

    Science.gov (United States)

    He, Xiaodong; Han, Jiecai; Zhang, Xinghong

    2002-01-01

    rocket engine. As a result, TiB2-Cu FGM showed excellent resistant ablating properties. There is only a little loss of the mass after heated for 40 seconds in the wind tunnel. Meanwhile no cracks and breakup appeared in the FGM under the sharp thermal shock condition. Key words: functionally graded materials, combustion synthesis, ablation, thermal shock, thermal stress

  9. Arc-Heater Facility for Hot Hydrogen Exposure of Nuclear Thermal Rocket Materials

    Science.gov (United States)

    Litchford, Ron J.; Foote, John P.; Wang,Ten-See; Hickman, Robert; Panda, Binayak; Dobson, Chris; Osborne, Robin; Clifton, Scooter

    2006-01-01

    A hyper-thermal environment simulator is described for hot hydrogen exposure of nuclear thermal rocket material specimens and component development. This newly established testing capability uses a high-power, multi-gas, segmented arc-heater to produce high-temperature pressurized hydrogen flows representative of practical reactor core environments and is intended to serve. as a low cost test facility for the purpose of investigating and characterizing candidate fueUstructura1 materials and improving associated processing/fabrication techniques. Design and development efforts are thoroughly summarized, including thermal hydraulics analysis and simulation results, and facility operating characteristics are reported, as determined from a series of baseline performance mapping tests.

  10. Advanced acoustic cavity technology. [for hydrogen oxygen rocket engines

    Science.gov (United States)

    Hines, W. S.; Oberg, C. L.; Kusak, L.

    1974-01-01

    A series of rocket motor firings was performed in a modified linear aerospike thrust chamber with the H2/O2 propellant combination to allow determination of the physical properties of the combustion gases in acoustic cavities located in the chamber side walls. A preliminary analytical study was first conducted to define theoretically both the appropriate cavity dimensions and the combustion gas flow field adjacent to the cavity openings. During the subsequent motor firings, cavity gas temperature profiles were measured and gas samples were withdrawn from the bottom of the cavities for compositional analysis by measurement of pressure/temperature variation and gas chromatography. Data were obtained with both radially and axially oriented cavities and with and without hydrogen bleed flow through the cavities. A simplified procedure was developed for predicting gas cavity and acoustic velocity for use in acoustic cavity design analyses.

  11. High Frequency Combustion Instabilities of LOx/CH4 Spray Flames in Rocket Engine Combustion Chambers

    NARCIS (Netherlands)

    Sliphorst, M.

    2011-01-01

    Ever since the early stages of space transportation in the 1940’s, and the related liquid propellant rocket engine development, combustion instability has been a major issue. High frequency combustion instability (HFCI) is the interaction between combustion and the acoustic field in the combustion

  12. A feasibility study on using inkjet technology, micropumps, and MEMs as fuel injectors for bipropellant rocket engines

    OpenAIRE

    Glynne-Jones, Peter; Coletti, M.; White, N.M.; Gabriel, S.B.; Bramanti, C.

    2010-01-01

    Control over drop size distributions, injection rates, and geometrical distribution of fuel and oxidizer sprays in bi-propellant rocket engines has the potential to produce more efficient, more stable, less polluting rocket engines. This control also offers the potential of an engine that can be throttled, working efficiently over a wide range of output thrusts. Inkjet printing technologies, MEMS fuel atomizers, and piezoelectric injectors similar in concept to those used in diesel engines ar...

  13. Theory of intrachamber processes and design of solid-propellant rocket engines

    Science.gov (United States)

    Erokhin, Boris T.

    The theory of processes taking place inside the combustion chamber of solid-propellant rocket engines is presented, and methods of optimizing the design of such engines are discussed. In particular, attention is given to the selection of the solid propellant charge and igniters, volume density of the charge, structural and heat insulation materials, and thrust vector control devices. Examples of calculations of the nonstationary and quasi-stationary operating regimes and energy and mass characteristics are presented.

  14. Lightweight Exit Cone for Liquid Rocket Engines, Phase I

    Data.gov (United States)

    National Aeronautics and Space Administration — The Pratt and Whitney Rocketdyne (PWR) J-2X engine will power the upper stage of the Ares I and the earth departure stage (EDS) of the Ares V, which will enable...

  15. Development of Advanced Rocket Engine Technology for Precision Guided Missiles

    National Research Council Canada - National Science Library

    Nusca, Michael J; Michaels, R. S

    2004-01-01

    ...) that can power tactical missiles for both current and future combat systems. The use of gel propellants brings the advantages of selectable thrust and the promise of small engine size but also introduces new challenges in combustion control...

  16. Rocket Engine Health Management: Early Definition of Critical Flight Measurements

    Science.gov (United States)

    Christenson, Rick L.; Nelson, Michael A.; Butas, John P.

    2003-01-01

    The NASA led Space Launch Initiative (SLI) program has established key requirements related to safety, reliability, launch availability and operations cost to be met by the next generation of reusable launch vehicles. Key to meeting these requirements will be an integrated vehicle health management ( M) system that includes sensors, harnesses, software, memory, and processors. Such a system must be integrated across all the vehicle subsystems and meet component, subsystem, and system requirements relative to fault detection, fault isolation, and false alarm rate. The purpose of this activity is to evolve techniques for defining critical flight engine system measurements-early within the definition of an engine health management system (EHMS). Two approaches, performance-based and failure mode-based, are integrated to provide a proposed set of measurements to be collected. This integrated approach is applied to MSFC s MC-1 engine. Early identification of measurements supports early identification of candidate sensor systems whose design and impacts to the engine components must be considered in engine design.

  17. Up the Technology Readiness Level (TRL) Scale to Demonstrate a Robust, Long Life, Liquid Rocket Engine Combustion Chamber, or...Up the Downstairs

    Science.gov (United States)

    Holmes, Richard; Elam, Sandra; McKechnie, Timothy; Power, Christopher

    2008-01-01

    Advanced vacuum plasma spray (VPS) technology, utilized to successfully apply thermal barrier coatings to space shuttle main engine turbine blades, was further refined as a functional gradient material (FGM) process for space furnace cartridge experiments at 1600 C and for robust, long life combustion chambers for liquid rocket engines. A VPS/FGM 5K (5,000 lb. thrust) thruster has undergone 220 hot firing tests, in pristine condition, showing no wear, blanching or cooling channel cracks. Most recently, this technology has been applied to a 40K thruster, with scale up planned for a 194K Ares I, J-2X engine.

  18. Fundamentals of aircraft and rocket propulsion

    CERN Document Server

    El-Sayed, Ahmed F

    2016-01-01

    This book provides a comprehensive basics-to-advanced course in an aero-thermal science vital to the design of engines for either type of craft. The text classifies engines powering aircraft and single/multi-stage rockets, and derives performance parameters for both from basic aerodynamics and thermodynamics laws. Each type of engine is analyzed for optimum performance goals, and mission-appropriate engines selection is explained. Fundamentals of Aircraft and Rocket Propulsion provides information about and analyses of: thermodynamic cycles of shaft engines (piston, turboprop, turboshaft and propfan); jet engines (pulsejet, pulse detonation engine, ramjet, scramjet, turbojet and turbofan); chemical and non-chemical rocket engines; conceptual design of modular rocket engines (combustor, nozzle and turbopumps); and conceptual design of different modules of aero-engines in their design and off-design state. Aimed at graduate and final-year undergraduate students, this textbook provides a thorough grounding in th...

  19. Artificial intelligence techniques for ground test monitoring of rocket engines

    Science.gov (United States)

    Ali, Moonis; Gupta, U. K.

    1990-01-01

    An expert system is being developed which can detect anomalies in Space Shuttle Main Engine (SSME) sensor data significantly earlier than the redline algorithm currently in use. The training of such an expert system focuses on two approaches which are based on low frequency and high frequency analyses of sensor data. Both approaches are being tested on data from SSME tests and their results compared with the findings of NASA and Rocketdyne experts. Prototype implementations have detected the presence of anomalies earlier than the redline algorithms that are in use currently. It therefore appears that these approaches have the potential of detecting anomalies early eneough to shut down the engine or take other corrective action before severe damage to the engine occurs.

  20. Combustion and Magnetohydrodynamic Processes in Advanced Pulse Detonation Rocket Engines

    Science.gov (United States)

    2012-10-01

    pressure Qrs Elastic collision cross-section of the rth and sth species Ru Universal gas constant Rprod Specific gas constant (products) Rreac...cross-section and collisional frequency of the sth species, ve is the electron thermal velocity, and f is the electron distribution function. Electron

  1. Transient Mathematical Modeling for Liquid Rocket Engine Systems: Methods, Capabilities, and Experience

    Science.gov (United States)

    Seymour, David C.; Martin, Michael A.; Nguyen, Huy H.; Greene, William D.

    2005-01-01

    The subject of mathematical modeling of the transient operation of liquid rocket engines is presented in overview form from the perspective of engineers working at the NASA Marshall Space Flight Center. The necessity of creating and utilizing accurate mathematical models as part of liquid rocket engine development process has become well established and is likely to increase in importance in the future. The issues of design considerations for transient operation, development testing, and failure scenario simulation are discussed. An overview of the derivation of the basic governing equations is presented along with a discussion of computational and numerical issues associated with the implementation of these equations in computer codes. Also, work in the field of generating usable fluid property tables is presented along with an overview of efforts to be undertaken in the future to improve the tools use for the mathematical modeling process.

  2. Design of a 500 lbf liquid oxygen and liquid methane rocket engine for suborbital flight

    Science.gov (United States)

    Trillo, Jesus Eduardo

    Liquid methane (LCH4)is the most promising rocket fuel for our journey to Mars and other space entities. Compared to liquid hydrogen, the most common cryogenic fuel used today, methane is denser and can be stored at a more manageable temperature; leading to more affordable tanks and a lighter system. The most important advantage is it can be produced from local sources using in-situ resource utilization (ISRU) technology. This will allow the production of the fuel needed to come back to earth on the surface of Mars, or the space entity being explored, making the overall mission more cost effective by enabling larger usable mass. The major disadvantage methane has over hydrogen is it provides a lower specific impulse, or lower rocket performance. The UTEP Center for Space Exploration and Technology Research (cSETR) in partnership with the National Aeronautics and Space Administration (NASA) has been the leading research center for the advancement of Liquid Oxygen (LOX) and Liquid Methane (LCH4) propulsion technologies. Through this partnership, the CROME engine, a throattable 500 lbf LOX/LCH4 rocket engine, was designed and developed. The engine will serve as the main propulsion system for Daedalus, a suborbital demonstration vehicle being developed by the cSETR. The purpose of Daedalus mission and the engine is to fire in space under microgravity conditions to demonstrate its restartability. This thesis details the design process, decisions, and characteristics of the engine to serve as a complete design guide.

  3. Conventional and Bimodal Nuclear Thermal Rocket (NTR) Artificial Gravity Mars Transfer Vehicle Concepts

    Science.gov (United States)

    Borowski, Stanley K.; McCurdy, David R.; Packard, Thomas W.

    2016-01-01

    A variety of countermeasures have been developed to address the debilitating physiological effects of zero-gravity (0-g) experienced by cosmonauts and astronauts during their approximately 0.5 to 1.2 year long stays in low Earth orbit (LEO). Longer interplanetary flights, combined with possible prolonged stays in Mars orbit, could subject crewmembers to up to approximately 2.5 years of weightlessness. In view of known and recently diagnosed problems associated with 0-g, an artificial gravity (AG) spacecraft offers many advantages and may indeed be an enabling technology for human flights to Mars. A number of important human factors must be taken into account in selecting the rotation radius, rotation rate, and orientation of the habitation module or modules. These factors include the gravity gradient effect, radial and tangential Coriolis forces, along with cross-coupled acceleration effects. Artificial gravity Mars transfer vehicle (MTV) concepts are presented that utilize both conventional NTR, as well as, enhanced bimodal nuclear thermal rocket (BNTR) propulsion. The NTR is a proven technology that generates high thrust and has a specific impulse (Isp) capability of approximately 900 s-twice that of today's best chemical rockets. The AG/MTV concepts using conventional Nuclear Thermal Propulsion (NTP) carry twin cylindrical International Space Station (ISS)- type habitation modules with their long axes oriented either perpendicular or parallel to the longitudinal spin axis of the MTV and utilize photovoltaic arrays (PVAs) for spacecraft power. The twin habitat modules are connected to a central operations hub located at the front of the MTV via two pressurized tunnels that provide the rotation radius for the habitat modules. For the BNTR AG/MTV option, each engine has its own closed secondary helium(He)-xenon (Xe) gas loop and Brayton Rotating Unit (BRU) that can generate 10s of kilowatts (kWe) of spacecraft electrical power during the mission coast phase

  4. Application of C/C composites to the combustion chamber of rocket engines. Part 1: Heating tests of C/C composites with high temperature combustion gases

    Science.gov (United States)

    Tadano, Makoto; Sato, Masahiro; Kuroda, Yukio; Kusaka, Kazuo; Ueda, Shuichi; Suemitsu, Takeshi; Hasegawa, Satoshi; Kude, Yukinori

    1995-04-01

    Carbon fiber reinforced carbon composite (C/C composite) has various superior properties, such as high specific strength, specific modulus, and fracture strength at high temperatures of more than 1800 K. Therefore, C/C composite is expected to be useful for many structural applications, such as combustion chambers of rocket engines and nose-cones of space-planes, but C/C composite lacks oxidation resistivity in high temperature environments. To meet the lifespan requirement for thermal barrier coatings, a ceramic coating has been employed in the hot-gas side wall. However, the main drawback to the use of C/C composite is the tendency for delamination to occur between the coating layer on the hot-gas side and the base materials on the cooling side during repeated thermal heating loads. To improve the thermal properties of the thermal barrier coating, five different types of 30-mm diameter C/C composite specimens constructed with functionally gradient materials (FGM's) and a modified matrix coating layer were fabricated. In this test, these specimens were exposed to the combustion gases of the rocket engine using nitrogen tetroxide (NTO) / monomethyl hydrazine (MMH) to evaluate the properties of thermal and erosive resistance on the thermal barrier coating after the heating test. It was observed that modified matrix and coating with FGM's are effective in improving the thermal properties of C/C composite.

  5. Signal Processing Methods for Liquid Rocket Engine Combustion Spontaneous Stability and Rough Combustion Assessments

    Science.gov (United States)

    Kenny, R. Jeremy; Casiano, Matthew; Fischbach, Sean; Hulka, James R.

    2012-01-01

    Liquid rocket engine combustion stability assessments are traditionally broken into three categories: dynamic stability, spontaneous stability, and rough combustion. This work focuses on comparing the spontaneous stability and rough combustion assessments for several liquid engine programs. The techniques used are those developed at Marshall Space Flight Center (MSFC) for the J-2X Workhorse Gas Generator program. Stability assessment data from the Integrated Powerhead Demonstrator (IPD), FASTRAC, and Common Extensible Cryogenic Engine (CECE) programs are compared against previously processed J-2X Gas Generator data. Prior metrics for spontaneous stability assessments are updated based on the compilation of all data sets.

  6. Design Considerations for Human Rating of Liquid Rocket Engines

    Science.gov (United States)

    Parkinson, Douglas

    2010-01-01

    I.Human-rating is specific to each engine; a. Context of program/project must be understood. b. Engine cannot be discussed independently from vehicle and mission. II. Utilize a logical combination of design, manufacturing, and test approaches a. Design 1) It is crucial to know the potential ways a system can fail, and how a failure can propagate; 2) Fault avoidance, fault tolerance, DFMR, caution and warning all have roles to play. b. Manufacturing and Assembly; 1) As-built vs. as-designed; 2) Review procedures for assembly and maintenance periodically; and 3) Keep personnel trained and certified. c. There is no substitute for test: 1) Analytical tools are constantly advancing, but still need test data for anchoring assumptions; 2) Demonstrate robustness and explore sensitivities; 3) Ideally, flight will be encompassed by ground test experience. III. Consistency and repeatability is key in production a. Maintain robust processes and procedures for inspection and quality control based upon development and qualification experience; b. Establish methods to "spot check" quality and consistency in parts: 1) Dedicated ground test engines; 2) Random components pulled from the line/lot to go through "enhanced" testing.

  7. Nuclear Thermal Rocket (Ntr) Propulsion: A Proven Game-Changing Technology for Future Human Exploration Missions

    Science.gov (United States)

    Borowski, Stanley K.; McCurdy, David R.; Packard, Thomas W.

    2012-01-01

    The NTR represents the next evolutionary step in high performance rocket propulsion. It generates high thrust and has a specific impulse (Isp) of approx.900 seconds (s) or more V twice that of today s best chemical rockets. The technology is also proven. During the previous Rover and NERVA (Nuclear Engine for Rocket Vehicle Applications) nuclear rocket programs, 20 rocket reactors were designed, built and ground tested. These tests demonstrated: (1) a wide range of thrust; (2) high temperature carbide-based nuclear fuel; (3) sustained engine operation; (4) accumulated lifetime; and (5) restart capability V all the requirements needed for a human mission to Mars. Ceramic metal cermet fuel was also pursued, as a backup option. The NTR also has significant growth and evolution potential. Configured as a bimodal system, it can generate electrical power for the spacecraft. Adding an oxygen afterburner nozzle introduces a variable thrust and Isp capability and allows bipropellant operation. In NASA s recent Mars Design Reference Architecture (DRA) 5.0 study, the NTR was selected as the preferred propulsion option because of its proven technology, higher performance, lower launch mass, simple assembly and mission operations. In contrast to other advanced propulsion options, NTP requires no large technology scale-ups. In fact, the smallest engine tested during the Rover program V the 25,000 lbf (25 klbf) Pewee engine is sufficient for human Mars missions when used in a clustered engine arrangement. The Copernicus crewed spacecraft design developed in DRA 5.0 has significant capability and a human exploration strategy is outlined here that uses Copernicus and its key components for precursor near Earth asteroid (NEA) and Mars orbital missions prior to a Mars landing mission. Initially, the basic Copernicus vehicle can enable reusable 1-year round trip human missions to candidate NEAs like 1991 JW and Apophis in the late 2020 s to check out vehicle systems. Afterwards, the

  8. Review of fuel element development for nuclear rocket engines

    International Nuclear Information System (INIS)

    Taub, J.M.

    1975-06-01

    The Los Alamos Scientific Laboratory (LASL) entered the nuclear propulsion field in 1955 and began work on all aspects of a nuclear propulsion program involving uranium-loaded graphite fuels, hydrogen propellant, and a target exhaust temperature of approximately 2500 0 C. A very extensive uranium-loaded graphite fuel element technology evolved from the program. Selection and composition of raw materials for the extrusion mix had to be coupled with heat treatment studies to give optimum element properties. The highly enriched uranium in the element was incorporated as UO 2 , pyrocarbon-coated UC 2 , or solid solution UC . ZrC particles. An extensive development program resulted in successful NbC or ZrC coatings on elements to withstand hydrogen corrosion at elevated temperatures. Hot gas, thermal shock, thermal stress, and NDT evaluation procedures were developed to monitor progress in preparation of elements with optimum properties. Final evaluation was made in reactor tests at NRDS. Aerojet-General, Westinghouse Astronuclear Laboratory, and the Oak Ridge Y-12 Plant of Union Carbide Nuclear Company entered the program in the early 1960's, and their activities paralleled those of LASL in fuel element development. (U.S.)

  9. Analysis of startup strategies for a particle bed reactor nuclear rocket engine

    Science.gov (United States)

    Suzuki, D. E.

    1993-06-01

    This paper develops and analyzes engine system startup strategies for a particle bed reactor (PBR) nuclear rocket engine. The strategies are designed to maintain stable flow through the PBR fuel element while reaching the design conditions as quickly as possible. The analyses are conducted using a computer model of a representative particle bed reactor and engine system. Elements of the startup strategy considered include: the coordinated control of reactor power and coolant flow; turbine inlet temperature and flow control; and use of an external starter system. The simulation results indicate that the use of an external starter system enables the engine to reach design conditions very quickly while maintaining the flow well away from the unstable regime. If a bootstrap start is used instead, the transient does not progress as fast and approaches closer to the unstable flow regime, but allows for greater engine reusability. These results can provide important information for engine designers and mission planners.

  10. Development of test stand for experimental investigation of chemical and physical phenomena in Liquid Rocket Engine

    Directory of Open Access Journals (Sweden)

    Emerson Andrade Santos

    2011-05-01

    Full Text Available The main objective of this work was to present the specification of an experimental firing test stand for liquid rocket engines (LRE and develop a program for control and acquisition of data. It provides conditions to test rocket engines with thrust from 50 to 100 kgf. A methodology for laboratory work implementation using information technology, which will allow the automatic and remote functioning of the test stand, permits users to input the necessary data to conduct tests safely, achieve accurate measurements and obtain reliable results. The control of propellant mass flow rates by pressure regulators and other system valves, as well as the test stand data acquisition, are carried out automatically through LabVIEW commercial software. The test stand program is a readable, scalable and maintainable code. The test stand design and its development represent the state of art of experimental apparatus in LRE testing.

  11. Signal Processing Methods for Liquid Rocket Engine Combustion Stability Assessments

    Science.gov (United States)

    Kenny, R. Jeremy; Lee, Erik; Hulka, James R.; Casiano, Matthew

    2011-01-01

    The J2X Gas Generator engine design specifications include dynamic, spontaneous, and broadband combustion stability requirements. These requirements are verified empirically based high frequency chamber pressure measurements and analyses. Dynamic stability is determined with the dynamic pressure response due to an artificial perturbation of the combustion chamber pressure (bomb testing), and spontaneous and broadband stability are determined from the dynamic pressure responses during steady operation starting at specified power levels. J2X Workhorse Gas Generator testing included bomb tests with multiple hardware configurations and operating conditions, including a configuration used explicitly for engine verification test series. This work covers signal processing techniques developed at Marshall Space Flight Center (MSFC) to help assess engine design stability requirements. Dynamic stability assessments were performed following both the CPIA 655 guidelines and a MSFC in-house developed statistical-based approach. The statistical approach was developed to better verify when the dynamic pressure amplitudes corresponding to a particular frequency returned back to pre-bomb characteristics. This was accomplished by first determining the statistical characteristics of the pre-bomb dynamic levels. The pre-bomb statistical characterization provided 95% coverage bounds; these bounds were used as a quantitative measure to determine when the post-bomb signal returned to pre-bomb conditions. The time for post-bomb levels to acceptably return to pre-bomb levels was compared to the dominant frequency-dependent time recommended by CPIA 655. Results for multiple test configurations, including stable and unstable configurations, were reviewed. Spontaneous stability was assessed using two processes: 1) characterization of the ratio of the peak response amplitudes to the excited chamber acoustic mode amplitudes and 2) characterization of the variability of the peak response

  12. Development and Short-Range Testing of a 100 kW Side-Illuminated Millimeter-Wave Thermal Rocket

    Science.gov (United States)

    Bruccoleri, Alexander; Eilers, James A.; Lambot, Thomas; Parkin, Kevin

    2015-01-01

    The objective of the phase described here of the Millimeter-Wave Thermal Launch System (MTLS) Project was to launch a small thermal rocket into the air using millimeter waves. The preliminary results of the first MTLS flight vehicle launches are presented in this work. The design and construction of a small thermal rocket with a planar ceramic heat exchanger mounted along the axis of the rocket is described. The heat exchanger was illuminated from the side by a millimeter-wave beam and fed propellant from above via a small tank containing high pressure argon or nitrogen. Short-range tests where the rocket was launched, tracked, and heated with the beam are described. The rockets were approximately 1.5 meters in length and 65 millimeters in diameter, with a liftoff mass of 1.8 kilograms. The rocket airframes were coated in aluminum and had a parachute recovery system activated via a timer and Pyrodex. At the rocket heat exchanger, the beam distance was 40 meters with a peak power intensity of 77 watts per square centimeter. and a total power of 32 kilowatts in a 30 centimeter diameter circle. An altitude of approximately 10 meters was achieved. Recommendations for improvements are discussed.

  13. Integrated Ceramic Matrix Composite and Carbon/Carbon Structures for Large Rocket Engine Nozzles and Nozzle Extensions Project

    Data.gov (United States)

    National Aeronautics and Space Administration — Low-cost access to space demands durable, cost-effective, efficient, and low-weight propulsion systems. Key components include rocket engine nozzles and nozzle...

  14. Bi-Metallic Additive Manufacturing Close-Out of Coolant Channels for Large Liquid Rocket Engine (LRE) Nozzles, Phase I

    Data.gov (United States)

    National Aeronautics and Space Administration — This NASA sponsored STTR project will investigate methods for close-out of large, liquid rocket engine nozzle, coolant channels utilizing robotic laser and...

  15. Paraffin-based hybrid rocket engines applications: A review and a market perspective

    Science.gov (United States)

    Mazzetti, Alessandro; Merotto, Laura; Pinarello, Giordano

    2016-09-01

    Hybrid propulsion technology for aerospace applications has received growing attention in recent years due to its important advantages over competitive solutions. Hybrid rocket engines have a great potential for several aeronautics and aerospace applications because of their safety, reliability, low cost and high performance. As a consequence, this propulsion technology is feasible for a number of innovative missions, including space tourism. On the other hand, hybrid rocket propulsion's main drawback, i.e. the difficulty in reaching high regression rate values using standard fuels, has so far limited the maturity level of this technology. The complex physico-chemical processes involved in hybrid rocket engines combustion are of major importance for engine performance prediction and control. Therefore, further investigation is ongoing in order to achieve a more complete understanding of such phenomena. It is well known that one of the most promising solutions for overcoming hybrid rocket engines performance limits is the use of liquefying fuels. Such fuels can lead to notably increased solid fuel regression rate due to the so-called "entrainment phenomenon". Among liquefying fuels, paraffin-based formulations have great potentials as solid fuels due to their low cost, availability (as they can be derived from industrial waste), low environmental impact and high performance. Despite the vast amount of literature available on this subject, a precise focus on market potential of paraffins for hybrid propulsion aerospace applications is lacking. In this work a review of hybrid rocket engines state of the art was performed, together with a detailed analysis of the possible applications of such a technology. A market study was carried out in order to define the near-future foreseeable development needs for hybrid technology application to the aforementioned missions. Paraffin-based fuels are taken into account as the most promising segment for market development

  16. Nuclear Thermal Rocket Element Environmental Simulator (NTREES) Phase II Upgrade Activities

    Science.gov (United States)

    Emrich, William J.; Moran, Robert P.; Pearson, J. Bose

    2013-01-01

    To support the on-going nuclear thermal propulsion effort, a state-of-the-art non nuclear experimental test setup has been constructed to evaluate the performance characteristics of candidate fuel element materials and geometries in representative environments. The facility to perform this testing is referred to as the Nuclear Thermal Rocket Element Environment Simulator (NTREES). This device can simulate the environmental conditions (minus the radiation) to which nuclear rocket fuel components will be subjected during reactor operation. Test articles mounted in the simulator are inductively heated in such a manner so as to accurately reproduce the temperatures and heat fluxes which would normally occur as a result of nuclear fission and would be exposed to flowing hydrogen. Initial testing of a somewhat prototypical fuel element has been successfully performed in NTREES and the facility has now been shutdown to allow for an extensive reconfiguration of the facility which will result in a significant upgrade in its capabilities. Keywords: Nuclear Thermal Propulsion, Simulator

  17. Preliminary design of a pressurization system for small bipropellant rocket engines

    Science.gov (United States)

    Stanley, Steven

    A study was conducted on the feasibility of developing a device or system that would improve the performance of small, bipropellant rockets through pressurization of the propellants. Due to the limitations in the space industry, namely high development costs and resistance to change, the new approach needed to be as simple and robust as possible. After reviewing several different potential methodologies, a concept was developed from first principles based on small gas turbine engine fuel injection approaches. The concept is simple and has heritage in the field of gas turbine engines, but it is new for the field of rocket propulsion. Using the basic physics of the proposed baseline concept, a simulation was developed to optimize the design parameters and to explore the trade space. Exercising the resulting simulation led to the identification of the critical design parameters and key performance metrics. During the iteration process, the design was updated and finalized. The resulting configuration appears to be feasible and has the potential of providing a new capability for small bipropellant rockets. Based upon the results of the study, recommendations were developed and a plan was created to further the development of the pump.

  18. New Frontiers AO: Advanced Materials Bi-propellant Rocket (AMBR) Engine Information Summary

    Science.gov (United States)

    Liou, Larry C.

    2008-01-01

    The Advanced Material Bi-propellant Rocket (AMBR) engine is a high performance (I(sub sp)), higher thrust, radiation cooled, storable bi-propellant space engine of the same physical envelope as the High Performance Apogee Thruster (HiPAT(TradeMark)). To provide further information about the AMBR engine, this document provides details on performance, development, mission implementation, key spacecraft integration considerations, project participants and approach, contact information, system specifications, and a list of references. The In-Space Propulsion Technology (ISPT) project team at NASA Glenn Research Center (GRC) leads the technology development of the AMBR engine. Their NASA partners were Marshall Space Flight Center (MSFC) and Jet Propulsion Laboratory (JPL). Aerojet leads the industrial partners selected competitively for the technology development via the NASA Research Announcement (NRA) process.

  19. Micro-thermal engineering; Microthermique

    Energy Technology Data Exchange (ETDEWEB)

    Pelzi, J.; Chotikaprakhan, S.; Fotsing, J.L.N.; Dietzel, D.; Meckenstock, R. [Ruhr Univ., Experimental Physics 3, Solid State Spectroscopy, Bochum (Germany); Cassette, S. [Thales Research and Technology France, 91 - Orsay (France); Polignano, M.L.; Mica, I. [STMicroelectronics, Agrate Brianza (Italy); Lallemand, M. [Institut National des Sciences Appliquees (INSA), CETHIL, UMR 5008 INSA-UCB-CNRS, 69 - Villeurbanne (France); Ayela, F. [CRTBT, UPR CRNS 5001, 38 - Grenoble (France); Favre-Marine, M.; Marty, Ph. [Institut National Polytechnique, LEGI, UMR 5519 INPG-UJF-CNRS, 38 - Grenoble (France); Gruss, A. [CEA Grenoble, GRETh, 38 (France); Maillet, D. [LEMTA, UMR 7563 ENSEM-CNRS, 54 - Vandoeuvre (France); Peerhossaini, H. [LT, UMR 6607 EPUN-CNRS, 44 - Nantes (France); Tadrist, L. [IUSTI, UMR 6595 EPU-CNRS, UMR 6595, 13 - Marseille (France); Charlot, B. [TIMA, Techniques de l' Informatique et de la Microelectronique pour l' Architecture d' Ordinateurs, 38 - Grenoble (France)

    2005-07-01

    This session about heat transfers in micro-systems gathers 3 articles dealing with: the thermal characterization of electronic components by thermal microscopies; the heat transfers in micro-channels and the application to micro-exchangers; and the heat transfers in micro-electromechanical systems. (J.S.)

  20. Laboratory Facilities for Testing Thermal Engines

    Directory of Open Access Journals (Sweden)

    Ioan Ruja

    2010-10-01

    Full Text Available This work presents an electromechanical plant through with which is realised couples different resistant, MR (0 ÷ MRN, on the gearbox shaft of internal combustion engine. The purpose is to study the plant in phase and stationary behaviour of the main technical parameters that define the engine operation such as: torque, speed, temperature, pressure, vibration, burnt gas, noise, forces. You can take measurements to determine engine performance testing and research on improving engine thermal efficiency. With the proposed plant is built by measuring the characteristic internal combustion engines (tuning characteristic and functional characteristic and determine the technical performance of interest, optimal.

  1. Design and test of a small two stage counter-rotating turbine for rocket engine application

    Science.gov (United States)

    Huber, F. W.; Branstrom, B. R.; Finke, A. K.; Johnson, P. D.; Rowey, R. J.; Veres, J. P.

    1993-01-01

    The aerodynamic design and rig test evaluation of a small counter-rotating turbine system is described. The technology represented by this turbine is being developed for application in an advanced upper stage rocket engine turbopump. This engine will employ an oxygen/hydrogen expander cycle and achieve high performance through efficient combustion, high combustion pressure, and high area ratio exhaust nozzle expansion. Engine performance goals require that the turbopump drive turbines achieve high efficiency at low gas flow rates. The low flow rates result in very small airfoil diameter, height and chord. The high efficiency and small size requirements present a challenging turbine design problem. The unconventional approach employed to meet this challenge is described, along with the detailed design process and resulting airfoil configurations. The method and results of full scale aerodynamic performance evaluation testing of both one and two stage configurations, as well as operation without the secondary stage stator are presented. The overall results of this effort illustrate that advanced aerodynamic design tools and hardware fabrication techniques have provided improved capability to produce small high performance turbines for advanced rocket engines.

  2. Coupled Lagrangian impingement spray model for doublet impinging injectors under liquid rocket engine operating conditions

    Directory of Open Access Journals (Sweden)

    Qiang WEI

    2017-08-01

    Full Text Available To predict the effect of the liquid rocket engine combustion chamber conditions on the impingement spray, the conventional uncoupled spray model for impinging injectors is extended by considering the coupling of the jet impingement process and the ambient gas field. The new coupled model consists of the plain-orifice sub-model, the jet-jet impingement sub-model and the droplet collision sub-model. The parameters of the child droplet are determined with the jet-jet impingement sub-model using correlations about the liquid jet parameters and the chamber conditions. The overall model is benchmarked under various impingement angles, jet momentum and off-center ratios. Agreement with the published experimental data validates the ability of the model to predict the key spray characteristics, such as the mass flux and mixture ratio distributions in quiescent air. Besides, impinging sprays under changing ambient pressure and non-uniform gas flow are investigated to explore the effect of liquid rocket engine chamber conditions. First, a transient impingement spray during engine start-up phase is simulated with prescribed pressure profile. The minimum average droplet diameter is achieved when the orifices work in cavitation state, and is about 30% smaller than the steady single phase state. Second, the effect of non-uniform gas flow produces off-center impingement and the rotated spray fan by 38°. The proposed model suggests more reasonable impingement spray characteristics than the uncoupled one and can be used as the first step in the complex simulation of coupling impingement spray and combustion in liquid rocket engines.

  3. A feasibility study on using inkjet technology, micropumps, and MEMs as fuel injectors for bipropellant rocket engines

    Science.gov (United States)

    Glynne-Jones, Peter; Coletti, M.; White, N. M.; Gabriel, S. B.; Bramanti, C.

    2010-07-01

    Control over drop size distributions, injection rates, and geometrical distribution of fuel and oxidizer sprays in bi-propellant rocket engines has the potential to produce more efficient, more stable, less polluting rocket engines. This control also offers the potential of an engine that can be throttled, working efficiently over a wide range of output thrusts. Inkjet printing technologies, MEMS fuel atomizers, and piezoelectric injectors similar in concept to those used in diesel engines are considered for their potential to yield a new, more active injection scheme for a rocket engine. Inkjets are found to be unable to pump at sufficient pressures, and have possibly dangerous failure modes. Active injection is found to be feasible if high pressure drop along the injector plate is used. A conceptual design is presented and its basic behavior assessed.

  4. Optical Measurement Techniques for Rocket Engine Testing and Component Applications: Digital Image Correlation and Dynamic Photogrammetry

    Science.gov (United States)

    Gradl, Paul

    2016-01-01

    NASA Marshall Space Flight Center (MSFC) has been advancing dynamic optical measurement systems, primarily Digital Image Correlation, for extreme environment rocket engine test applications. The Digital Image Correlation (DIC) technology is used to track local and full field deformations, displacement vectors and local and global strain measurements. This technology has been evaluated at MSFC through lab testing to full scale hotfire engine testing of the J-2X Upper Stage engine at Stennis Space Center. It has been shown to provide reliable measurement data and has replaced many traditional measurement techniques for NASA applications. NASA and AMRDEC have recently signed agreements for NASA to train and transition the technology to applications for missile and helicopter testing. This presentation will provide an overview and progression of the technology, various testing applications at NASA MSFC, overview of Army-NASA test collaborations and application lessons learned about Digital Image Correlation.

  5. Optimization of the stand for test of hybrid rocket engines of solid fuel

    Directory of Open Access Journals (Sweden)

    Zolotorev Nikolay

    2017-01-01

    Full Text Available In the paper the laboratory experimental stand of the hybrid rocket engine of solid fuel to study ballistic parameters of the engine at burning of high-energy materials in flow of hot gas is presented. Mixture of air with nitrogen with a specified content of active oxygen is used as a gaseous oxidizer. The experimental stand has modular design and consists of system of gas supply, system of heating of gas, system for monitoring gas parameters, to which a load cell with a model engine was connected. The modular design of the stand allows to change its configuration under specific objective. This experimental stand allows to conduct a wide range of the pilot studies at interaction of a hot stream of gas with samples high-energy materials.

  6. Development of an intelligent diagnostic system for reusable rocket engine control

    Science.gov (United States)

    Anex, R. P.; Russell, J. R.; Guo, T.-H.

    1991-01-01

    A description of an intelligent diagnostic system for the Space Shuttle Main Engines (SSME) is presented. This system is suitable for incorporation in an intelligent controller which implements accommodating closed-loop control to extend engine life and maximize available performance. The diagnostic system architecture is a modular, hierarchical, blackboard system which is particularly well suited for real-time implementation of a system which must be repeatedly updated and extended. The diagnostic problem is formulated as a hierarchical classification problem in which the failure hypotheses are represented in terms of predefined data patterns. The diagnostic expert system incorporates techniques for priority-based diagnostics, the combination of analytical and heuristic knowledge for diagnosis, integration of different AI systems, and the implementation of hierarchical distributed systems. A prototype reusable rocket engine diagnostic system (ReREDS) has been implemented. The prototype user interface and diagnostic performance using SSME test data are described.

  7. Development of Aluminum Composites for a Rocket Engine's Lightweight Thrust Cell

    Science.gov (United States)

    Lee, Jonathan A.; Elam, Sandy; Munafo, Paul M. (Technical Monitor)

    2001-01-01

    The Aerospike liquid fueled rocket engine for the X-33 aerospace vehicle consists of several thrust cells, which can comprise as much as 25% of the engine weight. The interior wall of the thrust cell chamber is exposed to high temperature combustion products and must be cooled by using liquid hydrogen. Ultimately, reducing engine weight would increase vehicle performance and allow heavier payload capabilities. Currently, the thrust cell's structural jacket and manifolds are made of stainless steel 347, which can potentially be replaced by a lighter material such as an Aluminum (Al) Metal Matrix Composites (MMC). Up to 50% weight reduction can be expected for each of the thrust cell chambers using particulate SiC reinforced Al MMC. Currently, several Al MMC thrust cell structural jackets have been produced, using cost-effective processes such as gravity casting and plasma spray deposition, to demonstrate MMC technology readiness for NASA's advanced propulsion systems.

  8. Application of High Speed Digital Image Correlation in Rocket Engine Hot Fire Testing

    Science.gov (United States)

    Gradl, Paul R.; Schmidt, Tim

    2016-01-01

    Hot fire testing of rocket engine components and rocket engine systems is a critical aspect of the development process to understand performance, reliability and system interactions. Ground testing provides the opportunity for highly instrumented development testing to validate analytical model predictions and determine necessary design changes and process improvements. To properly obtain discrete measurements for model validation, instrumentation must survive in the highly dynamic and extreme temperature application of hot fire testing. Digital Image Correlation has been investigated and being evaluated as a technique to augment traditional instrumentation during component and engine testing providing further data for additional performance improvements and cost savings. The feasibility of digital image correlation techniques were demonstrated in subscale and full scale hotfire testing. This incorporated a pair of high speed cameras to measure three-dimensional, real-time displacements and strains installed and operated under the extreme environments present on the test stand. The development process, setup and calibrations, data collection, hotfire test data collection and post-test analysis and results are presented in this paper.

  9. Modular simulation software development for liquid propellant rocket engines based on MATLAB Simulink

    Directory of Open Access Journals (Sweden)

    Naderi Mahyar

    2017-01-01

    Full Text Available Focusing on Liquid Propellant Rocket Engine (LPRE major components, a steady state modular simulation software has been established in MATLAB Simulink. For integrated system analysis, a new algorithm dependant on engine inlet mass flow rate and pressure is considered. Using the suggested algorithm, it is possible to evaluate engine component operation, similar to the known initial parameters during hot fire test of the engine on stand. As a case study, the reusable Space Shuttle Main Engine (SSME has been selected and the simulation has been performed to predict engine’s throttled operation at 109 percent of the nominal thrust value. For this purpose the engine actual flow diagram has been converted to 34 numerical modules and the engine has been modelled by solving a total of 101 steady state mathematic equations. The mean error of the simulation results is found to be less than 5% compared with the published SSME data. Using the presented idea and developed modules, it is possible to build up the numerical model and simulate other LPREs.

  10. Operation of a cryogenic rocket engine an outline with down-to-earth and up-to-space remarks

    CERN Document Server

    Kitsche, Wolfgang

    2010-01-01

    This book presents the operational aspects of the rocket engine on a test facility. It will be useful to engineers and scientists who are in touch with the test facility. To aerospace students it shall provide an insight of the job on the test facility. And to interest readers it shall provide an impression of this thrilling area of aerospace.

  11. Thermal integrity in mechanics and engineering

    CERN Document Server

    Shorr, Boris F

    2015-01-01

    The book is targeted at engineers, university lecturers, postgraduates, and final year undergraduate students involved in computational modelling and experimental and theoretical analysis of the high-temperature behavior of engineering structures. It will also be of interest to researchers developing the thermal strength theory as a branch of continuum mechanics. Thermal integrity is a multidisciplinary field combining the expertise of mechanical engineers, material scientists and applied mathematicians, each approaching the problem from their specific viewpoint. This monograph draws on the research of a broad scientific community including the author’s contribution. The scope of thermal strength analysis was considerably extended thanks to modern computers and the implementation of FEM codes. However, the author believes that some material models adopted in the advanced high-performance software, are not sufficiently justificated due to lack of easy-to-follow books on the theoretical and experimental aspec...

  12. Influence of atomization quality modulation on flame dynamics in a hypergolic rocket engine

    Directory of Open Access Journals (Sweden)

    Moritz Schulze

    2016-09-01

    Full Text Available For the numerical evaluation of the thermoacoustic stability of rocket engines often hybrid methods are applied, which separate the computation of wave propagation in the combustor from the analysis of the flame response to acoustic perturbations. Closure requires a thermoacoustic feedback model which provides the heat release fluctuation in the source term of the employed wave transport equations. The influence of the acoustic fluctuations in the combustion chamber on the heat release fluctuations from the modulation of the atomization of the propellants in a hypergolic upper stage rocket engine is studied. Numerical modeling of a single injector provides the time mean reacting flow field. A network of transfer functions representing all aspects relevant for the feedback model is presented. Analytical models for the injector admittances and for the atomization transfer functions are provided. The dynamics of evaporation and combustion are studied numerically and the numerical results are analyzed. An analytical approximation of the computed flame transfer function is combined with the analytical models for the injector and the atomization quality to derive the feedback model for the wave propagation code. The evaluation of this model on the basis of the Rayleigh index reveals the thermoacoustic driving potential originating from the fluctuating spray quality.

  13. Engine Cycle Analysis of Air Breathing Microwave Rocket with Reed Valves

    Science.gov (United States)

    Fukunari, Masafumi; Komatsu, Reiji; Yamaguchi, Toshikazu; Komurasaki, Kimiya; Arakawa, Yoshihiro; Katsurayama, Hiroshi

    2011-11-01

    The Microwave Rocket is a candidate for a low cost launcher system. Pulsed plasma generated by a high power millimeter wave beam drives a blast wave, and a vehicle acquires impulsive thrust by exhausting the blast wave. The thrust generation process of the Microwave Rocket is similar to a pulse detonation engine. In order to enhance the performance of its air refreshment, the air-breathing mechanism using reed valves is under development. Ambient air is taken to the thruster through reed valves. Reed valves are closed while the inside pressure is high enough. After the time when the shock wave exhausts at the open end, an expansion wave is driven and propagates to the thrust-wall. The reed valve is opened by the negative gauge pressure induced by the expansion wave and its reflection wave. In these processes, the pressure oscillation is important parameter. In this paper, the pressure oscillation in the thruster was calculated by CFD combined with the flux through from reed valves, which is estimated analytically. As a result, the air-breathing performance is evaluated using Partial Filling Rate (PFR), the ratio of thruster length to diameter L/D, and ratio of opening area of reed valves to superficial area α. An engine cycle and predicted thrust was explained.

  14. Subcontinuum thermal transport in tip-based thermal engineering

    Science.gov (United States)

    Hamian, Sina

    For the past two decades, tip-based thermal engineering has made remarkable advances to realize unprecedented nanoscale thermal applications, such as thermomechanical data storage, thermophysical/chemical property characterization of materials in nanometer scale, and scanning thermal imaging and analysis. All these applications involve localized heating with elevated temperature, generally in the order of mean free paths of heat carriers, thus necessitates fundamental understanding of sub-continuum thermal transport across point constrictions and within thin films. Considering the demands, this dissertation is divided into three main scopes providing: (1) a numerical model that provides insight onto nanoscale thermal transport, (2) an electrothermal characterization of a heated microcantilever as a localized heating source, and (3) qualitative measurement of tip-substrate thermal transport using high resolution nanothermometer/heater. This dissertation starts with a literature review on the three aforementioned scopes followed by a numerical model for two-dimensional transient ballistic-diffusive heat transfer combining finite element analysis with discrete ordinate method (DOM-FEA), seeking to provide insight on subcontinuum thermal transport. The phonon Boltzmann transport equation (BTE) under grey relaxation time approximation is solved for different Knudsen numbers. Next, a thermal microcantilever, as one of the main tools in tip-based thermal engineering, is characterized under periodic heating operation in air and vacuum using 3o technique. A three-dimensional FEA simulation of a thermal microcantilever is used to model heat transfer in frequency domain resulting in good agreement with the experiment. Next, quantitative thermal transport is measured by a home-built nanothermometer fabricated using combination of electron-beam lithography and photolithography. An atomic force microscope (AFM) cantilever is used to scan over the sensing probe of the

  15. Digital Image Correlation Techniques Applied to Large Scale Rocket Engine Testing

    Science.gov (United States)

    Gradl, Paul R.

    2016-01-01

    Rocket engine hot-fire ground testing is necessary to understand component performance, reliability and engine system interactions during development. The J-2X upper stage engine completed a series of developmental hot-fire tests that derived performance of the engine and components, validated analytical models and provided the necessary data to identify where design changes, process improvements and technology development were needed. The J-2X development engines were heavily instrumented to provide the data necessary to support these activities which enabled the team to investigate any anomalies experienced during the test program. This paper describes the development of an optical digital image correlation technique to augment the data provided by traditional strain gauges which are prone to debonding at elevated temperatures and limited to localized measurements. The feasibility of this optical measurement system was demonstrated during full scale hot-fire testing of J-2X, during which a digital image correlation system, incorporating a pair of high speed cameras to measure three-dimensional, real-time displacements and strains was installed and operated under the extreme environments present on the test stand. The camera and facility setup, pre-test calibrations, data collection, hot-fire test data collection and post-test analysis and results are presented in this paper.

  16. Nuclear Cryogenic Propulsion Stage (NCPS) Fuel Element Testing in the Nuclear Thermal Rocket Element Environmental Simulator (NTREES)

    Science.gov (United States)

    Emrich, William J., Jr.

    2017-01-01

    To support the on-going nuclear thermal propulsion effort, a state-of-the-art non nuclear experimental test setup has been constructed to evaluate the performance characteristics of candidate fuel element materials and geometries in representative environments. The facility to perform this testing is referred to as the Nuclear Thermal Rocket Element Environment Simulator (NTREES). Last year NTREES was successfully used to satisfy a testing milestone for the Nuclear Cryogenic Propulsion Stage (NCPS) project and met or exceeded all required objectives.

  17. Ultraviolet photographic pyrometer used in rocket exhaust analysis

    Science.gov (United States)

    Levin, B. P.

    1966-01-01

    Ultraviolet photographic pyrometer investigates the role of carbon as a thermal radiator and determines the geometry, location, and progress of afterburning phenomena in the exhaust plume of rocket engines using liquid oxygen/RP-1 as propellant.

  18. Cooling Duct Analysis for Transpiration/Film Cooled Liquid Propellant Rocket Engines

    Science.gov (United States)

    Micklow, Gerald J.

    1996-01-01

    The development of a low cost space transportation system requires that the propulsion system be reusable, have long life, with good performance and use low cost propellants. Improved performance can be achieved by operating the engine at higher pressure and temperature levels than previous designs. Increasing the chamber pressure and temperature, however, will increase wall heating rates. This necessitates the need for active cooling methods such as film cooling or transpiration cooling. But active cooling can reduce the net thrust of the engine and add considerably to the design complexity. Recently, a metal drawing process has been patented where it is possible to fabricate plates with very small holes with high uniformity with a closely specified porosity. Such a metal plate could be used for an inexpensive transpiration/film cooled liner to meet the demands of advanced reusable rocket engines, if coolant mass flow rates could be controlled to satisfy wall cooling requirements and performance. The present study investigates the possibility of controlling the coolant mass flow rate through the porous material by simple non-active fluid dynamic means. The coolant will be supplied to the porous material by series of constant geometry slots machined on the exterior of the engine.

  19. Investigation of a Tricarbide Grooved Ring Fuel Element for a Nuclear Thermal Rocket

    Science.gov (United States)

    Taylor, Brian D.; Emrich, Bill; Tucker, Dennis; Barnes, Marvin; Donders, Nicolas; Benensky, Kelsa

    2017-01-01

    Deep space exploration, especially that of Mars, is on the horizon as the next big challenge for space exploration. Nuclear propulsion, through which high thrust and efficiency can be achieved, is a promising option for decreasing the cost and logistics of such a mission. Work on nuclear thermal engines goes back to the days of the NERVA program. Currently, nuclear thermal propulsion is under development again in various forms to provide a superior propulsion system for deep space exploration. The authors have been working to develop a concept nuclear thermal engine that uses a grooved ring fuel element as an alternative to the traditional hexagonal rod design. The authors are also studying the use of carbide fuels. The concept was developed in order to increase surface area and heat transfer to the propellant. The use of carbides would also raise the temperature limitations of the reactor. It is hoped that this could lead to a higher thrust to weight nuclear thermal engine. This paper describes the modeling of neutronics, heat transfer, and fluid dynamics of this alternative nuclear fuel element geometry. Fabrication experiments of grooved rings from carbide refractory metals are also presented along with material characterization and interactions with a hot hydrogen environment.

  20. Extensions to the time lag models for practical application to rocket engine stability design

    Science.gov (United States)

    Casiano, Matthew J.

    The combustion instability problem in liquid-propellant rocket engines (LREs) has remained a tremendous challenge since their discovery in the 1930s. Improvements are usually made in solving the combustion instability problem primarily using computational fluid dynamics (CFD) and also by testing demonstrator engines. Another approach is to use analytical models. Analytical models can be used such that design, redesign, or improvement of an engine system is feasible in a relatively short period of time. Improvements to the analytical models can greatly aid in design efforts. A thorough literature review is first conducted on liquid-propellant rocket engine (LRE) throttling. Throttling is usually studied in terms of vehicle descent or ballistic missile control however there are many other cases where throttling is important. It was found that combustion instabilities are one of a few major issues that occur during deep throttling (other major issues are heat transfer concerns, performance loss, and pump dynamics). In the past and again recently, gas injected into liquid propellants has shown to be a viable solution to throttle engines and to eliminate some forms of combustion instability. This review uncovered a clever solution that was used to eliminate a chug instability in the Common Extensible Cryogenic Engine (CECE), a modified RL10 engine. A separate review was also conducted on classic time lag combustion instability models. Several new stability models are developed by incorporating important features to the classic and contemporary models, which are commonly used in the aerospace rocket industry. The first two models are extensions of the original Crocco and Cheng concentrated combustion model with feed system contributions. A third new model is an extension to the Wenzel and Szuch double-time lag model also with feed system contributions. The first new model incorporates the appropriate injector acoustic boundary condition which is neglected in contemporary

  1. Thermal integrity in mechanics and engineering

    International Nuclear Information System (INIS)

    Shorr, Boris F.

    2015-01-01

    The book is targeted at engineers, university lecturers, postgraduates, and final year undergraduate students involved in computational modelling and experimental and theoretical analysis of the high-temperature behavior of engineering structures. It will also be of interest to researchers developing the thermal strength theory as a branch of continuum mechanics. Thermal integrity is a multidisciplinary field combining the expertise of mechanical engineers, material scientists and applied mathematicians, each approaching the problem from their specific viewpoint. This monograph draws on the research of a broad scientific community including the author's contribution. The scope of thermal strength analysis was considerably extended thanks to modern computers and the implementation of FEM codes. However, the author believes that some material models adopted in the advanced high-performance software, are not sufficiently justificated due to lack of easy-to-follow books on the theoretical and experimental aspects of thermal integrity. The author endeavors to provide a thorough yet sufficiently simple presentation of the underlying concepts, making the book compelling to a wide audience.

  2. Thermal integrity in mechanics and engineering

    Energy Technology Data Exchange (ETDEWEB)

    Shorr, Boris F. [Central Institute of Aviation Motors (CIAM), Moscow (Russian Federation)

    2015-07-01

    The book is targeted at engineers, university lecturers, postgraduates, and final year undergraduate students involved in computational modelling and experimental and theoretical analysis of the high-temperature behavior of engineering structures. It will also be of interest to researchers developing the thermal strength theory as a branch of continuum mechanics. Thermal integrity is a multidisciplinary field combining the expertise of mechanical engineers, material scientists and applied mathematicians, each approaching the problem from their specific viewpoint. This monograph draws on the research of a broad scientific community including the author's contribution. The scope of thermal strength analysis was considerably extended thanks to modern computers and the implementation of FEM codes. However, the author believes that some material models adopted in the advanced high-performance software, are not sufficiently justificated due to lack of easy-to-follow books on the theoretical and experimental aspects of thermal integrity. The author endeavors to provide a thorough yet sufficiently simple presentation of the underlying concepts, making the book compelling to a wide audience.

  3. Computational Analysis of an LOx Supply Line System of an Liquid Rocket Engine

    Directory of Open Access Journals (Sweden)

    Insang Moon

    2009-12-01

    Full Text Available A computational fluid analysis was performed on an LOx line system of a liquid rocket engine. The model was created with 3D CAD and imbedded to the 3D CFD program. Before the full scale analysis on the system was carried out, each components with simplified models was analyzed to save time and cost. As a result, the inlet pressure of the gas generator should be compensated with a certain device unless the inlet pressure of the line system is sufficiently high. The flow pattern of the exit of the system was dependant upon the location of the orifice as well as the size. As a whole the line system analyzed met the requirements, and will be tested and confirmed after being manufactured.

  4. Multidimensional Unstructured-Grid Liquid Rocket Engine Nozzle Performance and Heat Transfer Analysis

    Science.gov (United States)

    Wang, Ten-See

    2004-01-01

    The objective of this study is to conduct a unified computational analysis for computing design parameters such as axial thrust, convective and radiative wall heat fluxes for regeneratively cooled liquid rocket engine nozzles, so as to develop a computational strategy for computing those parameters through parametric investigations. The computational methodology is based on a multidimensional, finite-volume, turbulent, chemically reacting, radiating, unstructured-grid, and pressure-based formulation, with grid refinement capabilities. Systematic parametric studies on effects of wall boundary conditions, combustion chemistry, radiation coupling, computational cell shape, and grid refinement were performed and assessed. Under the computational framework of this study, it is found that the computed axial thrust performance, flow features, and wall heat fluxes compared well with those of available data and calculations, using a strategy of structured-grid dominated mesh, finite-rate chemistry, and cooled wall boundary condition.

  5. Very Low Thrust Gaseous Oxygen-hydrogen Rocket Engine Ignition Technology

    Science.gov (United States)

    Bjorklund, Roy A.

    1983-01-01

    An experimental program was performed to determine the minimum energy per spark for reliable and repeatable ignition of gaseous oxygen (GO2) and gaseous hydrogen (GH2) in very low thrust 0.44 to 2.22-N (0.10 to 0.50-lb sub f) rocket engines or spacecraft and satellite attitude control systems (ACS) application. Initially, the testing was conducted at ambient conditions, with the results subsequently verified under vacuum conditions. An experimental breadboard electrical exciter that delivered 0.2 to 0.3 mj per spark was developed and demonstrated by repeated ignitions of a 2.22-N (0.50-lb sub f) thruster in a vacuum chamber with test durations up to 30 min.

  6. Development of a model for baffle energy dissipation in liquid fueled rocket engines

    Science.gov (United States)

    Miller, Nathan A.

    In this thesis the energy dissipation from a combined hub and blade baffle structure in a combustion chamber of a liquid-fueled rocket engine is modeled and computed. An analytical model of the flow stabilization due to the effect of combined radial and hub blades was developed. The rate of energy dissipation of the baffle blades was computed using a corner-flow model that included unsteady flow separation and turbulence effects. For the inviscid portion of the flow field, a solution methodology was formulated using an eigenfunction expansion and a velocity potential matching technique. Parameters such as local velocity, elemental path length, effective viscosity, and local energy dissipation rate were computed as a function of the local angle alpha for a representative baffle blade, and compared to results predicted by the Baer-Mitchell blade dissipation model. The sensitivity of the model to the overall engine acoustic oscillation mode, blade length, and thickness was also computed and compared to previous results. Additional studies were performed to determine the sensitivity to input parameters such as the dimensionless turbulence coefficient, the location of the potential difference in the generation of the dividing streamline, the number of baffle blades and the size of the central hub. Stability computations of a test engine indicated that when the baffle length is increased, the baffles provide increased stabilization effects. The model predicts greatest dissipation for radial modes with a hub radius at approximately half the chamber's radius.

  7. Fuzzy/Neural Software Estimates Costs of Rocket-Engine Tests

    Science.gov (United States)

    Douglas, Freddie; Bourgeois, Edit Kaminsky

    2005-01-01

    The Highly Accurate Cost Estimating Model (HACEM) is a software system for estimating the costs of testing rocket engines and components at Stennis Space Center. HACEM is built on a foundation of adaptive-network-based fuzzy inference systems (ANFIS) a hybrid software concept that combines the adaptive capabilities of neural networks with the ease of development and additional benefits of fuzzy-logic-based systems. In ANFIS, fuzzy inference systems are trained by use of neural networks. HACEM includes selectable subsystems that utilize various numbers and types of inputs, various numbers of fuzzy membership functions, and various input-preprocessing techniques. The inputs to HACEM are parameters of specific tests or series of tests. These parameters include test type (component or engine test), number and duration of tests, and thrust level(s) (in the case of engine tests). The ANFIS in HACEM are trained by use of sets of these parameters, along with costs of past tests. Thereafter, the user feeds HACEM a simple input text file that contains the parameters of a planned test or series of tests, the user selects the desired HACEM subsystem, and the subsystem processes the parameters into an estimate of cost(s).

  8. Thermal barrier coatings - Technology for diesel engines

    International Nuclear Information System (INIS)

    Harris, D.H.; Lutz, J.

    1988-01-01

    Thermal Barrier Coatings (TBC) are a development of the aerospace industry primarily aimed at hot gas flow paths in turbine engines. TBC consists of zirconia ceramic coatings applied over (M)CrAlY. These coatings can provide three benefits: (1) a reduction of metal surface operating temperatures, (2) a deterrent to hot gas corrosion, and (3) improved thermal efficiencies. TBC brings these same benefits to reciprocal diesel engines but coating longevity must be demonstrated. Diesels require thicker deposits and have challenging geometries for the arc-plasma spray (APS) deposition process. Different approaches to plasma spraying TBC are required for diesels, especially where peripheral edge effects play a major role. Bondcoats and ceramic top coats are modified to provide extended life as determined by burner rig tests, using ferrous and aluminum substrates

  9. Nuclear thermal propulsion engine cost trade studies

    International Nuclear Information System (INIS)

    Paschall, R.K.

    1993-01-01

    The NASA transportation strategy for the Mars Exploration architecture includes the use of nuclear thermal propulsion as the primary propulsion system for Mars transits. It is anticipated that the outgrowth of the NERVA/ROVER programs will be a nuclear thermal propulsion (NTP) system capable of providing the propulsion for missions to Mars. The specific impulse (Isp) for such a system is expected to be in the 870 s range. Trade studies were conducted to investigate whether or not it may be cost effective to invest in a higher performance (Isp>870 s) engine for nuclear thermal propulsion for missions to Mars. The basic cost trades revolved around the amount of mass that must be transported to low-earth orbit prior to each Mars flight and the cost to launch that mass. The mass required depended on the assumptions made for Mars missions scenarios including piloted/cargo flights, number of Mars missions, and transit time to Mars. Cost parameters included launch cost, program schedule for development and operations, and net discount rate. The results were very dependent on the assumptions that were made. Under some assumptions, higher performance engines showed cost savings in the billions of dollars; under other assumptions, the additional cost to develop higher performance engines was not justified

  10. Ground Testing a Nuclear Thermal Rocket: Design of a sub-scale demonstration experiment

    Energy Technology Data Exchange (ETDEWEB)

    David Bedsun; Debra Lee; Margaret Townsend; Clay A. Cooper; Jennifer Chapman; Ronald Samborsky; Mel Bulman; Daniel Brasuell; Stanley K. Borowski

    2012-07-01

    In 2008, the NASA Mars Architecture Team found that the Nuclear Thermal Rocket (NTR) was the preferred propulsion system out of all the combinations of chemical propulsion, solar electric, nuclear electric, aerobrake, and NTR studied. Recently, the National Research Council committee reviewing the NASA Technology Roadmaps recommended the NTR as one of the top 16 technologies that should be pursued by NASA. One of the main issues with developing a NTR for future missions is the ability to economically test the full system on the ground. In the late 1990s, the Sub-surface Active Filtering of Exhaust (SAFE) concept was first proposed by Howe as a method to test NTRs at full power and full duration. The concept relied on firing the NTR into one of the test holes at the Nevada Test Site which had been constructed to test nuclear weapons. In 2011, the cost of testing a NTR and the cost of performing a proof of concept experiment were evaluated.

  11. Fast reconstruction of an unmanned engineering vehicle and its application to carrying rocket

    Directory of Open Access Journals (Sweden)

    Jun Qian

    2014-04-01

    Full Text Available Engineering vehicle is widely used as a huge moving platform for transporting heavy goods. However, traditional human operations have a great influence on the steady movement of the vehicle. In this Letter, a fast reconstruction process of an unmanned engineering vehicle is carried out. By adding a higher-level controller and two two-dimensional laser scanners on the moving platform, the vehicle could perceive the surrounding environment and locate its pose according to extended Kalman filter. Then, a closed-loop control system is formed by communicating with the on-board lower-level controller. To verify the performance of automatic control system, the unmanned vehicle is automatically navigated when carrying a rocket towards a launcher in a launch site. The experimental results show that the vehicle could align with the launcher smoothly and safely within a small lateral deviation of 1 cm. This fast reconstruction presents an efficient way of rebuilding low-cost unmanned special vehicles and other automatic moving platforms.

  12. Derating design for optimizing reliability and cost with an application to liquid rocket engines

    International Nuclear Information System (INIS)

    Kim, Kyungmee O.; Roh, Taeseong; Lee, Jae-Woo; Zuo, Ming J.

    2016-01-01

    Derating is the operation of an item at a stress that is lower than its rated design value. Previous research has indicated that reliability can be increased from operational derating. In order to derate an item in field operation, however, an engineer must rate the design of the item at a stress level higher than the operational stress level, which increases the item's nominal failure rate and development costs. At present, there is no model available to quantify the cost and reliability that considers the design uprating as well as the operational derating. In this paper, we establish the reliability expression in terms of the derating level assuming that the nominal failure rate is constant with time for a fixed rated design value. The total development cost is expressed in terms of the rated design value and the number of tests necessary to demonstrate the reliability requirement. The properties of the optimal derating level are explained for maximizing the reliability or for minimizing the cost. As an example, the proposed model is applied to the design of liquid rocket engines. - Highlights: • Modeled the effect of derating design on the reliability and the development cost. • Discovered that derating design may reduce the cost of reliability demonstration test. • Optimized the derating design parameter for reliability maximization or cost minimization.

  13. Integrated System Health Management: Pilot Operational Implementation in a Rocket Engine Test Stand

    Science.gov (United States)

    Figueroa, Fernando; Schmalzel, John L.; Morris, Jonathan A.; Turowski, Mark P.; Franzl, Richard

    2010-01-01

    This paper describes a credible implementation of integrated system health management (ISHM) capability, as a pilot operational system. Important core elements that make possible fielding and evolution of ISHM capability have been validated in a rocket engine test stand, encompassing all phases of operation: stand-by, pre-test, test, and post-test. The core elements include an architecture (hardware/software) for ISHM, gateways for streaming real-time data from the data acquisition system into the ISHM system, automated configuration management employing transducer electronic data sheets (TEDS?s) adhering to the IEEE 1451.4 Standard for Smart Sensors and Actuators, broadcasting and capture of sensor measurements and health information adhering to the IEEE 1451.1 Standard for Smart Sensors and Actuators, user interfaces for management of redlines/bluelines, and establishment of a health assessment database system (HADS) and browser for extensive post-test analysis. The ISHM system was installed in the Test Control Room, where test operators were exposed to the capability. All functionalities of the pilot implementation were validated during testing and in post-test data streaming through the ISHM system. The implementation enabled significant improvements in awareness about the status of the test stand, and events and their causes/consequences. The architecture and software elements embody a systems engineering, knowledge-based approach; in conjunction with object-oriented environments. These qualities are permitting systematic augmentation of the capability and scaling to encompass other subsystems.

  14. Genetic algorithm to optimize the design of main combustor and gas generator in liquid rocket engines

    Science.gov (United States)

    Son, Min; Ko, Sangho; Koo, Jaye

    2014-06-01

    A genetic algorithm was used to develop optimal design methods for the regenerative cooled combustor and fuel-rich gas generator of a liquid rocket engine. For the combustor design, a chemical equilibrium analysis was applied, and the profile was calculated using Rao's method. One-dimensional heat transfer was assumed along the profile, and cooling channels were designed. For the gas-generator design, non-equilibrium properties were derived from a counterflow analysis, and a vaporization model for the fuel droplet was adopted to calculate residence time. Finally, a genetic algorithm was adopted to optimize the designs. The combustor and gas generator were optimally designed for 30-tonf, 75-tonf, and 150-tonf engines. The optimized combustors demonstrated superior design characteristics when compared with previous non-optimized results. Wall temperatures at the nozzle throat were optimized to satisfy the requirement of 800 K, and specific impulses were maximized. In addition, the target turbine power and a burned-gas temperature of 1000 K were obtained from the optimized gas-generator design.

  15. Thermal stability engineering of Glomerella cingulata cutinase.

    Science.gov (United States)

    Chin, Iuan-Sheau; Abdul Murad, Abdul Munir; Mahadi, Nor Muhammad; Nathan, Sheila; Abu Bakar, Farah Diba

    2013-05-01

    Cutinase has been ascertained as a biocatalyst for biotechnological and industrial bioprocesses. The Glomerella cingulata cutinase was genetically modified to enhance its enzymatic performance to fulfill industrial requirements. Two sites were selected for mutagenesis with the aim of altering the surface electrostatics as well as removing a potentially deamidation-prone asparagine residue. The N177D cutinase variant was affirmed to be more resilient to temperature increase with a 2.7-fold increase in half-life at 50°C as compared with wild-type enzyme, while, the activity at 25°C is not compromised. Furthermore, the increase in thermal tolerance of this variant is accompanied by an increase in optimal temperature. Another variant, the L172K, however, exhibited higher enzymatic performance towards phenyl ester substrates of longer carbon chain length, yet its thermal stability is inversely affected. In order to restore the thermal stability of L172K, we constructed a L172K/N177D double variant and showed that these two mutations yield an improved variant with enhanced activity towards phenyl ester substrates and enhanced thermal stability. Taken together, our study may provide valuable information for enhancing catalytic performance and thermal stability in future engineering endeavors.

  16. Continuous Space Education System and its Role in Increasing Efficiency of Engineering Staff Training for Ukraine Space Rocket Industry

    Science.gov (United States)

    Novykov, O.; Perlik, V.; Polyakov, N.; Khytorniy, V.

    2009-01-01

    Adjustment to new economical and social conditions in Ukraine, being a space-faring country, requires a new concept in improving efficiency of engineering education to provide rocket and space field with highly qualified engineers and scientists. General strategy to solve this task is to combine the efforts of the secondary schools, academic institutions, R&D institutes and production enterprises in order to educe gifted youth as early as possible and to train them into rocket and space field specialists following the scheme of continuous education: school - institution of higher education - enterprise. This report analyzes the 20-year experience of Dniepropetrovsk State University, Yuzhnoye State Design Office and Ukraine's National Center for the Aerospace Education of Youth in their joint efforts to organize the system of continuous education and how it managed to enhance the training efficiency of the engineering skills.

  17. Improving the performance of LOX/kerosene upper stage rocket engines

    Directory of Open Access Journals (Sweden)

    IgorN. Nikischenko

    2017-09-01

    Full Text Available Improved liquid rocket engine cycles were proposed and analyzed via comparison with existing staged combustion and gas-generator cycles. The key features of the proposed cycles are regenerative cooling of thrust chamber by oxygen and subsequent use of this oxygen for driving one or two oxygen pumps. The fuel pump(s are driven in a conventional manner, for example, using a fuel-rich gas-generator cycle. Comparison with staged combustion cycle based on oxygen-rich pre-burner showed that one of the proposed semi-expander cycles has a specific impulse only on 0.4% lower while providing much lower oxygen temperature, more efficient tank pressurizing system and built-in roll control. This semi-expander cycle can be considered as a more reliable and cost-effective alternative of staged combustion cycle. Another semi-expander cycle can be considered as an improvement of gas-generator cycle. All proposed semi-expander cycles were developed as a derivative of thrust chamber regenerative cooling performed by oxygen. Analysis of existing oxygen/kerosene engines showed that replacing of kerosene regenerative cooling with oxygen allows a significant increase of achievable specific impulse, via optimization of mixture ratio. It is especially the case for upper stage engines. The increasing of propellants average density can be considered as an additional benefit of mixture ratio optimization. It was demonstrated that oxygen regenerative cooling of thrust chamber is a feasible and the most promising option for oxygen/kerosene engines. Combination of oxygen regenerative cooling and semi-expander cycles potentially allows creating the oxygen/kerosene propulsion systems with minimum specific impulse losses. It is important that such propulsion systems can be fully based on inherited and well-proven technical solutions. A hypothetic upper stage engine with thrust 19.6 kN was chosen as a prospective candidate for theoretical analysis of the proposed semi

  18. Neural Network and Response Surface Methodology for Rocket Engine Component Optimization

    Science.gov (United States)

    Vaidyanathan, Rajkumar; Papita, Nilay; Shyy, Wei; Tucker, P. Kevin; Griffin, Lisa W.; Haftka, Raphael; Fitz-Coy, Norman; McConnaughey, Helen (Technical Monitor)

    2000-01-01

    The goal of this work is to compare the performance of response surface methodology (RSM) and two types of neural networks (NN) to aid preliminary design of two rocket engine components. A data set of 45 training points and 20 test points obtained from a semi-empirical model based on three design variables is used for a shear coaxial injector element. Data for supersonic turbine design is based on six design variables, 76 training, data and 18 test data obtained from simplified aerodynamic analysis. Several RS and NN are first constructed using the training data. The test data are then employed to select the best RS or NN. Quadratic and cubic response surfaces. radial basis neural network (RBNN) and back-propagation neural network (BPNN) are compared. Two-layered RBNN are generated using two different training algorithms, namely solverbe and solverb. A two layered BPNN is generated with Tan-Sigmoid transfer function. Various issues related to the training of the neural networks are addressed including number of neurons, error goals, spread constants and the accuracy of different models in representing the design space. A search for the optimum design is carried out using a standard gradient-based optimization algorithm over the response surfaces represented by the polynomials and trained neural networks. Usually a cubic polynominal performs better than the quadratic polynomial but exceptions have been noticed. Among the NN choices, the RBNN designed using solverb yields more consistent performance for both engine components considered. The training of RBNN is easier as it requires linear regression. This coupled with the consistency in performance promise the possibility of it being used as an optimization strategy for engineering design problems.

  19. Thermal barrier coatings for heat engine components

    Science.gov (United States)

    Levine, S. R.; Miller, R. A.; Hodge, P. E.

    1980-01-01

    A comprehensive NASA-Lewis program of coating development for aircraft gas turbine blades and vanes is presented. Improved ceramic layer compositions are investigated, along the MCrAlY bond films and the methods of uniform deposition of the coatings; the thermomechanical and fuel impurity tolerance limits of the coatings are being studied. Materials include the ZrO2-Y2O3/NiCrAlY system; the effects of the bond coat and zirconia composition on coating life and Mach 1 burner rig test results are discussed. It is concluded that Diesel engines can also utilize thermal barrier coatings; they have been used successfully on piston crowns and exhaust valves of shipboard engines to combat lower grade fuel combustion corrosion.

  20. Concept Assessment of a Fission Fragment Rocket Engine (FFRE) Propelled Spacecraft

    Science.gov (United States)

    Werka, Robert; Clark, Rod; Sheldon, Rob; Percy, Tom

    2012-01-01

    The March, 2012 issue of Aerospace America stated that ?the near-to-medium prospects for applying advanced propulsion to create a new era of space exploration are not very good. In the current world, we operate to the Moon by climbing aboard a Carnival Cruise Lines vessel (Saturn 5), sail from the harbor (liftoff) shedding whole decks of the ship (staging) along the way and, having reached the return leg of the journey, sink the ship (burnout) and return home in a lifeboat (Apollo capsule). Clearly this is an illogical way to travel, but forced on Explorers by today's propulsion technology. However, the article neglected to consider the one propulsion technology, using today's physical principles that offer continuous, substantial thrust at a theoretical specific impulse of 1,000,000 sec. This engine unequivocally can create a new era of space exploration that changes the way spacecraft operate. Today's space Explorers could travel in Cruise Liner fashion using the technology not considered by Aerospace America, the novel Dusty Plasma Fission Fragment Rocket Engine (FFRE). This NIAC study addresses the FFRE as well as its impact on Exploration Spacecraft design and operation. It uses common physics of the relativistic speed of fission fragments to produce thrust. It radiatively cools the fissioning dusty core and magnetically controls the fragments direction to practically implement previously patented, but unworkable designs. The spacecraft hosting this engine is no more complex nor more massive than the International Space Station (ISS) and would employ the successful ISS technology for assembly and check-out. The elements can be lifted in "chunks" by a Heavy Lift Launcher. This Exploration Spacecraft would require the resupply of small amounts of nuclear fuel for each journey and would be an in-space asset for decades just as any Cruise Liner on Earth. This study has synthesized versions of the FFRE, integrated one concept onto a host spacecraft designed for

  1. Modeling the Thermal Rocket Fuel Preparation Processes in the Launch Complex Fueling System

    Directory of Open Access Journals (Sweden)

    A. V. Zolin

    2015-01-01

    Full Text Available It is necessary to carry out fuel temperature preparation for space launch vehicles using hydrocarbon propellant components. A required temperature is reached with cooling or heating hydrocarbon fuel in ground facilities fuel storages. Fuel temperature preparing processes are among the most energy-intensive and lengthy processes that require the optimal technologies and regimes of cooling (heating fuel, which can be defined using the simulation of heat exchange processes for preparing the rocket fuel.The issues of research of different technologies and simulation of cooling processes of rocket fuel with liquid nitrogen are given in [1-10]. Diagrams of temperature preparation of hydrocarbon fuel, mathematical models and characteristics of cooling fuel with its direct contact with liquid nitrogen dispersed are considered, using the numerical solution of a system of heat transfer equations, in publications [3,9].Analytical models, allowing to determine the necessary flow rate and the mass of liquid nitrogen and the cooling (heating time fuel in specific conditions and requirements, are preferred for determining design and operational characteristics of the hydrocarbon fuel cooling system.A mathematical model of the temperature preparation processes is developed. Considered characteristics of these processes are based on the analytical solutions of the equations of heat transfer and allow to define operating parameters of temperature preparation of hydrocarbon fuel in the design and operation of the filling system of launch vehicles.The paper considers a technological system to fill the launch vehicles providing the temperature preparation of hydrocarbon gases at the launch site. In this system cooling the fuel in the storage tank before filling the launch vehicle is provided by hydrocarbon fuel bubbling with liquid nitrogen. Hydrocarbon fuel is heated with a pumping station, which provides fuel circulation through the heat exchanger-heater, with

  2. Experimental investigation of high-frequency combustion instabilities in liquid rocket engine

    Science.gov (United States)

    Richecoeur, F.; Ducruix, S.; Scouflaire, P.; Candel, S.

    2008-01-01

    High-frequency instabilities in liquid propellant rocket engines are experimentally investigated in a model scale research facility. Liquid oxygen and gaseous methane are injected in the combustion chamber at 0.9 MPa through three coaxial injectors vertically aligned. High-amplitude transverse pressure fluctuations are generated in the chamber at frequencies above 1 kHz by a rotating toothed wheel actuator which periodically blocks an auxiliary lateral nozzle. The chamber eigenmodes are identified in a first stage by examining the response of the system to a linear frequency sweep. In a second stage the chamber is excited at the frequency corresponding to the first transverse (1T) mode. The effect of the pressure mode on combustion is observed with intensified and high-speed cameras. Photo-multipliers and pressure sensors are also used to characterize the system behavior and examine phase relations between the corresponding signals. Flame structure modifications observed for specific injection conditions correspond to a strong coupling between acoustics and combustion which notably modifies the flow dynamics, augments the flame expansion rate and enhances heat transfer to the wall.

  3. Simulation of supercritical flows in rocket-motor engines: application to cooling channel and injection system

    Science.gov (United States)

    Ribert, G.; Taieb, D.; Petit, X.; Lartigue, G.; Domingo, P.

    2013-03-01

    To address physical modeling of supercritical multicomponent fluid flows, ideal-gas law must be changed to real-gas equation of state (EoS), thermodynamic and transport properties have to incorporate dense fluid corrections, and turbulence modeling has to be reconsidered compared to classical approaches. Real-gas thermodynamic is presently investigated with validation by NIST (National Institute of Standards and Technology) data. Two major issues of Liquid Rocket Engines (LRE) are also presented. The first one is the supercritical fluid flow inside small cooling channels. In a context of LRE, a strong heat flux coming from the combustion chamber (locally Φ ≈ 80 MW/m2) may lead to very steep density gradients close to the wall. These gradients have to be thermodynamically and numerically captured to properly reproduce in the simulation the mechanism of heat transfer from the wall to the fluid. This is done with a shock-capturing weighted essentially nonoscillatory (WENO) numerical discretization scheme. The second issue is a supercritical fluid injection following experimental conditions [1] in which a trans- or supercritical nitrogen is injected into warm nitrogen. The two-dimensional results show vortex structures with high fluid density detaching from the main jet and persisting in the low-speed region with low fluid density.

  4. Highly resolved numerical simulation of combustion downstream of a rocket engine igniter

    Science.gov (United States)

    Buttay, R.; Gomet, L.; Lehnasch, G.; Mura, A.

    2017-07-01

    We study ignition processes in the turbulent reactive flow established downstream of highly under-expanded coflowing jets. The corresponding configuration is typical of a rocket engine igniter, and to the best knowledge of the authors, this study is the first that documents highly resolved numerical simulations of such a reactive flowfield. Considering the discharge of axisymmetric coaxial under-expanded jets, various morphologies are expected, depending on the value of the nozzle pressure ratio, a key parameter used to classify them. The present computations are conducted with a value of this ratio set to fifteen. The simulations are performed with the massively parallel CREAMS solver on a grid featuring approximately 440,000,000 computational nodes. In the main zone of interest, the level of spatial resolution is D/74, with D the central inlet stream diameter. The computational results reveal the complex topology of the compressible flowfield. The obtained results also bring new and useful insights into the development of ignition processes. In particular, ignition is found to take place rather far downstream of the shock barrel, a conclusion that contrasts with early computational studies conducted within the unsteady RANS computational framework. Consideration of detailed chemistry confirms the essential role of hydroperoxyl radicals, while the analysis of the Takeno index reveals the predominance of a non-premixed combustion mode.

  5. Method and apparatus to produce high specific impulse and moderate thrust from a fusion-powered rocket engine

    Energy Technology Data Exchange (ETDEWEB)

    Cohen, Samuel A.; Pajer, Gary A.; Paluszek, Michael A.; Razin, Yosef S.

    2017-11-21

    A system and method for producing and controlling high thrust and desirable specific impulse from a continuous fusion reaction is disclosed. The resultant relatively small rocket engine will have lower cost to develop, test, and operate that the prior art, allowing spacecraft missions throughout the planetary system and beyond. The rocket engine method and system includes a reactor chamber and a heating system for heating a stable plasma to produce fusion reactions in the stable plasma. Magnets produce a magnetic field that confines the stable plasma. A fuel injection system and a propellant injection system are included. The propellant injection system injects cold propellant into a gas box at one end of the reactor chamber, where the propellant is ionized into a plasma. The propellant and fusion products are directed out of the reactor chamber through a magnetic nozzle and are detached from the magnetic field lines producing thrust.

  6. An Approximate Analysis of the Inner Wall Loading of a Bimetallic Camera Shell of Reusable Rocket Engine

    Directory of Open Access Journals (Sweden)

    V. S. Zarubin

    2016-01-01

    Full Text Available Various technical devices quite widely use bimetallic shells as the structural elements. A chamber combustion design of the liquid rocket engine (LRE is a typical use of the bimetallic shells.In LRE operation a combustion chamber shell is subject to intense thermal and mechanical effects, which necessitates cooling. A cooling shell path is formed by a gap between its inner and outer walls connected to each other by milled or grooved spacer ribs. The outer wall of the shell serves as a load-bearing element, the inner wall is in direct contact with high-temperature combustion products and exposed to intense heat. The difference in functions of shell walls calls for their manufacturing from different materials with different thermophysical and mechanical properties.Interaction between the shell walls of different materials in heating and cooling leads to emerging thermal strains of various values in the walls. In terms of mechanical properties the inner wall material, usually ranks below the outer wall material strength, which uses the high strength stainless steel 12Х21Н5Т. The inner wall is typically made from copper-based highly heat-conductive alloys. (eg.: chromium bronze. Therefore, the result of the difference in temperature deformations, arising in the walls,  is inelastic nonisothermal strain of the inner wall material with (usually elastic behavior of the outer wall material.For reusable LRE, a cyclic sequence of the loading steps of the inner wall can lead to accumulating damages in its material because of the low-cycle fatigue and cause destruction of the wall or the loss of the cooling tract tightness. The main parameter that determines the level of low-cycle fatigue, is an absolute value of the accumulated inelastic strain (both plastic and evolving over time creep deformation. Quantitative evaluation of this parameter involves analysis of the inner wall loading with multiple starts and shutdowns of LRE. The paper represents an

  7. 2D and 3D Modeling Efforts in Fuel Film Cooling of Liquid Rocket Engines (Conference Paper with Briefing Charts)

    Science.gov (United States)

    2017-01-12

    Conference Paper with Briefing Charts 3. DATES COVERED (From - To) 17 November 2016 – 12 January 2017 4. TITLE AND SUBTITLE 2D and 3D Modeling ...98) Prescribed by ANSI Std. 239.18 2D and 3D Modeling Efforts in Fuel Film Cooling of Liquid Rocket Engines Kevin C. Brown∗, Edward B. Coy†, and...wide. As a consequence, the 3D simulations may better model the experimental setup used, but are perhaps not representative of the long circumferential

  8. Low-Cost High-Performance Non-Toxic Self-Pressurizing Storable Liquid Bi-Propellant Pressure-Fed Rocket Engine, Phase I

    Data.gov (United States)

    National Aeronautics and Space Administration — Exquadrum proposes a high-performance liquid bi-propellant rocket engine that uses propellants that are non-toxic, self-pressurizing, and low cost. The proposed...

  9. Vibration, acoustic, and shock design and test criteria for components on the Solid Rocket Boosters (SRB), Lightweight External Tank (LWT), and Space Shuttle Main Engines (SSME)

    Science.gov (United States)

    1984-01-01

    The vibration, acoustics, and shock design and test criteria for components and subassemblies on the space shuttle solid rocket booster (SRB), lightweight tank (LWT), and main engines (SSME) are presented. Specifications for transportation, handling, and acceptance testing are also provided.

  10. A thermal engine for underwater glider driven by ocean thermal energy

    International Nuclear Information System (INIS)

    Yang, Yanan; Wang, Yanhui; Ma, Zhesong; Wang, Shuxin

    2016-01-01

    Highlights: • Thermal engine with a double-tube structure is developed for underwater glider. • Isostatic pressing technology is effective to increase volumetric change rate. • Actual volumetric change rate reaches 89.2% of the theoretical value. • Long term sailing of 677 km and 27 days is achieved by thermal underwater glider. - Graphical Abstract: - Abstract: Underwater glider is one of the most popular platforms for long term ocean observation. Underwater glider driven by ocean thermal energy extends the duration and range of underwater glider powered by battery. Thermal engine is the core device of underwater glider to harvest ocean thermal energy. In this paper, (1) model of thermal engine was raised by thermodynamics method and the performance of thermal engine was investigated, (2) thermal engine with a double-tube structure was developed and isostatic pressing technology was applied to improve the performance for buoyancy driven, referencing powder pressing theory, (3) wall thickness of thermal engine was optimized to reduce the overall weight of thermal engine, (4) material selection and dimension determination were discussed for a faster heat transfer design, by thermal resistance analysis, (5) laboratory test and long term sea trail were carried out to test the performance of thermal engine. The study shows that volumetric change rate is the most important indicator to evaluating buoyancy-driven performance of a thermal engine, isostatic pressing technology is effective to improve volumetric change rate, actual volumetric change rate can reach 89.2% of the theoretical value and the average power is about 124 W in a typical diving profile. Thermal engine developed by Tianjin University is a superior thermal energy conversion device for underwater glider. Additionally, application of thermal engine provides a new solution for miniaturization of ocean thermal energy conversion.

  11. Reuse fo a Cold War Surveillance Drone to Flight Test a NASA Rocket Based Combined Cycle Engine

    Science.gov (United States)

    Brown, T. M.; Smith, Norm

    1999-01-01

    Plans for and early feasibility investigations into the modification of a Lockheed D21B drone to flight test the DRACO Rocket Based Combined Cycle (RBCC) engine are discussed. Modifications include the addition of oxidizer tanks, modern avionics systems, actuators, and a vehicle recovery system. Current study results indicate that the D21B is a suitable candidate for this application and will allow demonstrations of all DRACO engine operating modes at Mach numbers between 0.8 and 4.0. Higher Mach numbers may be achieved with more extensive modification. Possible project risks include low speed stability and control, and recovery techniques.

  12. Techniques to assess acoustic-structure interaction in liquid rocket engines

    Science.gov (United States)

    Davis, R. Benjamin

    Acoustoelasticity is the study of the dynamic interaction between elastic structures and acoustic enclosures. In this dissertation, acoustoelasticity is considered in the context of liquid rocket engine design. The techniques presented here can be used to determine which forcing frequencies are important in acoustoelastic systems. With a knowledge of these frequencies, an analyst can either find ways to attenuate the excitation at these frequencies or alter the system in such a way that the prescribed excitations do result in a resonant condition. The end result is a structural component that is less susceptible to failure. The research scope is divided into three parts. In the first part, the dynamics of cylindrical shells submerged in liquid hydrogen (LH2) and liquid oxygen (LOX) are considered. The shells are bounded by rigid outer cylinders. This configuration gives rise to two fluid-filled cavities---an inner cylindrical cavity and an outer annular cavity. Such geometries are common in rocket engine design. The natural frequencies and modes of the fluid-structure system are computed by combining the rigid wall acoustic cavity modes and the in vacuo structural modes into a system of coupled ordinary differential equations. Eigenvalue veering is observed near the intersections of the curves representing natural frequencies of the rigid wall acoustic and the in vacuo structural modes. In the case of a shell submerged in LH2, system frequencies near these intersections are as much as 30% lower than the corresponding in vacuo structural frequencies. Due to its high density, the frequency reductions in the presence of LOX are even more dramatic. The forced responses of a shell submerged in LH2 and LOX while subject to a harmonic point excitation are also presented. The responses in the presence of fluid are found to be quite distinct from those of the structure in vacuo. In the second part, coupled mode theory is used to explore the fundamental features of

  13. Using Decision Trees to Detect and Isolate Simulated Leaks in the J-2X Rocket Engine

    Science.gov (United States)

    Schwabacher, Mark A.; Aguilar, Robert; Figueroa, Fernando F.

    2009-01-01

    The goal of this work was to use data-driven methods to automatically detect and isolate faults in the J-2X rocket engine. It was decided to use decision trees, since they tend to be easier to interpret than other data-driven methods. The decision tree algorithm automatically "learns" a decision tree by performing a search through the space of possible decision trees to find one that fits the training data. The particular decision tree algorithm used is known as C4.5. Simulated J-2X data from a high-fidelity simulator developed at Pratt & Whitney Rocketdyne and known as the Detailed Real-Time Model (DRTM) was used to "train" and test the decision tree. Fifty-six DRTM simulations were performed for this purpose, with different leak sizes, different leak locations, and different times of leak onset. To make the simulations as realistic as possible, they included simulated sensor noise, and included a gradual degradation in both fuel and oxidizer turbine efficiency. A decision tree was trained using 11 of these simulations, and tested using the remaining 45 simulations. In the training phase, the C4.5 algorithm was provided with labeled examples of data from nominal operation and data including leaks in each leak location. From the data, it "learned" a decision tree that can classify unseen data as having no leak or having a leak in one of the five leak locations. In the test phase, the decision tree produced very low false alarm rates and low missed detection rates on the unseen data. It had very good fault isolation rates for three of the five simulated leak locations, but it tended to confuse the remaining two locations, perhaps because a large leak at one of these two locations can look very similar to a small leak at the other location.

  14. Development and Hotfire Testing of Additively Manufactured Copper Combustion Chambers for Liquid Rocket Engine Applications

    Science.gov (United States)

    Gradl, Paul R.; Greene, Sandy; Protz, Chris

    2017-01-01

    NASA and industry partners are working towards fabrication process development to reduce costs and schedules associated with manufacturing liquid rocket engine components with the goal of reducing overall mission costs. One such technique being evaluated is powder-bed fusion or selective laser melting (SLM), commonly referred to as additive manufacturing (AM). The NASA Low Cost Upper Stage Propulsion (LCUSP) program was designed to develop processes and material characterization for GRCop-84 (a NASA Glenn Research Center-developed copper, chrome, niobium alloy) commensurate with powder bed AM, evaluate bimetallic deposition, and complete testing of a full scale combustion chamber. As part of this development, the process has been transferred to industry partners to enable a long-term supply chain of monolithic copper combustion chambers. To advance the processes further and allow for optimization with multiple materials, NASA is also investigating the feasibility of bimetallic AM chambers. In addition to the LCUSP program, NASA’s Marshall Space Flight Center (MSFC) has completed a series of development programs and hot-fire tests to demonstrate SLM GRCop-84 and other AM techniques. MSFC’s efforts include a 4,000 pounds-force thrust liquid oxygen/methane (LOX/CH4) combustion chamber. Small thrust chambers for 1,200 pounds-force LOX/hydrogen (H2) applications have also been designed and fabricated with SLM GRCop-84. Similar chambers have also completed development with an Inconel 625 jacket bonded to the GRCop-84 material, evaluating direct metal deposition (DMD) laser- and arc-based techniques. The same technologies for these lower thrust applications are being applied to 25,000-35,000 pounds-force main combustion chamber (MCC) designs. This paper describes the design, development, manufacturing and testing of these numerous combustion chambers, and the associated lessons learned throughout their design and development processes.

  15. Introduction to thermal and fluid engineering

    CERN Document Server

    Kraus, Allan D; Aziz, Abdul; Ghajar, Afshin J

    2011-01-01

    The Thermal/Fluid Sciences: Introductory ConceptsThermodynamicsFluid MechanicsHeat TransferEngineered Systems and ProductsHistorical DevelopmentThe Thermal/Fluid Sciences and the EnvironmentThermodynamics: Preliminary Concepts and DefinitionsThe Study of ThermodynamicsSome DefinitionsDimensions and UnitsDensity and Related PropertiesPressureTemperature and the Zeroth Law of ThermodynamicsProblem-Solving MethodologyEnergy and the First Law of ThermodynamicsKinetic, Potential, and Internal EnergyWorkHeatThe First Law of ThermodynamicsThe Energy Balance for Closed SystemsThe Ideal Gas ModelIdeal Gas Enthalpy and Specific HeatsProcesses of an Ideal GasProperties of Pure, Simple Compressible SubstancesThe State PostulateP-v-T RelationshipsThermodynamic Property DataThe T-s and h-s DiagramsReal Gas BehaviorEquations of StateThe Polytropic Process for an Ideal GasControl Volume Mass and Energy Analysis The Control VolumeConservation of MassConservation of Energy for a Control VolumeSpecific Heats of Incompressible S...

  16. Development and Testing of Carbon-Carbon Nozzle Extensions for Upper Stage Liquid Rocket Engines

    Science.gov (United States)

    Valentine, Peter G.; Gradl, Paul R.; Greene, Sandra E.

    2017-01-01

    Carbon-carbon (C-C) composite nozzle extensions are of interest for use on a variety of launch vehicle upper stage engines and in-space propulsion systems. The C-C nozzle extension technology and test capabilities being developed are intended to support National Aeronautics and Space Administration (NASA) and Department of Defense (DOD) requirements, as well as those of the broader Commercial Space industry. For NASA, C-C nozzle extension technology development primarily supports the NASA Space Launch System (SLS) and NASA's Commercial Space partners. Marshall Space Flight Center (MSFC) efforts are aimed at both (a) further developing the technology and databases needed to enable the use of composite nozzle extensions on cryogenic upper stage engines, and (b) developing and demonstrating low-cost capabilities for testing and qualifying composite nozzle extensions. Recent, on-going, and potential future work supporting NASA, DOD, and Commercial Space needs will be discussed. Information to be presented will include (a) recent and on-going mechanical, thermal, and hot-fire testing, as well as (b) potential future efforts to further develop and qualify domestic C-C nozzle extension solutions for the various upper stage engines under development.

  17. Improving of technical characteristics of launch vehicles with liquid rocket engines using active onboard de-orbiting systems

    Science.gov (United States)

    Trushlyakov, V.; Shatrov, Ya.

    2017-09-01

    In this paper, the analysis of technical requirements (TR) for the development of modern space launch vehicles (LV) with main liquid rocket engines (LRE) is fulfilled in relation to the anthropogenic impact decreasing. Factual technical characteristics on the example of a promising type of rocket ;Soyuz-2.1.v.; are analyzed. Meeting the TR in relation to anthropogenic impact decrease based on the conventional design approach and the content of the onboard system does not prove to be efficient and leads to depreciation of the initial technical characteristics obtained at the first design stage if these requirements are not included. In this concern, it is shown that the implementation of additional active onboard de-orbiting system (AODS) of worked-off stages (WS) into the onboard LV stages systems allows to meet the TR related to the LV environmental characteristics, including fire-explosion safety. In some cases, the orbital payload mass increases.

  18. Thermodynamic analysis of thermal efficiency and power of Minto engine

    International Nuclear Information System (INIS)

    He, Wei; Hou, Jingxin; Zhang, Yang; Ji, Jie

    2011-01-01

    Minto engine is a kind of liquid piston heat engine that operates on a small temperature gradient. But there is no power formula for it yet. And its thermal efficiency is low and formula sometimes is misused. In this paper, deriving the power formula and simplifying the thermal efficiency formula of Minto engine based on energy distribution analysis will be discussed. To improve the original Minto engine, a new design of improved Minto engine is proposed and thermal efficiency formula and power formula are also given. A computer program was developed to analyze thermal efficiency and power of original and improved Minto engines operating between low and high-temperature heat sources. The simulation results show that thermal efficiency of improved Minto engine can reach over 7% between 293.15 K and 353.15 K which is much higher than that of original one; the temperature difference between upper and lower containers is lower than half of that between low and high temperature of heat sources when the original Minto engines output the maximum power; on the contrary, it is higher in the improved Minto engines. -- Highlights: ► The thermal efficiency formula of Minto engine is simplified and the power formula is established. ► A high-powered design of improved Minto engine is proposed. ► A computer simulation program based on real operating environment is developed.

  19. Buffer thermal energy storage for an air Brayton solar engine

    Science.gov (United States)

    Strumpf, H. J.; Barr, K. P.

    1981-01-01

    The application of latent-heat buffer thermal energy storage to a point-focusing solar receiver equipped with an air Brayton engine was studied. To demonstrate the effect of buffer thermal energy storage on engine operation, a computer program was written which models the recuperator, receiver, and thermal storage device as finite-element thermal masses. Actual operating or predicted performance data are used for all components, including the rotating equipment. Based on insolation input and a specified control scheme, the program predicts the Brayton engine operation, including flows, temperatures, and pressures for the various components, along with the engine output power. An economic parametric study indicates that the economic viability of buffer thermal energy storage is largely a function of the achievable engine life.

  20. The Pressure Field Measurement for Researching Inducer Flow of Booster Rocket Engine Turbopump

    Directory of Open Access Journals (Sweden)

    N. S. Dorosh

    2014-01-01

    Full Text Available When designing a feed system for modern main rocket engine development, designers have to pay special attention to energy efficiency of units and their reliability. One of the most important conditions of reliability is to provide non-cavitation operation of the main turbo-pump, which is impossible without using the booster turbo-pumps, considering the current levels of pressure in the combustion chamber. Thanks to high suction properties and processability, axial inducers with screw geometry became the most widely used in booster turbo-pumps. At the same time, the flow in the inducers of progressive geometry has complex spatial nature that makes their designing and detailed flow studying to be a difficult task.Based on the need of detailed understanding the flow structure in inducer channels a number of investigation methods are considered, including: analytical calculation, visual research methods, direct flow measurement, and numerical simulation. Analysis of the characteristics of each method shows the need to combine several methods to achieve the best results. Using a numerical simulation becomes the most effective strategy to obtain a wide range of data and confirm their authenticity by experimental measurements at characteristic points. The features of such kind of measurements in the inducer flow and measuring device requirements are considered.Based on this, an original design experimental booster turbo-pump, equipped with a pressure measuring system behind the inducer and automatic unloader device simulator is developed. Using these systems a radial pressure diagram of inducer flow as well as axial the force acting on the inducer can be experimentally obtained. It is shown that the offered measuring system satisfies those requirements and provides data at the various operation modes of the booster turbopump unit. A developed test program allows us to obtain required data: the pressure values in the flow behind inducer and axial force

  1. The engineering of a nuclear thermal landing and ascent vehicle utilizing indigenous Martian propellant

    International Nuclear Information System (INIS)

    Zubrin, R.M.

    1991-01-01

    The following paper reports on a design study of a novel space transportation concept known as a ''NIMF'' (Nuclear rocket using Indigenous Martian Fuel.) The NIMF is a ballistic vehicle which obtains its propellant out of the Martian air by compression and liquefaction of atmospheric CO 2 . This propellant is subsequently used to generate rocket thrust at a specific impulse of 264 s by being heated to high temperature (2800 K) gas in the NIMFs' nuclear thermal rocket engines. The vehicle is designed to provide surface to orbit and surface to surface transportation, as well as housing, for a crew of three astronauts. It is capable of refueling itself for a flight to its maximum orbit in less than 50 days. The ballistic NIMF has a mass of 44.7 tonnes and, with the assumed 2800 K propellant temperature, is capable of attaining highly energetic (250 km by 34000 km elliptical) orbits. This allows it to rendezvous with interplanetary transfer vehicles which are only very loosely bound into orbit around Mars. If a propellant temperature of 2000 K is assumed, then low Mars orbit can be attained; while if 3100 K is assumed, then the ballistic NIMF is capable of injecting itself onto a minimum energy transfer orbit to Earth in a direct ascent from the Martian surface

  2. Rockets: Educator's Guide with Activities in Science, Technology, Engineering and Mathematics

    Science.gov (United States)

    Shearer, Deborah A.; Vogt, Gregory L.

    2008-01-01

    This guide provides teachers and students many opportunities. Chapters within the guide present the history of rocketry, National Aeronautics and Space Administration's (NASA's) 21st Century Space Exploration Policy, rocketry principles, and practical rocketry. These topics lay the foundation for what follows--a wealth of dynamic rocket science…

  3. Thermal Expansion Behavior of Hot-Pressed Engineered Matrices

    Science.gov (United States)

    Raj, S. V.

    2016-01-01

    Advanced engineered matrix composites (EMCs) require that the coefficient of thermal expansion (CTE) of the engineered matrix (EM) matches those of the fiber reinforcements as closely as possible in order to reduce thermal compatibility strains during heating and cooling of the composites. The present paper proposes a general concept for designing suitable matrices for long fiber reinforced composites using a rule of mixtures (ROM) approach to minimize the global differences in the thermal expansion mismatches between the fibers and the engineered matrix. Proof-of-concept studies were conducted to demonstrate the validity of the concept.

  4. Nuclear thermal propulsion engine based on particle bed reactor using light water steam as a propellant

    Science.gov (United States)

    Powell, James R.; Ludewig, Hans; Maise, George

    1993-01-01

    In this paper the possibility of configuring a water cooled Nuclear Thermal Propulsion (NTP) rocket, based on a Particle Bed Reactor (PBR) is investigated. This rocket will be used to operate on water obtained from near earth objects. The conclusions reached in this paper indicate that it is possible to configure a PBR based NTP rocket to operate on water and meet the mission requirements envisioned for it. No insurmountable technology issues have been identified.

  5. THE POSSIBILITY OF USING LASER-ULTRASOUND TO MONITOR THE QUALITY SOLDERED CONNECTIONS CHAMBERS OF LIQUID ROCKET ENGINES

    Directory of Open Access Journals (Sweden)

    N. V. Astredinova

    2014-01-01

    Full Text Available During the manufacturing process to the design of modern liquid rocket engines are presented important requirements, such as minimum weight, maximum stiffness and strength of nodes, maximum service life in operation, high reliability and quality of soldered and welded seams. Due to the high quality requirements soldered connections and the specific design of the nozzle, it became necessary in the development and testing of a new non-conventional non-destructive testing method – laser-ultrasound diagnosis. In accordance with regulatory guidelines, quality control soldered connections is allowed to use an acoustic kind of control methods of the reflected light, transmitted light, resonant, free vibration and acoustic emission. Attempts to use traditional methods of non-destructive testing did not lead to positive results. This is due primarily to the size of typical solder joint defects, as well as the structural features of the rocket engine, the data structure is not controllable. In connection with this, a new method that provides quality control soldered connections cameras LRE based on the thermo generation of ultrasound. Methods of ultrasonic flaw detection of photoacoustic effect, in most cases, have a number of advantages over methods that use standard (traditional piezo transducers. In the course of studies have found that the sensitivity of the laser-ultrasonic method and flaw detector UDL-2M can detect lack of adhesion in the solder joints on the upper edges of the nozzle in the sub-header area of the site.

  6. Engine cycle design considerations for nuclear thermal propulsion systems

    International Nuclear Information System (INIS)

    Pelaccio, D.G.; Scheil, C.M.; Collins, J.T.

    1993-01-01

    A top-level study was performed which addresses nuclear thermal propulsion system engine cycle options and their applicability to support future Space Exploration Initiative manned lunar and Mars missions. Technical and development issues associated with expander, gas generator, and bleed cycle near-term, solid core nuclear thermal propulsion engines are identified and examined. In addition to performance and weight the influence of the engine cycle type on key design selection parameters such as design complexity, reliability, development time, and cost are discussed. Representative engine designs are presented and compared. Their applicability and performance impact on typical near-term lunar and Mars missions are shown

  7. The Primary Experiments of an Analysis of Pareto Solutions for Conceptual Design Optimization Problem of Hybrid Rocket Engine

    Science.gov (United States)

    Kudo, Fumiya; Yoshikawa, Tomohiro; Furuhashi, Takeshi

    Recentry, Multi-objective Genetic Algorithm, which is the application of Genetic Algorithm to Multi-objective Optimization Problems is focused on in the engineering design field. In this field, the analysis of design variables in the acquired Pareto solutions, which gives the designers useful knowledge in the applied problem, is important as well as the acquisition of advanced solutions. This paper proposes a new visualization method using Isomap which visualizes the geometric distances of solutions in the design variable space considering their distances in the objective space. The proposed method enables a user to analyze the design variables of the acquired solutions considering their relationship in the objective space. This paper applies the proposed method to the conceptual design optimization problem of hybrid rocket engine and studies the effectiveness of the proposed method.

  8. Operational Modal Analysis and Force Characterization of an Unstable Liquid Rocket Engine

    OpenAIRE

    Buechele, Kevin Charles

    2013-01-01

    Combustion instability has plagued the rocket industry since its beginnings. It is characterized by sustained pressure oscillations in the combustion chamber due to the coupling of natural pressure fluctuations with unsteady heat release. Typically, combustion instabilities are identified and mitigated during ground testing. Occasionally combustion instabilities do not present themselves until after a system is fielded and may only occur in a flight environment. While ground tests are heavily...

  9. Spectrally-engineered solar thermal photovoltaic devices

    Energy Technology Data Exchange (ETDEWEB)

    Lenert, Andrej; Bierman, David; Chan, Walker; Celanovic, Ivan; Soljacic, Marin; Wang, Evelyn N.; Nam, Young Suk; McEnaney, Kenneth; Kraemer, Daniel; Chen, Gang

    2018-03-27

    A solar thermal photovoltaic device, and method of forming same, includes a solar absorber and a spectrally selective emitter formed on either side of a thermally conductive substrate. The solar absorber is configured to absorb incident solar radiation. The solar absorber and the spectrally selective emitter are configured with an optimized emitter-to-absorber area ratio. The solar thermal photovoltaic device also includes a photovoltaic cell in thermal communication with the spectrally selective emitter. The spectrally selective emitter is configured to permit high emittance for energies above a bandgap of the photovoltaic cell and configured to permit low emittance for energies below the bandgap.

  10. Modeling of Uneven Flow and Electromagnetic Field Parameters in the Combustion Chamber of Liquid Rocket Engine with a Near-wall Layer Available

    Directory of Open Access Journals (Sweden)

    A. V. Rudinskii

    2015-01-01

    Full Text Available The paper concerns modeling of an uneven flow and electromagnetic field parameters in the combustion chamber of the liquid rocket engine with a near-wall layer available.The research objective was to evaluate quantitatively influence of changing model chamber mode of the liquid rocket engine on the electro-physical characteristics of the hydrocarbon fuel combustion by-products.The main method of research was based on development of a final element model of the flowing path of the rocket engine chamber and its adaptation to the boundary conditions.The paper presents a developed two-dimensional non-stationary mathematical model of electro-physical processes in the liquid rocket engine chamber using hydrocarbon fuel. The model takes into consideration the features of a gas-dynamic contour of the engine chamber and property of thermo-gas-dynamic characteristics of the ionized products of combustion of hydrocarbonic fuel. Distributions of magnetic field intensity and electric conductivity received and analyzed taking into account a low-temperature near-wall layer. Special attention is paid to comparison of obtained calculation values of the electric current, which is taken out from intrachamber space of the engine with earlier published data of other authors.

  11. Research on shock wave characteristics in the isolator of central strut rocket-based combined cycle engine under Ma5.5

    Science.gov (United States)

    Wei, Xianggeng; Xue, Rui; Qin, Fei; Hu, Chunbo; He, Guoqiang

    2017-11-01

    A numerical calculation of shock wave characteristics in the isolator of central strut rocket-based combined cycle (RBCC) engine fueled by kerosene was carried out in this paper. A 3D numerical model was established by the DES method. The kerosene chemical kinetic model used the 9-component and 12-step simplified mechanism model. Effects of fuel equivalence ratio, inflow total temperature and central strut rocket on-off on shock wave characteristics were studied under Ma5.5. Results demonstrated that with the increase of equivalence ratio, the leading shock wave moves toward upstream, accompanied with higher possibility of the inlet unstart. However, the leading shock wave moves toward downstream as the inflow total temperature rises. After the central strut rocket is closed, the leading shock wave moves toward downstream, which can reduce risks of the inlet unstart. State of the shear layer formed by the strut rocket jet flow and inflow can influence the shock train structure significantly.

  12. Models of Non-Stationary Thermodynamic Processes in Rocket Engines Taking into Account a Chemical Equilibrium of Combustion Products

    Directory of Open Access Journals (Sweden)

    A. V. Aliev

    2015-01-01

    Full Text Available The paper considers the two approach-based techniques for calculating the non-stationary intra-chamber processes in solid-propellant rocket engine (SPRE. The first approach assumes that the combustion products are a mechanical mix while the other one supposes it to be the mix, which is in chemical equilibrium. To enhance reliability of solution of the intra ballistic tasks, which assume a chemical equilibrium of combustion products, the computing algorithms to calculate a structure of the combustion products are changed. The algorithm for solving a system of the nonlinear equations of chemical equilibrium, when determining the iterative amendments, uses the orthogonal QR method instead of a method of Gauss. Besides, a possibility to apply genetic algorithms in a task about a structure of combustion products is considered.It is shown that in the tasks concerning the prediction of non-stationary intra ballistic characteristics in a solid propellant rocket engine, application of models of mechanical mix and chemically equilibrium structure of combustion products leads to qualitatively and quantitatively coinciding results. The maximum difference in parameters is 5-10%, at most. In tasks concerning the starting operation of a solid sustainer engine with high-temperature products of combustion difference in results is more essential, and can reach 20% and more.A technique to calculate the intra ballistic parameters, in which flotation of combustion products is considered in the light of a spatial statement, requires using the high-performance computer facilities. For these tasks it is offered to define structure of products of combustion and its thermo-physical characteristics, using the polynoms coefficients of which should be predefined.

  13. Carbon-Carbon Nozzle Extension Development in Support of In-Space and Upper-Stage Liquid Rocket Engines

    Science.gov (United States)

    Gradl, Paul R.; Valentine, Peter G.

    2017-01-01

    Upper stage and in-space liquid rocket engines are optimized for performance through the use of high area ratio nozzles to fully expand combustion gases to low exit pressures, increasing exhaust velocities. Due to the large size of such nozzles, and the related engine performance requirements, carbon-carbon (C-C) composite nozzle extensions are being considered to reduce weight impacts. Currently, the state-of-the-art is represented by the metallic and foreign composite nozzle extensions limited to approximately 2000 degrees F. used on the Atlas V, Delta IV, Falcon 9, and Ariane 5 launch vehicles. NASA and industry partners are working towards advancing the domestic supply chain for C-C composite nozzle extensions. These development efforts are primarily being conducted through the NASA Small Business Innovation Research (SBIR) program in addition to other low level internal research efforts. This has allowed for the initial material development and characterization, subscale hardware fabrication, and completion of hot-fire testing in relevant environments. NASA and industry partners have designed, fabricated and hot-fire tested several subscale domestically produced C-C extensions to advance the material and coatings fabrication technology for use with a variety of liquid rocket and scramjet engines. Testing at NASA's Marshall Space Flight Center (MSFC) evaluated heritage and state-of-the-art C-C materials and coatings, demonstrating the initial capabilities of the high temperature materials and their fabrication methods. This paper discusses the initial material development, design and fabrication of the subscale carbon-carbon nozzle extensions, provides an overview of the test campaign, presents results of the hot fire testing, and discusses potential follow-on development work. The follow on work includes the fabrication of ultra-high temperature materials, larger C-C nozzle extensions, material characterization, sub-element testing and hot-fire testing at

  14. Subscale Carbon-Carbon Nozzle Extension Development and Hot Fire Testing in Support of Upper Stage Liquid Rocket Engines

    Science.gov (United States)

    Gradl, Paul; Valentine, Peter; Crisanti, Matthew; Greene, Sandy Elam

    2016-01-01

    Upper stage and in-space liquid rocket engines are optimized for performance through the use of high area ratio nozzles to fully expand combustion gases to low exit pressures increasing exhaust velocities. Due to the large size of such nozzles and the related engine performance requirements, carbon-carbon (C/C) composite nozzle extensions are being considered for use in order to reduce weight impacts. NASA and industry partner Carbon-Carbon Advanced Technologies (C-CAT) are working towards advancing the technology readiness level of large-scale, domestically-fabricated, C/C nozzle extensions. These C/C extensions have the ability to reduce the overall costs of extensions relative to heritage metallic and composite extensions and to decrease weight by 50%. Material process and coating developments have advanced over the last several years, but hot fire testing to fully evaluate C/C nozzle extensions in relevant environments has been very limited. NASA and C-CAT have designed, fabricated and hot fire tested multiple subscale nozzle extension test articles of various C/C material systems, with the goal of assessing and advancing the manufacturability of these domestically producible materials as well as characterizing their performance when subjected to the typical environments found in a variety of liquid rocket and scramjet engines. Testing at the MSFC Test Stand 115 evaluated heritage and state-of-the-art C/C materials and coatings, demonstrating the capabilities of the high temperature materials and their fabrication methods. This paper discusses the design and fabrication of the 1.2k-lbf sized carbon-carbon nozzle extensions, provides an overview of the test campaign, presents results of the hot fire testing, and discusses potential follow-on development work.

  15. Rocket University at KSC

    Science.gov (United States)

    Sullivan, Steven J.

    2014-01-01

    "Rocket University" is an exciting new initiative at Kennedy Space Center led by NASA's Engineering and Technology Directorate. This hands-on experience has been established to develop, refine & maintain targeted flight engineering skills to enable the Agency and KSC strategic goals. Through "RocketU", KSC is developing a nimble, rapid flight engineering life cycle systems knowledge base. Ongoing activities in RocketU develop and test new technologies and potential customer systems through small scale vehicles, build and maintain flight experience through balloon and small-scale rocket missions, and enable a revolving fresh perspective of engineers with hands on expertise back into the large scale NASA programs, providing a more experienced multi-disciplined set of systems engineers. This overview will define the Program, highlight aspects of the training curriculum, and identify recent accomplishments and activities.

  16. Buffer thermal energy storage for a solar Brayton engine

    Science.gov (United States)

    Strumpf, H. J.; Barr, K. P.

    1981-01-01

    A study has been completed on the application of latent-heat buffer thermal energy storage to a point-focusing solar receiver equipped with an air Brayton engine. To aid in the study, a computer program was written for complete transient/stead-state Brayton cycle performance. The results indicated that thermal storage can afford a significant decrease in the number of engine shutdowns as compared to operating without thermal storage. However, the number of shutdowns does not continuously decrease as the storage material weight increases. In fact, there appears to be an optimum weight for minimizing the number of shutdowns.

  17. Base Flow and Heat Transfer Characteristics of a Four-Nozzle Clustered Rocket Engine: Effect of Nozzle Pressure Ratio

    Science.gov (United States)

    Nallasamy, R.; Kandula, M.; Duncil, L.; Schallhorn, P.

    2010-01-01

    The base pressure and heating characteristics of a four-nozzle clustered rocket configuration is studied numerically with the aid of OVERFLOW Navier-Stokes code. A pressure ratio (chamber pressure to freestream static pressure) range of 990 to 5,920 and a freestream Mach number range of 2.5 to 3.5 are studied. The qualitative trends of decreasing base pressure with increasing pressure ratio and increasing base heat flux with increasing pressure ratio are correctly predicted. However, the predictions for base pressure and base heat flux show deviations from the wind tunnel data. The differences in absolute values between the computation and the data are attributed to factors such as perfect gas (thermally and calorically perfect) assumption, turbulence model inaccuracies in the simulation, and lack of grid adaptation.

  18. Design of Cooling Channels of Preburners for Small Liquid Rocket Engines with Computational Flow and Heat Transfer Analysis

    Directory of Open Access Journals (Sweden)

    Insang Moon

    2011-09-01

    Full Text Available A series of computational analyses was performed to predict the cooling process by the cooling channel of preburners used for kerosene-liquid oxygen staged combustion cycle rocket engines. As an oxygen-rich combustion occurs in the kerosene fueled preburner, it is of great importance to control the wall temperature so that it does not exceed the critical temperature. However, since the heat transfer is proportional to the speed of fluid running inside the channel, the high heat transfer leads to a trade-off of pressure loss. For this reason, it is necessary to establish a certain criteria between the pressure loss and the heat transfer or the wall surface temperature. The design factors of the cooling channel were determined by the computational research, and a test model was manufactured. The test model was used for the hot fire tests to prove the function of the cooling mechanism, among other purposes.

  19. The issue of ensuring the safe explosion of the spent orbital stages of a launch vehicle with propulsion rocket engine

    Directory of Open Access Journals (Sweden)

    Trushlyakov Valeriy I.

    2017-01-01

    Full Text Available A method for increasing the safe explosion of the spent orbital stages of a space launch vehicle (SLV with a propulsion rocket engine (PRE based on the gasification of unusable residues propellant and venting fuel tanks. For gasification and ventilation the hot gases used produced by combustion of the specially selected gas generating composition (GGC with a set of physical and chemical properties. Excluding the freezing of the drainage system on reset gasified products (residues propellant+pressurization gas+hot gases in the near-Earth space is achieved by selecting the physical-chemical characteristics of the GGC. Proposed steps to ensure rotation of gasified products due to dumping through the drainage system to ensure the most favorable conditions for propellant gasification residues. For example, a tank with liquid oxygen stays with the orbital spent second stage of the SLV “Zenit”, which shows the effectiveness of the proposed method.

  20. The Technique for CFD-Simulation of Fuel Valve from Pneumatic-Hydraulic System of Liquid-Propellant Rocket Engine

    Science.gov (United States)

    Shabliy, L. S.; Malov, D. V.; Bratchinin, D. S.

    2018-01-01

    In the article the description of technique for simulation of valves for pneumatic-hydraulic system of liquid-propellant rocket engine (LPRE) is given. Technique is based on approach of computational hydrodynamics (Computational Fluid Dynamics – CFD). The simulation of a differential valve used in closed circuit LPRE supply pipes of fuel components is performed to show technique abilities. A schematic and operation algorithm of this valve type is described in detail. Also assumptions made in the construction of the geometric model of the hydraulic path of the valve are described in detail. The calculation procedure for determining valve hydraulic characteristics is given. Based on these calculations certain hydraulic characteristics of the valve are given. Some ways of usage of the described simulation technique for research the static and dynamic characteristics of the elements of the pneumatic-hydraulic system of LPRE are proposed.

  1. Orbit transfer rocket engine technology program: Automated preflight methods concept definition

    Science.gov (United States)

    Erickson, C. M.; Hertzberg, D. W.

    1991-01-01

    The possibility of automating preflight engine checkouts on orbit transfer engines is discussed. The minimum requirements in terms of information and processing necessary to assess the engine'e integrity and readiness to perform its mission were first defined. A variety of ways for remotely obtaining that information were generated. The sophistication of these approaches varied from a simple preliminary power up, where the engine is fired up for the first time, to the most advanced approach where the sensor and operational history data system alone indicates engine integrity. The critical issues and benefits of these methods were identified, outlined, and prioritized. The technology readiness of each of these automated preflight methods were then rated on a NASA Office of Exploration scale used for comparing technology options for future mission choices. Finally, estimates were made of the remaining cost to advance the technology for each method to a level where the system validation models have been demonstrated in a simulated environment.

  2. Thermal Barrier Coatings for Advanced Gas Turbine and Diesel Engines

    Science.gov (United States)

    Zhu, Dongming; Miller, Robert A.

    1999-01-01

    Ceramic thermal barrier coatings (TBCS) have been developed for advanced gas turbine and diesel engine applications to improve engine reliability and fuel efficiency. However, durability issues of these thermal barrier coatings under high temperature cyclic conditions are still of major concern. The coating failure depends not only on the coating, but also on the ceramic sintering/creep and bond coat oxidation under the operating conditions. Novel test approaches have been established to obtain critical thermomechanical and thermophysical properties of the coating systems under near-realistic transient and steady state temperature and stress gradients encountered in advanced engine systems. This paper presents detailed experimental and modeling results describing processes occurring in the ZrO2-Y2O3 thermal barrier coating systems, thus providing a framework for developing strategies to manage ceramic coating architecture, microstructure and properties.

  3. Observed and modelled effects of auroral precipitation on the thermal ionospheric plasma: comparing the MICA and Cascades2 sounding rocket events

    Science.gov (United States)

    Lynch, K. A.; Gayetsky, L.; Fernandes, P. A.; Zettergren, M. D.; Lessard, M.; Cohen, I. J.; Hampton, D. L.; Ahrns, J.; Hysell, D. L.; Powell, S.; Miceli, R. J.; Moen, J. I.; Bekkeng, T.

    2012-12-01

    Auroral precipitation can modify the ionospheric thermal plasma through a variety of processes. We examine and compare the events seen by two recent auroral sounding rockets carrying in situ thermal plasma instrumentation. The Cascades2 sounding rocket (March 2009, Poker Flat Research Range) traversed a pre-midnight poleward boundary intensification (PBI) event distinguished by a stationary Alfvenic curtain of field-aligned precipitation. The MICA sounding rocket (February 2012, Poker Flat Research Range) traveled through irregular precipitation following the passage of a strong westward-travelling surge. Previous modelling of the ionospheric effects of auroral precipitation used a one-dimensional model, TRANSCAR, which had a simplified treatment of electric fields and did not have the benefit of in situ thermal plasma data. This new study uses a new two-dimensional model which self-consistently calculates electric fields to explore both spatial and temporal effects, and compares to thermal plasma observations. A rigorous understanding of the ambient thermal plasma parameters and their effects on the local spacecraft sheath and charging, is required for quantitative interpretation of in situ thermal plasma observations. To complement this TRANSCAR analysis we therefore require a reliable means of interpreting in situ thermal plasma observation. This interpretation depends upon a rigorous plasma sheath model since the ambient ion energy is on the order of the spacecraft's sheath energy. A self-consistent PIC model is used to model the spacecraft sheath, and a test-particle approach then predicts the detector response for a given plasma environment. The model parameters are then modified until agreement is found with the in situ data. We find that for some situations, the thermal plasma parameters are strongly driven by the precipitation at the observation time. For other situations, the previous history of the precipitation at that position can have a stronger

  4. Ramjets: Thermal Management an Integrated Engineering Approach

    Science.gov (United States)

    2010-09-01

    and Thermal Management (Propulsion a vitesse elevee : Conception du moteur - integration et gestion thermique ) 14. ABSTRACT Within the framework of...2.0 AERODYNAMIC HEATING 2.1 General Heat Transfer Relations The air flow around any vehicle moving through the atmosphere comes to rest at the...resulting in a convective heat flux from the air flow to the structure of the vehicle. The basic equation describing convective heat transfer is: )( wawcc

  5. Technical engineering services in support of the Nike-Tomahawk sounding rocket vehicle system

    Science.gov (United States)

    1972-01-01

    Task assignments in support of the Nike-Tomahawk vehicles, which were completed from May, 1970 through November 1972 are reported. The services reported include: analytical, design and drafting, fabrication and modification, and field engineering.

  6. Air Force Research Laboratory's Rocket Engine Program Enters Fast-Paced Test Phase

    National Research Council Canada - National Science Library

    Thornburg, Jeff

    2002-01-01

    .... Recent tests of the Integrated Powerhead Demonstration project here established a technical first for the United States and mark the first advancements in boost engine technology since the space...

  7. Measurement and Modeling of Surface Coking in Fuel-Film Cooled Liquid Rocket Engines

    Data.gov (United States)

    National Aeronautics and Space Administration — The development of future Kerosone/LOX engines will require higher chamber pressures to increase performance and reusability in order to decrease operating costs....

  8. Multiphysics Framework for Prediction of Dynamic Instability in Liquid Rocket Engines, Phase I

    Data.gov (United States)

    National Aeronautics and Space Administration — Mitigation of dynamic combustion instability is one of the most difficult engineering challenges facing NASA and industry in the development of new continuous-flow...

  9. Rankine-Brayton engine powered solar thermal aircraft

    Science.gov (United States)

    Bennett, Charles L [Livermore, CA

    2009-12-29

    A solar thermal powered aircraft powered by heat energy from the sun. A Rankine-Brayton hybrid cycle heat engine is carried by the aircraft body for producing power for a propulsion mechanism, such as a propeller or other mechanism for enabling sustained free flight. The Rankine-Brayton engine has a thermal battery, preferably containing a lithium-hydride and lithium mixture, operably connected to it so that heat is supplied from the thermal battery to a working fluid. A solar concentrator, such as reflective parabolic trough, is movably connected to an optically transparent section of the aircraft body for receiving and concentrating solar energy from within the aircraft. Concentrated solar energy is collected by a heat collection and transport conduit, and heat transported to the thermal battery. A solar tracker includes a heliostat for determining optimal alignment with the sun, and a drive motor actuating the solar concentrator into optimal alignment with the sun based on a determination by the heliostat.

  10. Rankline-Brayton engine powered solar thermal aircraft

    Science.gov (United States)

    Bennett, Charles L [Livermore, CA

    2012-03-13

    A solar thermal powered aircraft powered by heat energy from the sun. A Rankine-Brayton hybrid cycle heat engine is carried by the aircraft body for producing power for a propulsion mechanism, such as a propeller or other mechanism for enabling sustained free flight. The Rankine-Brayton engine has a thermal battery, preferably containing a lithium-hydride and lithium mixture, operably connected to it so that heat is supplied from the thermal battery to a working fluid. A solar concentrator, such as reflective parabolic trough, is movably connected to an optically transparent section of the aircraft body for receiving and concentrating solar energy from within the aircraft. Concentrated solar energy is collected by a heat collection and transport conduit, and heat transported to the thermal battery. A solar tracker includes a heliostat for determining optimal alignment with the sun, and a drive motor actuating the solar concentrator into optimal alignment with the sun based on a determination by the heliostat.

  11. A Collaborative Analysis Tool for Integrating Hypersonic Aerodynamics, Thermal Protection Systems, and RBCC Engine Performance for Single Stage to Orbit Vehicles

    Science.gov (United States)

    Stanley, Thomas Troy; Alexander, Reginald

    1999-01-01

    Presented is a computer-based tool that connects several disciplines that are needed in the complex and integrated design of high performance reusable single stage to orbit (SSTO) vehicles. Every system is linked to every other system, as is the case of SSTO vehicles with air breathing propulsion, which is currently being studied by NASA. The deficiencies in the scramjet powered concept led to a revival of interest in Rocket-Based Combined-Cycle (RBCC) propulsion systems. An RBCC propulsion system integrates airbreathing and rocket propulsion into a single engine assembly enclosed within a cowl or duct. A typical RBCC propulsion system operates as a ducted rocket up to approximately Mach 3. At this point the transitions to a ramjet mode for supersonic-to-hypersonic acceleration. Around Mach 8 the engine transitions to a scram4jet mode. During the ramjet and scramjet modes, the integral rockets operate as fuel injectors. Around Mach 10-12 (the actual value depends on vehicle and mission requirements), the inlet is physically closed and the engine transitions to an integral rocket mode for orbit insertion. A common feature of RBCC propelled vehicles is the high degree of integration between the propulsion system and airframe. At high speeds the vehicle forebody is fundamentally part of the engine inlet, providing a compression surface for air flowing into the engine. The compressed air is mixed with fuel and burned. The combusted mixture must be expanded to an area larger than the incoming stream to provide thrust. Since a conventional nozzle would be too large, the entire lower after body of the vehicle is used as an expansion surface. Because of the high external temperatures seen during atmospheric flight, the design of an airbreathing SSTO vehicle requires delicate tradeoffs between engine design, vehicle shape, and thermal protection system (TPS) sizing in order to produce an optimum system in terms of weight (and cost) and maximum performance.

  12. Theoretical and experimental analysis of liquid layer instability in hybrid rocket engines

    Science.gov (United States)

    Kobald, Mario; Verri, Isabella; Schlechtriem, Stefan

    2015-03-01

    The combustion behavior of different hybrid rocket fuels has been analyzed in the frame of this research. Tests have been performed in a 2D slab burner configuration with windows on two sides. Four different liquefying paraffin-based fuels, hydroxyl terminated polybutadiene (HTPB) and high-density polyethylene (HDPE) have been tested in combination with gaseous oxygen (GOX). Experimental high-speed video data have been analyzed manually and with the proper orthogonal decomposition (POD) technique. Application of POD enables the recognition of the main structures of the flow field and the combustion flame appearing in the video data. These results include spatial and temporal analysis of the structures. For liquefying fuels these spatial values relate to the wavelengths associated to the Kelvin Helmholtz Instability (KHI). A theoretical long-wave solution of the KHI problem shows good agreement with the experimental results. Distinct frequencies found in the POD analysis can be related to the precombustion chamber configuration which can lead to vortex shedding phenomena.

  13. Thermal Efficiency of a Combined Turbocharger Set with Gasoline Engine

    OpenAIRE

    Jarut Kunanoppadon

    2010-01-01

    Problem statement: The technology of turbocharger has been used with internal combustion engines since 1905 to increase intake air pressure prior to putting it into the cylinders to increase thermal efficiency of the engine. Based on previous researches and uses of turbochargers, the pattern of turbocharger installation remains the same, either in series or in parallel. Therefore, this research aims to study installation of the combined turbocharger. The combined turbocharger set comprised tw...

  14. Study on the Effect of water Injection Momentum on the Cooling Effect of Rocket Engine Exhaust Plume

    Science.gov (United States)

    Yang, Kan; Qiang, Yanhui; Zhong, Chenghang; Yu, Shaozhen

    2017-10-01

    For the study of water injection momentum factors impact on flow field of the rocket engine tail flame, the numerical computation model of gas-liquid two phase flow in the coupling of high temperature and high speed gas flow and low temperature liquid water is established. The accuracy and reliability of the numerical model are verified by experiments. Based on the numerical model, the relationship between the flow rate and the cooling effect is analyzed by changing the water injection momentum of the water spray pipes. And the effective mathematical expression is obtained. What’s more, by changing the number of the water spray and using small flow water injection, the cooling effect is analyzed to check the application range of the mathematical expressions. The results show that: the impact and erosion of the gas flow field could be reduced greatly by water injection, and there are two parts in the gas flow field, which are the slow cooling area and the fast cooling area. In the fast cooling area, the influence of the water flow momentum and nozzle quantity on the cooling effect can be expressed by mathematical functions without causing bifurcation flow for the mainstream gas. The conclusion provides a theoretical reference for the engineering application.

  15. Space Launch System Base Heating Test: Sub-Scale Rocket Engine/Motor Design, Development and Performance Analysis

    Science.gov (United States)

    Mehta, Manish; Seaford, Mark; Kovarik, Brian; Dufrene, Aaron; Solly, Nathan; Kirchner, Robert; Engel, Carl D.

    2014-01-01

    The Space Launch System (SLS) base heating test is broken down into two test programs: (1) Pathfinder and (2) Main Test. The Pathfinder Test Program focuses on the design, development, hot-fire test and performance analyses of the 2% sub-scale SLS core-stage and booster element propulsion systems. The core-stage propulsion system is composed of four gaseous oxygen/hydrogen RS-25D model engines and the booster element is composed of two aluminum-based model solid rocket motors (SRMs). The first section of the paper discusses the motivation and test facility specifications for the test program. The second section briefly investigates the internal flow path of the design. The third section briefly shows the performance of the model RS-25D engines and SRMs for the conducted short duration hot-fire tests. Good agreement is observed based on design prediction analysis and test data. This program is a challenging research and development effort that has not been attempted in 40+ years for a NASA vehicle.

  16. A Framework for Assessing the Reusability of Hardware (Reusable Rocket Engines)

    Science.gov (United States)

    Childress-Thompson, Rhonda; Thomas, Dale; Farrington, Philip

    2016-01-01

    Within the past few years, there has been a renewed interest in reusability as it applies to space flight hardware. Commercial companies such as Space Exploration Technologies Corporation (SpaceX), Blue Origin, and United Launch Alliance (ULA) are pursuing reusable hardware. Even foreign companies are pursuing this option. The Indian Space Research Organization (ISRO) launched a reusable space plane technology demonstrator and Airbus Defense and Space is planning to recover the main engines and avionics from its Advanced Expendable Launcher with Innovative engine Economy [1] [2]. To date, the Space Shuttle remains as the only Reusable Launch (RLV) to have flown repeated missions and the Space Shutte Main Engine (SSME) is the only demonstrated reusable engine. Whether the hardware being considered for reuse is a launch vehicle (fully reusable), a first stage (partially reusable), or a booster engine (single component), the overall governing process is the same; it must be recovered and recertified for flight. Therefore, there is a need to identify the key factors in determining the reusability of flight hardware. This paper begins with defining reusability to set the context, addresses the significance of reuse, and discusses areas that limit successful implementation. Finally, this research identifies the factors that should be considered when incorporating reuse.

  17. ELM - A SIMPLE TOOL FOR THERMAL-HYDRAULIC ANALYSIS OF SOLID-CORE NUCLEAR ROCKET FUEL ELEMENTS

    Science.gov (United States)

    Walton, J. T.

    1994-01-01

    ELM is a simple computational tool for modeling the steady-state thermal-hydraulics of propellant flow through fuel element coolant channels in nuclear thermal rockets. Written for the nuclear propulsion project of the Space Exploration Initiative, ELM evaluates the various heat transfer coefficient and friction factor correlations available for turbulent pipe flow with heat addition. In the past, these correlations were found in different reactor analysis codes, but now comparisons are possible within one program. The logic of ELM is based on the one-dimensional conservation of energy in combination with Newton's Law of Cooling to determine the bulk flow temperature and the wall temperature across a control volume. Since the control volume is an incremental length of tube, the corresponding pressure drop is determined by application of the Law of Conservation of Momentum. The size, speed, and accuracy of ELM make it a simple tool for use in fuel element parametric studies. ELM is a machine independent program written in FORTRAN 77. It has been successfully compiled on an IBM PC compatible running MS-DOS using Lahey FORTRAN 77, a DEC VAX series computer running VMS, and a Sun4 series computer running SunOS UNIX. ELM requires 565K of RAM under SunOS 4.1, 360K of RAM under VMS 5.4, and 406K of RAM under MS-DOS. Because this program is machine independent, no executable is provided on the distribution media. The standard distribution medium for ELM is one 5.25 inch 360K MS-DOS format diskette. ELM was developed in 1991. DEC, VAX, and VMS are trademarks of Digital Equipment Corporation. Sun4 and SunOS are trademarks of Sun Microsystems, Inc. IBM PC is a registered trademark of International Business Machines. MS-DOS is a registered trademark of Microsoft Corporation.

  18. RP-2 Thermal Stability and Heat Transfer Investigation for Hydrocarbon Boost Engines

    Science.gov (United States)

    VanNoord, J. L.; Stiegemeier, B. R.

    2010-01-01

    A series of electrically heated tube tests were performed at the NASA Glenn Research Center s Heated Tube Facility to investigate the use of RP-2 as a fuel for next generation regeneratively cooled hydrocarbon boost engines. The effect that test duration, operating condition and test piece material have on the overall thermal stability and materials compatibility characteristics of RP-2 were evaluated using copper and 304 stainless steel test sections. The copper tests were run at 1000 psia, heat flux up to 6.0 Btu/in.2-sec, and wall temperatures up to 1180 F. Preliminary results, using measured wall temperature as an indirect indicator of the carbon deposition process, show that in copper test pieces above approximately 850 F, RP-2 begins to undergo thermal decomposition resulting in local carbon deposits. Wall temperature traces show significant local temperature increases followed by near instantaneous drops which have been attributed to the carbon deposition/shedding process in previous investigations. Data reduction is currently underway for the stainless steel test sections and carbon deposition measurements will be performed in the future for all test sections used in this investigation. In conjunction with the existing thermal stability database, these findings give insight into the feasibility of cooling a long life, high performance, high-pressure liquid rocket combustor and nozzle with RP-2.

  19. A computer program for performance prediction of tripropellant rocket engines with tangential slot injection

    Science.gov (United States)

    Dang, Anthony; Nickerson, Gary R.

    1987-01-01

    For the development of a Heavy Lift Launch Vehicle (HLLV) several engines with different operating cycles and using LOX/Hydrocarbon propellants are presently being examined. Some concepts utilize hydrogen for thrust chamber wall cooling followed by a gas generator turbine drive cycle with subsequent dumping of H2/O2 combustion products into the nozzle downstream of the throat. In the Space Transportation Booster Engine (STBE) selection process the specific impulse will be one of the optimization criteria; however, the current performance prediction programs do not have the capability to include a third propellant in this process, nor to account for the effect of dumping the gas-generator product tangentially inside the nozzle. The purpose is to describe a computer program for accurately predicting the performance of such an engine. The code consists of two modules; one for the inviscid performance, and the other for the viscous loss. For the first module, the two-dimensional kinetics program (TDK) was modified to account for tripropellant chemistry, and for the effect of tangential slot injection. For the viscous loss, the Mass Addition Boundary Layer program (MABL) was modified to include the effects of the boundary layer-shear layer interaction, and tripropellant chemistry. Calculations were made for a real engine and compared with available data.

  20. Unsteady flowfield in an integrated rocket ramjet engine and combustion dynamics of a gas turbine swirl-stabilized injector

    Science.gov (United States)

    Sung, Hong-Gye

    This research focuses on the time-accurate simulation and analysis of the unsteady flowfield in an integrated rocket-ramjet engine (IRR) and combustion dynamics of a swirl-stabilized gas turbine engine. The primary objectives are: (1) to establish a unified computational framework for studying unsteady flow and flame dynamics in ramjet propulsion systems and gas turbine combustion chambers, and (2) to investigate the parameters and mechanisms responsible for driving flow oscillations. The first part of the thesis deals with a complete axi-symmetric IRR engine. The domain of concern includes a supersonic inlet diffuser, a combustion chamber, and an exhaust nozzle. This study focused on the physical mechanism of the interaction between the oscillatory terminal shock in the inlet diffuser and the flame in the combustion chamber. In addition, the flow and ignition transitions from the booster to the sustainer phase were analyzed comprehensively. Even though the coupling between the inlet dynamics and the unsteady motions of flame shows that they are closely correlated, fortunately, those couplings are out of phase with a phase lag of 90 degrees, which compensates for the amplification of the pressure fluctuation in the inlet. The second part of the thesis treats the combustion dynamics of a lean-premixed gas turbine swirl injector. A three-dimensional computation method utilizing the message passing interface (MPI) Parallel architecture and large-eddy-simulation technique was applied. Vortex breakdown in the swirling flow is clearly visualized and explained on theoretical bases. The unsteady turbulent flame dynamics are carefully simulated so that the flow motion can be characterized in detail. It was observed that some fuel lumps escape from the primary combustion zone, and move downstream and consequently produce hot spots and large vortical structures in the azimuthal direction. The correlation between pressure oscillation and unsteady heat release is examined by

  1. Thermal analysis of the effect of thick thermal barrier coatings on diesel engine performance

    International Nuclear Information System (INIS)

    Hoag, K.L.; Frisch, S.R.; Yonushonis, T.M.

    1986-01-01

    The reduction of heat rejection from the diesel engine combustion chamber has been the subject of a great deal of focus in recent years. In the pursuit of this goal, Cummins Engine Company has received a contract from the Department of Energy for the development of thick thermal barrier coatings for combustion chamber surfaces. This contract involves the analysis of the impact of coatings on diesel engine performance, bench test evaluation of various coating designs, and single cylinder engine tests. The efforts reported in this paper center on the analysis of the effects of coatings on engine performance and heat rejection. For this analysis the conventional water cooled engine was compared with an engine having limited oil cooling, and utilizing zirocnia coated cylinder had firedecks and piston crowns. The analysis showed little or no benefits of similarly coating the valves or cylinder liner

  2. Space Launch System Base Heating Test: Sub-Scale Rocket Engine/Motor Design, Development & Performance Analysis

    Science.gov (United States)

    Mehta, Manish; Seaford, Mark; Kovarik, Brian; Dufrene, Aaron; Solly, Nathan

    2014-01-01

    ATA-002 Technical Team has successfully designed, developed, tested and assessed the SLS Pathfinder propulsion systems for the Main Base Heating Test Program. Major Outcomes of the Pathfinder Test Program: Reach 90% of full-scale chamber pressure Achieved all engine/motor design parameter requirements Reach steady plume flow behavior in less than 35 msec Steady chamber pressure for 60 to 100 msec during engine/motor operation Similar model engine/motor performance to full-scale SLS system Mitigated nozzle throat and combustor thermal erosion Test data shows good agreement with numerical prediction codes Next phase of the ATA-002 Test Program Design & development of the SLS OML for the Main Base Heating Test Tweak BSRM design to optimize performance Tweak CS-REM design to increase robustness MSFC Aerosciences and CUBRC have the capability to develop sub-scale propulsion systems to meet desired performance requirements for short-duration testing.

  3. Electric field and radio frequency measurements for rocket engine health monitoring applications

    Science.gov (United States)

    Valenti, Elizabeth L.

    1992-01-01

    Electric-field (EF) and radio-frequency (RF) emissions generated in the exhaust plumes of the diagnostic testbed facility thruster (DTFT) and the SSME are examined briefly for potential applications to plume diagnostics and engine health monitoring. Hypothetically, anomalous engine conditions could produce measurable changes in any characteristic EF and RF spectral signatures identifiable with a 'healthly' plumes. Tests to determine the presence of EF and RF emissions in the DTFT and SSME exhaust plumes were conducted. EF and RF emissions were detected using state-of-the-art sensors. Analysis of limited data sets show some apparent consistencies in spectral signatures. Significant emissions increases were detected during controlled tests using dopants injected into the DTFT.

  4. 'Bimodal' Nuclear Thermal Rocket (BNTR) propulsion for an artificial gravity HOPE mission to Callisto

    International Nuclear Information System (INIS)

    Borowski, Stanley K.; McGuire, Melissa L.; Mason, Lee M.; Gilland, James H.; Packard, Thomas W.

    2003-01-01

    This paper summarizes the results of a year long, multi-center NASA study which examined the viability of nuclear fission propulsion systems for Human Outer Planet Exploration (HOPE). The HOPE mission assumes a crew of six is sent to Callisto. Jupiter's outermost large moon, to establish a surface base and propellant production facility. The Asgard asteroid formation, a region potentially rich in water-ice, is selected as the landing site. High thrust BNTR propulsion is used to transport the crew from the Earth-Moon L1 staging node to Callisto then back to Earth in less than 5 years. Cargo and LH2 'return' propellant for the piloted Callisto transfer vehicle (PCTV) is pre-deployed at the moon (before the crew's departure) using low thrust, high power, nuclear electric propulsion (NEP) cargo and tanker vehicles powered by hydrogen magnetoplasmadynamic (MPD) thrusters. The PCTV is powered by three 25 klbf BNTR engines which also produce 50 kWe of power for crew life support and spacecraft operational needs. To counter the debilitating effects of long duration space flight (∼855 days out and ∼836 days back) under '0-gE' conditions, the PCTV generates an artificial gravity environment of '1-gE' via rotation of the vehicle about its center-of-mass at a rate of ∼4 rpm. After ∼123 days at Callisto, the 'refueled' PCTV leaves orbit for the trip home. Direct capsule re-entry of the crew at mission end is assumed. Dynamic Brayton power conversion and high temperature uranium dioxide (UO2) in tungsten metal ''cermet'' fuel is used in both the BNTR and NEP vehicles to maximize hardware commonality. Technology performance levels and vehicle characteristics are presented, and requirements for PCTV reusability are also discussed

  5. Rocket engine plume diagnostics using video digitization and image processing - Analysis of start-up

    Science.gov (United States)

    Disimile, P. J.; Shoe, B.; Dhawan, A. P.

    1991-01-01

    Video digitization techniques have been developed to analyze the exhaust plume of the Space Shuttle Main Engine. Temporal averaging and a frame-by-frame analysis provide data used to evaluate the capabilities of image processing techniques for use as measurement tools. Capabilities include the determination of the necessary time requirement for the Mach disk to obtain a fully-developed state. Other results show the Mach disk tracks the nozzle for short time intervals, and that dominate frequencies exist for the nozzle and Mach disk movement.

  6. Torpedo Rockets

    Science.gov (United States)

    2004-01-01

    All through the 13th to the 15th Centuries there were reports of many rocket experiments. For example, Joanes de Fontana of Italy designed a surface-rurning, rocket-powered torpedo for setting enemy ships on fire

  7. Rocket Flight.

    Science.gov (United States)

    Van Evera, Bill; Sterling, Donna R.

    2002-01-01

    Describes an activity for designing, building, and launching rockets that provides students with an intrinsically motivating and real-life application of what could have been classroom-only concepts. Includes rocket design guidelines and a sample grading rubric. (KHR)

  8. Rocket-borne thermal plasma instrument "MIPEX" for the ionosphere D, E layer in-situ measurements

    Science.gov (United States)

    Fang, H. K.; Chen, A. B. C.; Lin, C. C. H.; Wu, T. J.; Liu, K. S.; Chuang, C. W.

    2017-12-01

    In this presentation, the design concepts, performances and status of a thermal plasma particle instrument package "Mesosphere and Ionosphere Plasma Exploration complex (MIPEX)", which is going to be installed onboard a NSPO-funded hybrid rocket, to investigate the electrodynamic processes in ionosphere D, E layers above Taiwan are reported. MIPEX is capable of measuring plasma characteristics including ion temperature, ion composition, ion drift, electron temperature and plasma density at densities as low as 1-10 cm-1. This instrument package consists of an improved retarding potential analyzer with a channel electron multiplier (CEM), a simplified ion drift meter and a planar Langmuir probe. To achieve the working atmospheric pressure of CEM at the height of lower D layer ( 70km), a portable vacuum pump is also placed in the package. A prototype set of the MIPEX has been developed and tested in the Space Plasma Operation Chamber (SPOC) at NCKU, where in ionospheric plasma is generated by back-diffusion plasma sources. A plasma density of 10-106 cm-1, ion temperature of 300-1500 K and electron temperature of 1000-3000K is measured and verified. Limited by the flight platform and the performance of the instruments, the in-situ plasma measurements at the Mesosphere and lower Thermosphere is very challenging and rare. MIPEX is capable of extending the altitude of the effective plasma measurement down to 70 km height and this experiment can provide unique high-quality data of the plasma environment to explore the ion distribution and the electrodynamic processes in the Ionosphere D, E layers at dusk.

  9. Rocket engine high-enthalpy flow simulation using heated CO2 gas to verify the development of a rocket nozzle and combustion tests

    Science.gov (United States)

    Takeishi, K.; Ishizaka, K.; Okamoto, J.; Watanabe, Y.

    2017-03-01

    The LE-7A engine is the first-stage engine of the Japanese-made H-IIA launch vehicle. This engine has been developed by improving and reducing the price of the LE-7 engine used in the H-II launch vehicle. In the qualification combustion tests, the original designed LE-7A (LE-7A-OR) engine experienced two major problems, a large side load in the transient state of engine start and stop and melt on nozzle generative cooling tubes. The reason for the troubles of the LE-7A-OR engine was investigated by conducting experimental and numerical studies. In actual engine conditions, the main hot gas stream is a heated steam. Furthermore, the main stream temperature in the nozzle changes from approximately 3500 K at the throat to 500 K at the exit. In such a case, the specific heat ratio changes depending on the temperature. A similarity of the Mach number should be considered when conducting a model flow test with a similar flow condition of the Mach number between an actual engine combustion test and a model flow test. High-speed flow tests were conducted using CO2 gas heated up to 673 K as a working fluid and a 1:12 sub-scaled model nozzle of the LE-7A-OR engine configuration. The problems of the side force and the conducted form of the shock waves generated in the nozzle of the LE-7A-OR engine during engine start and stop were reproduced by the model tests of experimental and numerical investigations. This study presented that the model flow test using heated CO2 gas is useful and effective in verifying the numerical analysis and the design verification before actual engine combustion tests.

  10. Some Interesting Applications of Probabilistic Techiques in Structural Dynamic Analysis of Rocket Engines

    Science.gov (United States)

    Brown, Andrew M.

    2014-01-01

    Numerical and Analytical methods developed to determine damage accumulation in specific engine components when speed variation included. Dither Life Ratio shown to be well over factor of 2 for specific example. Steady-State assumption shown to be accurate for most turbopump cases, allowing rapid calculation of DLR. If hot-fire speed data unknown, Monte Carlo method developed that uses speed statistics for similar engines. Application of techniques allow analyst to reduce both uncertainty and excess conservatism. High values of DLR could allow previously unacceptable part to pass HCF criteria without redesign. Given benefit and ease of implementation, recommend that any finite life turbomachine component analysis adopt these techniques. Probability Values calculated, compared, and evaluated for several industry-proposed methods for combining random and harmonic loads. Two new excel macros written to calculate combined load for any specific probability level. Closed form Curve fits generated for widely used 3(sigma) and 2(sigma) probability levels. For design of lightweight aerospace components, obtaining accurate, reproducible, statistically meaningful answer critical.

  11. Engineering Aerothermal Analysis for X-34 Thermal Protection System Design

    Science.gov (United States)

    Wurster, Kathryn E.; Riley, Christopher J.; Zoby, E. Vincent

    1998-01-01

    Design of the thermal protection system for any hypersonic flight vehicle requires determination of both the peak temperatures over the surface and the heating-rate history along the flight profile. In this paper, the process used to generate the aerothermal environments required for the X-34 Testbed Technology Demonstrator thermal protection system design is described as it has evolved from a relatively simplistic approach based on engineering methods applied to critical areas to one of detailed analyses over the entire vehicle. A brief description of the trajectory development leading to the selection of the thermal protection system design trajectory is included. Comparisons of engineering heating predictions with wind-tunnel test data and with results obtained using a Navier-Stokes flowfield code and an inviscid/boundary layer method are shown. Good agreement is demonstrated among all these methods for both the ground-test condition and the peak heating flight condition. Finally, the detailed analysis using engineering methods to interpolate the surface-heating-rate results from the inviscid/boundary layer method to predict the required thermal environments is described and results presented.

  12. Perspective of Micro Process Engineering for Thermal Food Treatment.

    Science.gov (United States)

    Mathys, Alexander

    2018-01-01

    Micro process engineering as a process synthesis and intensification tool enables an ultra-short thermal treatment of foods within milliseconds (ms) using very high surface-area-to-volume ratios. The innovative application of ultra-short pasteurization and sterilization at high temperatures, but with holding times within the range of ms would allow the preservation of liquid foods with higher qualities, thereby avoiding many unwanted reactions with different temperature-time characteristics. Process challenges, such as fouling, clogging, and potential temperature gradients during such conditions need to be assessed on a case by case basis and optimized accordingly. Owing to the modularity, flexibility, and continuous operation of micro process engineering, thermal processes from the lab to the pilot and industrial scales can be more effectively upscaled. A case study on thermal inactivation demonstrated the feasibility of transferring lab results to the pilot scale. It was shown that micro process engineering applications in thermal food treatment may be relevant to both research and industrial operations. Scaling of micro structured devices is made possible through the use of numbering-up approaches; however, reduced investment costs and a hygienic design must be assured.

  13. A quantum Szilard engine without heat from a thermal reservoir

    Science.gov (United States)

    Hamed Mohammady, M.; Anders, Janet

    2017-11-01

    We study a quantum Szilard engine that is not powered by heat drawn from a thermal reservoir, but rather by projective measurements. The engine is constituted of a system { S }, a weight { W }, and a Maxwell demon { D }, and extracts work via measurement-assisted feedback control. By imposing natural constraints on the measurement and feedback processes, such as energy conservation and leaving the memory of the demon intact, we show that while the engine can function without heat from a thermal reservoir, it must give up at least one of the following features that are satisfied by a standard Szilard engine: (i) repeatability of measurements; (ii) invariant weight entropy; or (iii) positive work extraction for all measurement outcomes. This result is shown to be a consequence of the Wigner–Araki–Yanase theorem, which imposes restrictions on the observables that can be measured under additive conservation laws. This observation is a first-step towards developing ‘second-law-like’ relations for measurement-assisted feedback control beyond thermality.

  14. Orbit transfer rocket engine integrated control and health monitoring system technology readiness assessment

    Science.gov (United States)

    Bickford, R. L.; Collamore, F. N.; Gage, M. L.; Morgan, D. B.; Thomas, E. R.

    1992-01-01

    The objectives of this task were to: (1) estimate the technology readiness of an integrated control and health monitoring (ICHM) system for the Aerojet 7500 lbF Orbit Transfer Vehicle engine preliminary design assuming space based operations; and (2) estimate the remaining cost to advance this technology to a NASA defined 'readiness level 6' by 1996 wherein the technology has been demonstrated with a system validation model in a simulated environment. The work was accomplished through the conduct of four subtasks. In subtask 1 the minimally required functions for the control and monitoring system was specified. The elements required to perform these functions were specified in Subtask 2. In Subtask 3, the technology readiness level of each element was assessed. Finally, in Subtask 4, the development cost and schedule requirements were estimated for bringing each element to 'readiness level 6'.

  15. Effect of Surface Impulsive Thermal Loads on Fatigue Behavior of Constant Volume Propulsion Engine Combustor Materials

    National Research Council Canada - National Science Library

    Zhu, Dongming

    2004-01-01

    .... In this study, a simulated engine test rig has been established to evaluate thermal fatigue behavior of a candidate engine combustor material, Haynes 188, under superimposed CO2 laser surface impulsive thermal loads (30 to 100 Hz...

  16. High-frequency combustion instability control through acoustic modulation at the inlet boundary for liquid rocket engine applications

    Science.gov (United States)

    Bennewitz, John William

    model-predicted mode stability transition was consistent with experimental observations, supporting the premise that inlet acoustic modulation is a means to control high-frequency combustion instabilities. From the modal analysis, it may be deduced that the inlet impedance provides a damping mechanism for instability suppression. Combined, this work demonstrates the strategic application of acoustic modulation within an injector as a potential method to control high-frequency combustion instabilities for liquid rocket engine applications.

  17. Ceramic thermal barrier coatings for electric utility gas turbine engines

    Science.gov (United States)

    Miller, R. A.

    1986-01-01

    Research and development into thermal barrier coatings for electric utility gas turbine engines is reviewed critically. The type of coating systems developed for aircraft applications are found to be preferred for clear fuel electric utility applications. These coating systems consists of a layer of plasma sprayed zirconia-yttria ceramic over a layer of MCrAly bond coat. They are not recommended for use when molten salts are presented. Efforts to understand coating degradation in dirty environments and to develop corrosion resistant thermal barrier coatings are discussed.

  18. Simple-II: A new numerical thermal model for predicting thermal performance of Stirling engines

    International Nuclear Information System (INIS)

    Babaelahi, Mojtaba; Sayyaadi, Hoseyn

    2014-01-01

    A new thermal model called Simple-II was presented based on modification of the original Simple analysis. First, the engine was modeled considering adiabatic expansion and compression spaces, in which effect of gas leakage from cylinder to buffer space and shuttle effect of displacer were implemented in the basic differential equations. Moreover, non-ideal thermal operation of the regenerator and the longitudinal heat conduction between heater and cooler through the regenerator wall were considered. Based on the magnitudes of pressure drops in heat exchangers, values of pressure in the expansion and compression spaces were corrected. Furthermore, based on the theory of finite speed thermodynamics (FST), the corresponding power loss due to the piston motion and also the mechanical friction were considered. Simple-II was employed for thermal simulation of a prototype Stirling engine. Finally, result of the new model was evaluated by comprehensive comparison of experimental results with those of the previous models. The output power and thermal efficiency were predicted with +20.7% and +7.1% errors, respectively. Also, the regenerator was demonstrated to be the main source of power and heat losses; nevertheless, other loss mechanisms have reasonable effects on output power and/or thermal efficiency of Stirling engines. - Highlights: • A new thermal model was presented based on various loss mechanisms. • Shuttle effect and mass leakage were integrated into differential equations. • FST, mechanical friction and longitudinal conduction losses were considered. • A methodology was presented for numerical solution and correcting results based on losses. • The new model predicted thermal performance of engine with higher accuracy

  19. Transient Two-Dimensional Analysis of Side Load in Liquid Rocket Engine Nozzles

    Science.gov (United States)

    Wang, Ten-See

    2004-01-01

    Two-dimensional planar and axisymmetric numerical investigations on the nozzle start-up side load physics were performed. The objective of this study is to develop a computational methodology to identify nozzle side load physics using simplified two-dimensional geometries, in order to come up with a computational strategy to eventually predict the three-dimensional side loads. The computational methodology is based on a multidimensional, finite-volume, viscous, chemically reacting, unstructured-grid, and pressure-based computational fluid dynamics formulation, and a transient inlet condition based on an engine system modeling. The side load physics captured in the low aspect-ratio, two-dimensional planar nozzle include the Coanda effect, afterburning wave, and the associated lip free-shock oscillation. Results of parametric studies indicate that equivalence ratio, combustion and ramp rate affect the side load physics. The side load physics inferred in the high aspect-ratio, axisymmetric nozzle study include the afterburning wave; transition from free-shock to restricted-shock separation, reverting back to free-shock separation, and transforming to restricted-shock separation again; and lip restricted-shock oscillation. The Mach disk loci and wall pressure history studies reconfirm that combustion and the associated thermodynamic properties affect the formation and duration of the asymmetric flow.

  20. A Framework for Assessing the Reusability of Hardware (Reusable Rocket Engines)

    Science.gov (United States)

    Childress-Thompson, Rhonda; Thomas, Dale; Farrington, Phillip

    2016-01-01

    Within the space flight community, reusability has taken center stage as the new buzzword. In order for reusable hardware to be competitive with its expendable counterpart, two major elements must be closely scrutinized. First, recovery and refurbishment costs must be lower than the development and acquisition costs. Additionally, the reliability for reused hardware must remain the same (or nearly the same) as "first use" hardware. Therefore, it is imperative that a systematic approach be established to enhance the development of reusable systems. However, before the decision can be made on whether it is more beneficial to reuse hardware or to replace it, the parameters that are needed to deem hardware worthy of reuse must be identified. For reusable hardware to be successful, the factors that must be considered are reliability (integrity, life, number of uses), operability (maintenance, accessibility), and cost (procurement, retrieval, refurbishment). These three factors are essential to the successful implementation of reusability while enabling the ability to meet performance goals. Past and present strategies and attempts at reuse within the space industry will be examined to identify important attributes of reusability that can be used to evaluate hardware when contemplating reusable versus expendable options. This paper will examine why reuse must be stated as an initial requirement rather than included as an afterthought in the final design. Late in the process, changes in the overall objective/purpose of components typically have adverse effects that potentially negate the benefits. A methodology for assessing the viability of reusing hardware will be presented by using the Space Shuttle Main Engine (SSME) to validate the approach. Because reliability, operability, and costs are key drivers in making this critical decision, they will be used to assess requirements for reuse as applied to components of the SSME.

  1. High thermal conductivity in electrostatically engineered amorphous polymers

    Science.gov (United States)

    Shanker, Apoorv; Li, Chen; Kim, Gun-Ho; Gidley, David; Pipe, Kevin P.; Kim, Jinsang

    2017-01-01

    High thermal conductivity is critical for many applications of polymers (for example, packaging of light-emitting diodes), in which heat must be dissipated efficiently to maintain the functionality and reliability of a system. Whereas uniaxially extended chain morphology has been shown to significantly enhance thermal conductivity in individual polymer chains and fibers, bulk polymers with coiled and entangled chains have low thermal conductivities (0.1 to 0.4 W m−1 K−1). We demonstrate that systematic ionization of a weak anionic polyelectrolyte, polyacrylic acid (PAA), resulting in extended and stiffened polymer chains with superior packing, can significantly enhance its thermal conductivity. Cross-plane thermal conductivity in spin-cast amorphous films steadily grows with PAA degree of ionization, reaching up to ~1.2 W m−1 K−1, which is on par with that of glass and about six times higher than that of most amorphous polymers, suggesting a new unexplored molecular engineering strategy to achieve high thermal conductivities in amorphous bulk polymers. PMID:28782022

  2. Thermal engineering studies with Excel, Mathcad and Internet

    CERN Document Server

    2016-01-01

    This book provides the fundamentals of the application of mathematical methods, modern computational tools (Excel, Mathcad, SMath, etc.), and the Internet to solve the typical problems of heat and mass transfer, thermodynamics, fluid dynamics, energy conservation and energy efficiency. Chapters cover the technology for creating and using databases on various properties of working fluids, coolants and thermal materials. All calculation methods are provided with links to online computational pages where data can be inserted and recalculated. It discusses tasks involving the generation of electricity at thermal, nuclear, gas turbine and combined-cycle power plants, as well as processes of co- and trigeneration, conditioning facilities and heat pumps. This text engages students and researchers by using modern calculation tools and the Internet for thermal engineering applications. .

  3. Analysis of Thermal Radiation Effects on Temperatures in Turbine Engine Thermal Barrier Coatings

    Science.gov (United States)

    Siegel, Robert; Spuckler, Charles M.

    1998-01-01

    Thermal barrier coatings are important, and in some instances a necessity, for high temperature applications such as combustor liners, and turbine vanes and rotating blades for current and advanced turbine engines. Some of the insulating materials used for coatings, such as zirconia that currently has widespread use, are partially transparent to thermal radiation. A translucent coating permits energy to be transported internally by radiation, thereby increasing the total energy transfer and acting like an increase in thermal conductivity. This degrades the insulating ability of the coating. Because of the strong dependence of radiant emission on temperature, internal radiative transfer effects are increased as temperatures are raised. Hence evaluating the significance of internal radiation is of importance as temperatures are increased to obtain higher efficiencies in advanced engines.

  4. Some typical solid propellant rocket motors

    NARCIS (Netherlands)

    Zandbergen, B.T.C.

    2013-01-01

    Typical Solid Propellant Rocket Motors (shortly referred to as Solid Rocket Motors; SRM's) are described with the purpose to form a database, which allows for comparative analysis and applications in practical SRM engineering.

  5. Scaled Rocket Testing in Hypersonic Flow

    Science.gov (United States)

    Dufrene, Aaron; MacLean, Matthew; Carr, Zakary; Parker, Ron; Holden, Michael; Mehta, Manish

    2015-01-01

    NASA's Space Launch System (SLS) uses four clustered liquid rocket engines along with two solid rocket boosters. The interaction between all six rocket exhaust plumes will produce a complex and severe thermal environment in the base of the vehicle. This work focuses on a recent 2% scale, hot-fire SLS base heating test. These base heating tests are short-duration tests executed with chamber pressures near the full-scale values with gaseous hydrogen/oxygen engines and RSRMV analogous solid propellant motors. The LENS II shock tunnel/Ludwieg tube tunnel was used at or near flight duplicated conditions up to Mach 5. Model development was strongly based on the Space Shuttle base heating tests with several improvements including doubling of the maximum chamber pressures and duplication of freestream conditions. Detailed base heating results are outside of the scope of the current work, rather test methodology and techniques are presented along with broader applicability toward scaled rocket testing in supersonic and hypersonic flow.

  6. Development and Hot-fire Testing of Additively Manufactured Copper Combustion Chambers for Liquid Rocket Engine Applications

    Science.gov (United States)

    Gradl, Paul R.; Greene, Sandy Elam; Protz, Christopher S.; Ellis, David L.; Lerch, Bradley A.; Locci, Ivan E.

    2017-01-01

    NASA and industry partners are working towards fabrication process development to reduce costs and schedules associated with manufacturing liquid rocket engine components with the goal of reducing overall mission costs. One such technique being evaluated is powder-bed fusion or selective laser melting (SLM), commonly referred to as additive manufacturing (AM). The NASA Low Cost Upper Stage Propulsion (LCUSP) program was designed to develop processes and material characterization for GRCop-84 (a NASA Glenn Research Center-developed copper, chrome, niobium alloy) commensurate with powder-bed AM, evaluate bimetallic deposition, and complete testing of a full scale combustion chamber. As part of this development, the process has been transferred to industry partners to enable a long-term supply chain of monolithic copper combustion chambers. To advance the processes further and allow for optimization with multiple materials, NASA is also investigating the feasibility of bimetallic AM chambers. In addition to the LCUSP program, NASA has completed a series of development programs and hot-fire tests to demonstrate SLM GRCop-84 and other AM techniques. NASA's efforts include a 4K lbf thrust liquid oxygen/methane (LOX/CH4) combustion chamber and subscale thrust chambers for 1.2K lbf LOX/hydrogen (H2) applications that have been designed and fabricated with SLM GRCop-84. The same technologies for these lower thrust applications are being applied to 25-35K lbf main combustion chamber (MCC) designs. This paper describes the design, development, manufacturing and testing of these numerous combustion chambers, and the associated lessons learned throughout their design and development processes.

  7. Oxygen Containment System Options for Nuclear Thermal Propulsion Testing

    Data.gov (United States)

    National Aeronautics and Space Administration — All nuclear thermal propulsion (NTP) ground testing conducted in the 1950s and 1960s during the ROVER/(Nuclear Engine Rocket Vehicle Application (NERVA) program...

  8. Easier Analysis With Rocket Science

    Science.gov (United States)

    2003-01-01

    Analyzing rocket engines is one of Marshall Space Flight Center's specialties. When Marshall engineers lacked a software program flexible enough to meet their needs for analyzing rocket engine fluid flow, they overcame the challenge by inventing the Generalized Fluid System Simulation Program (GFSSP), which was named the co-winner of the NASA Software of the Year award in 2001. This paper describes the GFSSP in a wide variety of applications

  9. Thermal barrier coatings for thermal insulation and corrosion resistance in industrial gas turbine engines

    Science.gov (United States)

    Vogan, J. W.; Hsu, L.; Stetson, A. R.

    1981-01-01

    Four thermal barrier coatings were subjected to a 500-hour gas turbine engine test. The coatings were two yttria stabilized zirconias, calcium ortho silicate and calcium meta titanate. The calcium silicate coating exhibited significant spalling. Yttria stabilized zirconia and calcium titanate coatings showed little degradation except in blade leading edge areas. Post-test examination showed variations in the coating due to manual application techniques. Improved process control is required if engineering quality coatings are to be developed. The results indicate that some leading edge loss of the coating can be expected near the tip.

  10. Certain aspects of the environmental impact of nuclear power engineering and thermal power engineering

    International Nuclear Information System (INIS)

    Malenchenko, A.F.

    1979-01-01

    A review is made of the both environmental impact and hazard to man resulting from nuclear power engineering comparing with those of thermal power engineering. At present, in addition to such criteria, as physical-chemical characteristic of energy sources, their efficiency and accessibility for exploitation, new requirements were substantiated in relation to safety of their utilization for environment. So, one of essential problems of nuclear power engineering development consists in assessment and prediction of radioecological consequence. The analysis and operating experience of more than 1000 reactor/years with no accidents and harm for pupulation show, that in respect to impact on environment and man nuclear power engineering is much more safe in comparison with energy sources using tradidional fossile fuel

  11. Engineering-Based Thermal CFD Simulations on Massive Parallel Systems

    KAUST Repository

    Frisch, Jérôme

    2015-05-22

    The development of parallel Computational Fluid Dynamics (CFD) codes is a challenging task that entails efficient parallelization concepts and strategies in order to achieve good scalability values when running those codes on modern supercomputers with several thousands to millions of cores. In this paper, we present a hierarchical data structure for massive parallel computations that supports the coupling of a Navier–Stokes-based fluid flow code with the Boussinesq approximation in order to address complex thermal scenarios for energy-related assessments. The newly designed data structure is specifically designed with the idea of interactive data exploration and visualization during runtime of the simulation code; a major shortcoming of traditional high-performance computing (HPC) simulation codes. We further show and discuss speed-up values obtained on one of Germany’s top-ranked supercomputers with up to 140,000 processes and present simulation results for different engineering-based thermal problems.

  12. Thermal engineering and micro-technology; Thermique et microtechnologie

    Energy Technology Data Exchange (ETDEWEB)

    Kandlikar, S. [Rochester Inst. of Tech., NY (United States); Luo, L. [Institut National Polytechnique, 54 - Nancy (France); Gruss, A. [CEA Grenoble, GRETH, 38 (France); Wautelet, M. [Mons Univ. (Belgium); Gidon, S. [CEA Grenoble, Lab. d' Electronique et de Technologie de l' Informatique (LETI), 38 (France); Gillot, C. [Ecole Nationale Superieure d' Ingenieurs Electriciens de Grenoble, 38 - Saint Martin d' Heres (France)]|[CEA Grenoble, Lab. Electronique et de Technologie de l' Informatique (LETI), 38 (France); Therme, J.; Marvillet, Ch.; Vidil, R. [CEA Grenoble, 38 (France); Dutartre, D. [ST Microelectronique, France (France); Lefebvre, Ph. [SNECMA, 75 - Paris (France); Lallemand, M. [Institut National des Sciences Appliquees (INSA), 69 - Villeurbanne (France); Colin, S. [Institut National des Sciences Appliquees (INSA), 31 - Toulouse (France); Joulin, K. [Ecole Nationale Superieure de Mecanique et d' Aerotechnique (ENSMA), 86 - Poitiers (France); Gad el Hak, M. [Virginia Univ., Charlottesville, VA (United States)

    2003-07-01

    This document gathers the abstracts and transparencies of 5 invited conferences of this congress of the SFT about heat transfers and micro-technologies: Flow boiling in microchannels: non-dimensional groups and heat transfer mechanisms (S. Kandlikar); Intensification and multi-scale process units (L. Luo and A. Gruss); Macro-, micro- and nano-systems: different physics? (M. Wautelet); micro-heat pipes (M. Lallemand); liquid and gas flows inside micro-ducts (S. Colin). The abstracts of the following presentations are also included: Electro-thermal writing of nano-scale memory points in a phase change material (S. Gidon); micro-technologies for cooling in micro-electronics (C. Gillot); the Minatec project (J. Therme); importance and trends of thermal engineering in micro-electronics (D. Dutartre); Radiant heat transfers at short length scales (K. Joulain); Momentum and heat transfer in micro-electromechanical systems (M. Gad-el-Hak). (J.S.)

  13. Graphics tablet technology in second year thermal engineering teaching

    Directory of Open Access Journals (Sweden)

    Antonio Carrillo Andrés

    2013-12-01

    Full Text Available Graphics tablet technology is well known in markets such as manufacturing, graphics arts and design but they have not yet found widespread acceptance for university teaching. A graphics tablet is an affordable and efficient teaching tool that combines the best features from traditional and new media. It allows developing a progressive, interactive lecture (as a traditional blackboard does. However, the tablet is more versatile, being able to integrate graphic material such as tables, graphs, colours, etc. In addition to that, lecture notes can be saved and posted on a course website. The objective of this paper is to show the usefulness of tablet technology in undergraduate engineering teaching by sharing experiences made using a graphics tablet for lecturing a second year Thermal Engineering course. Students’ feedback is definitely positive, though there are some caveats regarding technical and operative problems.

  14. The dynamics of a gas-dust cloud expansion in the upper atmosphere at a shutdown of solid-propellant rocket engines

    Science.gov (United States)

    Nikolaishvili, S. Sh.; Platov, Yu. V.; Chernouss, S. A.

    2015-09-01

    The velocity of spherical gas-dust cloud expansion in the situation when the stages of solid-propellant rocket separate in the upper atmosphere have been determined. The measured velocity vary from 2.5 to 7.5 km/s. The dispersed component accelerates at the front of a shock that develops at engine-thrust shutdown. The model calculations of the gas-dust cloud luminosity intensity qualitatively coincide with the photometric profiles of object images. Such formations can vary from almost homogeneous ball-shaped clouds to rather thin spherical shells depending on the gas-dust cloud mass and the matter distribution within this cloud.

  15. Rotating and positive-displacement pumps for low-thrust rocket engines. Volume 1: Pump Evaluation and design. [of centrifugal pumps

    Science.gov (United States)

    Macgregor, C.; Csomor, A.

    1974-01-01

    Rotating and positive displacement pumps of various types were studied for pumping liquid fluorine for low-thrust, high-performance rocket engines. Included in the analysis were: centrifugal, pitot, Barske, Tesla, drag, gear, vane, axial piston, radial piston, diaphragm, and helirotor pump concepts. The centrifugal pump and the gear pump were selected and these were carried through detailed design and fabrication. Mechanical difficulties were encountered with the gear pump during the preliminary tests in Freon-12. Further testing and development was therefore limited to the centrifugal pump. Tests on the centrifugal pump were conducted in Freon-12 to determine the hydrodynamic performance and in liquid fluorine to demonstrate chemical compatibility.

  16. Grain boundary engineering to enhance thermal stability of electrodeposited nickel

    DEFF Research Database (Denmark)

    Alimadadi, Hossein

    Manufacturing technologies such as injection molding and micro electromechanical systems demand materials with improved mechanical properties (e.g. hardness, ductility) and high durability at elevated temperatures. Significant improvement in some of the mechanical properties is obtained by miniat......Manufacturing technologies such as injection molding and micro electromechanical systems demand materials with improved mechanical properties (e.g. hardness, ductility) and high durability at elevated temperatures. Significant improvement in some of the mechanical properties is obtained...... by miniaturization of the grains down to nano-meter scale. However, this augments the total grain boundary energy stored in the material, hence, making the material less thermally stable. Coherent twin boundaries are of very low energy and mobility compared to all other boundaries in a FCC material. Accordingly......, grain boundary engineering of electrodeposited nickel to achieve high population of coherent twin boundaries and, hence, higher thermal stability is a promising method to achieve simultaneous improvement in mechanical properties and thermal stability. This is of particular scientific and practical...

  17. Study on the special vision sensor for detecting position error in robot precise TIG welding of some key part of rocket engine

    Science.gov (United States)

    Zhang, Wenzeng; Chen, Nian; Wang, Bin; Cao, Yipeng

    2005-01-01

    Rocket engine is a hard-core part of aerospace transportation and thrusting system, whose research and development is very important in national defense, aviation and aerospace. A novel vision sensor is developed, which can be used for error detecting in arc length control and seam tracking in precise pulse TIG welding of the extending part of the rocket engine jet tube. The vision sensor has many advantages, such as imaging with high quality, compactness and multiple functions. The optics design, mechanism design and circuit design of the vision sensor have been described in detail. Utilizing the mirror imaging of Tungsten electrode in the weld pool, a novel method is proposed to detect the arc length and seam tracking error of Tungsten electrode to the center line of joint seam from a single weld image. A calculating model of the method is proposed according to the relation of the Tungsten electrode, weld pool, the mirror of Tungsten electrode in weld pool and joint seam. The new methodologies are given to detect the arc length and seam tracking error. Through analyzing the results of the experiments, a system error modifying method based on a linear function is developed to improve the detecting precise of arc length and seam tracking error. Experimental results show that the final precision of the system reaches 0.1 mm in detecting the arc length and the seam tracking error of Tungsten electrode to the center line of joint seam.

  18. Space shuttle SRM plume expansion sensitivity analysis. [flow characteristics of exhaust gases from solid propellant rocket engines

    Science.gov (United States)

    Smith, S. D.; Tevepaugh, J. A.; Penny, M. M.

    1975-01-01

    The exhaust plumes of the space shuttle solid rocket motors can have a significant effect on the base pressure and base drag of the shuttle vehicle. A parametric analysis was conducted to assess the sensitivity of the initial plume expansion angle of analytical solid rocket motor flow fields to various analytical input parameters and operating conditions. The results of the analysis are presented and conclusions reached regarding the sensitivity of the initial plume expansion angle to each parameter investigated. Operating conditions parametrically varied were chamber pressure, nozzle inlet angle, nozzle throat radius of curvature ratio and propellant particle loading. Empirical particle parameters investigated were mean size, local drag coefficient and local heat transfer coefficient. Sensitivity of the initial plume expansion angle to gas thermochemistry model and local drag coefficient model assumptions were determined.

  19. Thermal energy storage for the Stirling engine powered automobile

    Science.gov (United States)

    Morgan, D. T. (Editor)

    1979-01-01

    A thermal energy storage (TES) system developed for use with the Stirling engine as an automotive power system has gravimetric and volumetric storage densities which are competitive with electric battery storage systems, meets all operational requirements for a practical vehicle, and can be packaged in compact sized automobiles with minimum impact on passenger and freight volume. The TES/Stirling system is the only storage approach for direct use of combustion heat from fuel sources not suitable for direct transport and use on the vehicle. The particular concept described is also useful for a dual mode TES/liquid fuel system in which the TES (recharged from an external energy source) is used for short duration trips (approximately 10 miles or less) and liquid fuel carried on board the vehicle used for long duration trips. The dual mode approach offers the potential of 50 percent savings in the consumption of premium liquid fuels for automotive propulsion in the United States.

  20. Rocket noise - A review

    Science.gov (United States)

    McInerny, S. A.

    1990-10-01

    This paper reviews what is known about far-field rocket noise from the controlled studies of the late 1950s and 1960s and from launch data. The peak dimensionless frequency, the dependence of overall sound power on exhaust parameters, and the directivity of the overall sound power of rockets are compared to those of subsonic jets and turbo-jets. The location of the dominant sound source in the rocket exhaust plume and the mean flow velocity in this region are discussed and shown to provide a qualitative explanation for the low peak Strouhal number, fD(e)/V(e), and large angle of maximum directivity. Lastly, two empirical prediction methods are compared with data from launches of a Titan family vehicle (two, solid rocket motors of 5.7 x 10 to the 6th N thrust each) and the Saturn V (five, liquid oxygen/rocket propellant engines of 6.7 x 10 to the 6th N thrust, each). The agreement is favorable. In contrast, these methods appear to overpredict the far-field sound pressure levels generated by the Space Shuttle.

  1. Osmotic heat engine using thermally responsive ionic liquids

    KAUST Repository

    Zhong, Yujiang

    2017-07-11

    The osmotic heat engine (OHE) is a promising technology for converting low grade heat to electricity. Most of the existing studies have focused on thermolytic salt systems. Herein, for the first time, we proposed to use thermally responsive ionic liquids (TRIL) that have either an upper critical solution temperature (UCST) or lower critical solution temperature (LCST) type of phase behavior as novel thermolytic osmotic agents. Closed-loop TRIL-OHEs were designed based on these unique phase behaviors to convert low grade heat to work or electricity. Experimental studies using two UCST-type TRILs, protonated betaine bis(trifluoromethyl sulfonyl)imide ([Hbet][Tf2N]) and choline bis(trifluoromethylsulfonyl)imide ([Choline][Tf2N]) showed that (1) the specific energy of the TRIL-OHE system could reach as high as 4.0 times that of the seawater and river water system, (2) the power density measured from a commercial FO membrane reached up to 2.3 W/m2, and (3) the overall energy efficiency reached up to 2.6% or 18% of the Carnot efficiency at no heat recovery and up to 10.5% or 71% of the Carnet efficiency at 70% heat recovery. All of these results clearly demonstrated the great potential of using TRILs as novel osmotic agents to design high efficient OHEs for recovery of low grade thermal energy to work or electricity.

  2. Osmotic Heat Engine Using Thermally Responsive Ionic Liquids.

    Science.gov (United States)

    Zhong, Yujiang; Wang, Xinbo; Feng, Xiaoshuang; Telalovic, Selvedin; Gnanou, Yves; Huang, Kuo-Wei; Hu, Xiao; Lai, Zhiping

    2017-08-15

    The osmotic heat engine (OHE) is a promising technology for converting low grade heat to electricity. Most of the existing studies have focused on thermolytic salt systems. Herein, for the first time, we proposed to use thermally responsive ionic liquids (TRIL) that have either an upper critical solution temperature (UCST) or lower critical solution temperature (LCST) type of phase behavior as novel thermolytic osmotic agents. Closed-loop TRIL-OHEs were designed based on these unique phase behaviors to convert low grade heat to work or electricity. Experimental studies using two UCST-type TRILs, protonated betaine bis(trifluoromethyl sulfonyl)imide ([Hbet][Tf 2 N]) and choline bis(trifluoromethylsulfonyl)imide ([choline][Tf 2 N]) showed that (1) the specific energy of the TRIL-OHE system could reach as high as 4.0 times that of the seawater and river water system, (2) the power density measured from a commercial FO membrane reached up to 2.3 W/m 2 , and (3) the overall energy efficiency reached up to 2.6% or 18% of the Carnot efficiency at no heat recovery and up to 10.5% or 71% of the Carnet efficiency at 70% heat recovery. All of these results clearly demonstrated the great potential of using TRILs as novel osmotic agents to design high efficient OHEs for recovery of low grade thermal energy to work or electricity.

  3. Thermal Fault Tolerance Analysis of Carbon Fiber Rope Barrier Systems for Use in the Reusable Solid Rocket Motor ( RSRM) Nozzle Joints

    Science.gov (United States)

    Clayton, J. Louie; Phelps, Lisa (Technical Monitor)

    2001-01-01

    Carbon Fiber Rope (CFR) thermal barrier systems are being considered for use in several RSRM (Reusable Solid Rocket Motor) nozzle joints as a replacement for the current assembly gap close-out process/design. This study provides for development and test verification of analysis methods used for flow-thermal modeling of a CFR thermal barrier subject to fault conditions such as rope combustion gas blow-by and CFR splice failure. Global model development is based on a 1-D (one dimensional) transient volume filling approach where the flow conditions are calculated as a function of internal 'pipe' and porous media 'Darcy' flow correlations. Combustion gas flow rates are calculated for the CFR on a per-linear inch basis and solved simultaneously with a detailed thermal-gas dynamic model of a local region of gas blow by (or splice fault). Effects of gas compressibility, friction and heat transfer are accounted for the model. Computational Fluid Dynamic (CFD) solutions of the fault regions are used to characterize the local flow field, quantify the amount of free jet spreading and assist in the determination of impingement film coefficients on the nozzle housings. Gas to wall heat transfer is simulated by a large thermal finite element grid of the local structure. The employed numerical technique loosely couples the FE (Finite Element) solution with the gas dynamics solution of the faulted region. All free constants that appear in the governing equations are calibrated by hot fire sub-scale test. The calibrated model is used to make flight predictions using motor aft end environments and timelines. Model results indicate that CFR barrier systems provide a near 'vented joint' style of pressurization. Hypothetical fault conditions considered in this study (blow by, splice defect) are relatively benign in terms of overall heating to nozzle metal housing structures.

  4. Method of operating a thermal engine powered by a chemical reaction

    Science.gov (United States)

    Ross, John; Escher, Claus

    1988-01-01

    The invention involves a novel method of increasing the efficiency of a thermal engine. Heat is generated by a non-linear chemical reaction of reactants, said heat being transferred to a thermal engine such as Rankine cycle power plant. The novel method includes externally perturbing one or more of the thermodynamic variables of said non-linear chemical reaction.

  5. Application of Probabilistic Methods to Assess Risk Due to Resonance in the Design of J-2X Rocket Engine Turbine Blades

    Science.gov (United States)

    Brown, Andrew M.; DeHaye, Michael; DeLessio, Steven

    2011-01-01

    The LOX-Hydrogen J-2X Rocket Engine, which is proposed for use as an upper-stage engine for numerous earth-to-orbit and heavy lift launch vehicle architectures, is presently in the design phase and will move shortly to the initial development test phase. Analysis of the design has revealed numerous potential resonance issues with hardware in the turbomachinery turbine-side flow-path. The analysis of the fuel pump turbine blades requires particular care because resonant failure of the blades, which are rotating in excess of 30,000 revolutions/minutes (RPM), could be catastrophic for the engine and the entire launch vehicle. This paper describes a series of probabilistic analyses performed to assess the risk of failure of the turbine blades due to resonant vibration during past and present test series. Some significant results are that the probability of failure during a single complete engine hot-fire test is low (1%) because of the small likelihood of resonance, but that the probability increases to around 30% for a more focused turbomachinery-only test because all speeds will be ramped through and there is a greater likelihood of dwelling at more speeds. These risk calculations have been invaluable for use by program management in deciding if risk-reduction methods such as dampers are necessary immediately or if the test can be performed before the risk-reduction hardware is ready.

  6. Experimental investigation of solid rocket motors for small sounding rockets

    Science.gov (United States)

    Suksila, Thada

    2018-01-01

    Experimentation and research of solid rocket motors are important subjects for aerospace engineering students. However, many institutes in Thailand rarely include experiments on solid rocket motors in research projects of aerospace engineering students, mainly because of the complexity of mixing the explosive propellants. This paper focuses on the design and construction of a solid rocket motor for total impulse in the class I-J that can be utilised as a small sounding rocket by researchers in the near future. Initially, the test stands intended for measuring the pressure in the combustion chamber and the thrust of the solid rocket motor were designed and constructed. The basic design of the propellant configuration was evaluated. Several formulas and ratios of solid propellants were compared for achieving the maximum thrust. The convenience of manufacturing and casting of the fabricated solid rocket motors were a critical consideration. The motor structural analysis such as the combustion chamber wall thickness was also discussed. Several types of nozzles were compared and evaluated for ensuring the maximum thrust of the solid rocket motors during the experiments. The theory of heat transfer analysis in the combustion chamber was discussed and compared with the experimental data.

  7. A Rocket Powered Single-Stage-to-Orbit Launch Vehicle With U.S. and Soviet Engineers

    Science.gov (United States)

    MacConochie, Ian O.; Stnaley, Douglas O.

    1991-01-01

    A single-stage-to-orbit launch vehicle is used to assess the applicability of Soviet Energia high-pressure-hydrocarbon engine to advanced U.S. manned space transportation systems. Two of the Soviet engines are used with three Space Shuttle Main Engines. When applied to a baseline vehicle that utilized advanced hydrocarbon engines, the higher weight of the Soviet engines resulted in a 20 percent loss of payload capability and necessitated a change in the crew compartment size and location from mid-body to forebody in order to balance the vehicle. Various combinations of Soviet and Shuttle engines were evaluated for comparison purposes, including an all hydrogen system using all Space Shuttle Main Engines. Operational aspects of the baseline vehicle are also discussed. A new mass properties program entitles Weights and Moments of Inertia (WAMI) is used in the study.

  8. History of Thermal Barrier Coatings for Gas Turbine Engines: Emphasizing NASA's Role from 1942 to 1990

    Science.gov (United States)

    Miller, Robert A.

    2009-01-01

    NASA has played a central role in the development of thermal barrier coatings (TBCs) for gas turbine applications. This report discusses the history of TBCs emphasizing the role NASA has played beginning with (1) frit coatings in the 1940s and 1950s; (2) thermally sprayed coatings for rocket application in the 1960s and early 1970s; (3) the beginnings of the modern era of turbine section coatings in the mid 1970s; and (4) failure mechanism and life prediction studies in the 1980s and 1990s. More recent efforts are also briefly discussed.

  9. Incorporating engineering intuition for parameter estimation in thermal sciences

    Science.gov (United States)

    Balaji, C.; Reddy, B. Konda; Herwig, H.

    2013-12-01

    This paper proposes a new method of incorporating priors based on engineering intuition for solving inverse problems. The thesis of this paper is that if an asymptote can be found to a problem in applied sciences or engineering, estimation of parameters can be first done for this asymptotic variant, which in principle should be simpler, since one or more parameters of the original problem may vanish for the asymptotic variant. Even so, by solving the inverse problem associated with the asymptotic variant, estimates of key parameters of the full problem can be obtained. This information can then be quantitatively incorporated as priors in the estimation of parameters for the full version of the problem which we call as prior generation through asymptotic variant. The goal is to see if this methodology will significantly reduce the uncertainties in the resulting estimates. To demonstrate this methodology, the classic problem of unsteady heat transfer from a one dimensional fin is chosen. The inverse problem is posed as the simultaneous estimation of the temperature dependent transfer coefficient (h θ ) and the thermal diffusivity ( α) of the fin material, given experimentally measured temperature-time histories at various locations along the fin. The asymptotic variant θ ( x, t) is the steady state problem where the influence of thermal diffusivity vanishes. Using surrogate temperature data generated from assumed values of h θ , first the asymptotic variant of the problem is solved using the Markov Chain Monte Carlo method in a Bayesian framework to generate an estimate of h θ . The estimate of h θ is then used as an informative prior for solving the inverse problem of determining h θ and α from θ ( x, t), and the effect of prior is quantitatively assessed by performing estimation with and without the prior. Finally, for purposes of validation, in-house experiments have been done where θ ( x, t) is generated using liquid crystal thermography and these data

  10. CECE: Expanding the Envelope of Deep Throttling Technology in Liquid Oxygen/Liquid Hydrogen Rocket Engines for NASA Exploration Missions

    Science.gov (United States)

    Giuliano, Victor J.; Leonard, Timothy G.; Lyda, Randy T.; Kim, Tony S.

    2010-01-01

    As one of the first technology development programs awarded by NASA under the Vision for Space Exploration, the Pratt & Whitney Rocketdyne (PWR) Deep Throttling, Common Extensible Cryogenic Engine (CECE) program was selected by NASA in November 2004 to begin technology development and demonstration toward a deep throttling, cryogenic engine supporting ongoing trade studies for NASA s Lunar Lander descent stage. The CECE program leverages the maturity and previous investment of a flight-proven hydrogen/oxygen expander cycle engine, the PWR RL10, to develop and demonstrate an unprecedented combination of reliability, safety, durability, throttlability, and restart capabilities in high-energy, cryogenic, in-space propulsion. The testbed selected for the deep throttling demonstration phases of this program was a minimally modified RL10 engine, allowing for maximum current production engine commonality and extensibility with minimum program cost. Four series of demonstrator engine tests have been successfully completed between April 2006 and April 2010, accumulating 7,436 seconds of hot fire time over 47 separate tests. While the first two test series explored low power combustion (chug) and system instabilities, the third test series investigated and was ultimately successful in demonstrating several mitigating technologies for these instabilities and achieved a stable throttling ratio of 13:1. The fourth test series significantly expanded the engine s operability envelope by successfully demonstrating a closed-loop control system and extensive transient modeling to enable lower power engine starting, faster throttle ramp rates, and mission-specific ignition testing. The final hot fire test demonstrated a chug-free, minimum power level of 5.9%, corresponding to an overall 17.6:1 throttling ratio achieved. In total, these tests have provided an early technology demonstration of an enabling cryogenic propulsion concept with invaluable system-level technology data

  11. Experimental research of thermal loading of the rocket payload fairing element during the atmospheric phase of the descent trajectory

    Science.gov (United States)

    Trushlyakov, V.; Iordan, Yu; Davydovich, D.; Zharikov, K.; Dron, M.

    2018-01-01

    The thermal loading physical simulation in the experimental wind tunnel on the design element of the payload fairing made of carbon fiber was done. The experimental study is given in the speed range below 70 m/s, which corresponds to the interval of heights of the descent trajectory of the payload fairing half below 10 km. The values of heat transfer coefficient are obtained. The analysis of the results is carried out.

  12. Thermal balance of a LPG fuelled, four stroke SI engine with water addition

    International Nuclear Information System (INIS)

    Ozcan, Hakan; Soeylemez, M.S.

    2006-01-01

    The effect of water injection on a spark ignition engine thermal balance and performance has been experimentally investigated. A four stroke, four cylinder conventional engine was used with LPG (liquid petroleum gas) as fuel. Different water to fuel ratios by mass were used with variable engine speed ranging from 1000 to 4500 rpm. The results showed that as the water injection level to the engine increased, the percentage of useful work increased, while the losses other than unaccounted losses decreased. Additionally, the specific fuel consumption decreases, while the engine thermal efficiency increases. The average increase in the brake thermal efficiency for a 0.5 water to fuel mass ratio is approximately 2.7% over the use of LPG alone for the engine speed range studied

  13. Investigation into the Interactions between thermal management, lubrication and control systems of a diesel engine

    OpenAIRE

    Burke, Richard D

    2011-01-01

    Engine thermal and lubricant systems have only recently been a serious focus in engine design and in general remain under passive control. The introduction of active control has shown benefits in fuel consumption during the engine warm-up period, however there is a lack of rigorous calibration of these devices in conjunction with other engine systems.For these systems, benefits in fuel consumption (FC) are small and accurate measurement systems are required. Analysis of both FC and NOx emissi...

  14. Engineering the thermal conductivity along an individual silicon nanowire by selective helium ion irradiation

    Science.gov (United States)

    Zhao, Yunshan; Liu, Dan; Chen, Jie; Zhu, Liyan; Belianinov, Alex; Ovchinnikova, Olga S.; Unocic, Raymond R.; Burch, Matthew J.; Kim, Songkil; Hao, Hanfang; Pickard, Daniel S.; Li, Baowen; Thong, John T. L.

    2017-06-01

    The ability to engineer the thermal conductivity of materials allows us to control the flow of heat and derive novel functionalities such as thermal rectification, thermal switching and thermal cloaking. While this could be achieved by making use of composites and metamaterials at bulk length-scales, engineering the thermal conductivity at micro- and nano-scale dimensions is considerably more challenging. In this work, we show that the local thermal conductivity along a single Si nanowire can be tuned to a desired value (between crystalline and amorphous limits) with high spatial resolution through selective helium ion irradiation with a well-controlled dose. The underlying mechanism is understood through molecular dynamics simulations and quantitative phonon-defect scattering rate analysis, where the behaviour of thermal conductivity with dose is attributed to the accumulation and agglomeration of scattering centres at lower doses. Beyond a threshold dose, a crystalline-amorphous transition was observed.

  15. Engineering the thermal conductivity along an individual silicon nanowire by selective helium ion irradiation.

    Science.gov (United States)

    Zhao, Yunshan; Liu, Dan; Chen, Jie; Zhu, Liyan; Belianinov, Alex; Ovchinnikova, Olga S; Unocic, Raymond R; Burch, Matthew J; Kim, Songkil; Hao, Hanfang; Pickard, Daniel S; Li, Baowen; Thong, John T L

    2017-06-27

    The ability to engineer the thermal conductivity of materials allows us to control the flow of heat and derive novel functionalities such as thermal rectification, thermal switching and thermal cloaking. While this could be achieved by making use of composites and metamaterials at bulk length-scales, engineering the thermal conductivity at micro- and nano-scale dimensions is considerably more challenging. In this work, we show that the local thermal conductivity along a single Si nanowire can be tuned to a desired value (between crystalline and amorphous limits) with high spatial resolution through selective helium ion irradiation with a well-controlled dose. The underlying mechanism is understood through molecular dynamics simulations and quantitative phonon-defect scattering rate analysis, where the behaviour of thermal conductivity with dose is attributed to the accumulation and agglomeration of scattering centres at lower doses. Beyond a threshold dose, a crystalline-amorphous transition was observed.

  16. Replacement of Chromium Electroplating on Gas Turbine Engine Components Using Thermal Spray Coatings

    National Research Council Canada - National Science Library

    Sartwell, Bruce D; Legg, Keith O; Schell, Jerry; Bondaruk, Bob; Alford, Charles; Natishan, Paul; Lawrence, Steven; Shubert, Gary; Bretz, Philip; Kaltenhauser, Anne

    2005-01-01

    .... This document constitutes the final report on a project to qualify high-velocity oxygen-fuel (HVOF) and plasma thermal spray coatings as a replacement for hard chrome plating on gas turbine engine components...

  17. The Importance of Thermal Heat Bridges in Civil Engineering

    Directory of Open Access Journals (Sweden)

    Adriana Tokar

    2011-10-01

    Full Text Available Based on the heat transfer characteristics of a construction, the expected temperatures along interior surfaces must be evaluated in order to predict (and avoid areas of potential moisture condensation. Beyond preventing damage to building materials caused by mould growth, adequate surface temperatures are also a relevant factor in the thermal comfort of an interior environment. An agreable climate in a room can be obtained, when relative humidity is between 40 and 60%. As the air in a room is warmer, the more vapor can absorb (and vice versa, influencing the thermal comfort index. Heat losses are influenced largely by thermal bridges of construction. The importance of the thermal heat bridges is strongly increasing today. In new developments the thermal optimization of junctions in today common low energy constructions receives very special standing. The subject of avoiding thermal bridges in passive houses became predominant.

  18. True Concurrent Thermal Engineering Integrating CAD Model Building with Finite Element and Finite Difference Methods

    Science.gov (United States)

    Panczak, Tim; Ring, Steve; Welch, Mark

    1999-01-01

    Thermal engineering has long been left out of the concurrent engineering environment dominated by CAD (computer aided design) and FEM (finite element method) software. Current tools attempt to force the thermal design process into an environment primarily created to support structural analysis, which results in inappropriate thermal models. As a result, many thermal engineers either build models "by hand" or use geometric user interfaces that are separate from and have little useful connection, if any, to CAD and FEM systems. This paper describes the development of a new thermal design environment called the Thermal Desktop. This system, while fully integrated into a neutral, low cost CAD system, and which utilizes both FEM and FD methods, does not compromise the needs of the thermal engineer. Rather, the features needed for concurrent thermal analysis are specifically addressed by combining traditional parametric surface based radiation and FD based conduction modeling with CAD and FEM methods. The use of flexible and familiar temperature solvers such as SINDA/FLUINT (Systems Improved Numerical Differencing Analyzer/Fluid Integrator) is retained.

  19. Modeling Potential Carbon Monoxide Exposure Due to Operation of a Major Rocket Engine Altitude Test Facility Using Computational Fluid Dynamics

    Science.gov (United States)

    Blotzer, Michael J.; Woods, Jody L.

    2009-01-01

    This viewgraph presentation reviews computational fluid dynamics as a tool for modelling the dispersion of carbon monoxide at the Stennis Space Center's A3 Test Stand. The contents include: 1) Constellation Program; 2) Constellation Launch Vehicles; 3) J2X Engine; 4) A-3 Test Stand; 5) Chemical Steam Generators; 6) Emission Estimates; 7) Located in Existing Test Complex; 8) Computational Fluid Dynamics; 9) Computational Tools; 10) CO Modeling; 11) CO Model results; and 12) Next steps.

  20. A ``NEW'' Solid-Core Reactor Fuel Form that Maximizes the Performance of Nuclear Thermal and Electric Rockets

    Science.gov (United States)

    Rom, Frank E.; Finnegan, Patrick M.

    1994-07-01

    The ``NEW'' solid-core fuel form is the old Vapor Transport (VT) fuel pin investigated at NASA about 30 years ago. It is simply a tube sealed at both ends partially filled with UO2. During operation the UO2 forms an annular layer on the inside of the tube by vaporization and condensation. This form is an ideal structure for overall strength and retention of fission products. All of the structural material lies between the fuel (including fission products) and the reactor coolant. The isothermal inside fuel surface temperature that results from the vaporization and condensation of fuel during operation eliminates hotspots, significantly increasing the design fuel pin surface temperature. For NTP, W-UO2 fuel pins yield higher operating temperatures than for other fuel forms, because W has about a ten-fold lower vaporization rate compared to any other known material. The use of perigee propulsion using W-UO2 fuel pins can result in a more than ten-fold reduction in reactor power. Lower reactor power, together with zero fission product release potential, and the simplicity of fabrication of VT fuel pins should greatly simplify and reduce the cost of development of NTP. For NEP, VT fuel pins can increase fast neutron spectrum reactor life with no fission product release. Thermal spectrum NEP reactors using W184 or Mo VT fuel pins, with only small amounts of high neutron absorbing additives, offer benefits because of much lower fissionable fuel requirements. The VT fuel pin has application to commercial power reactors with similar benefits.

  1. V-2 Rocket at White Sands

    Science.gov (United States)

    1946-01-01

    A V-2 rocket takes flight at White Sands, New Mexico, in 1946. The German engineers and scientists who developed the V-2 came to the United States at the end of World War II and continued rocket testing under the direction of the U. S. Army, launching more than sixty V-2s.

  2. Rocket + Science = Dialogue

    Science.gov (United States)

    Morris,Bruce; Sullivan, Greg; Burkey, Martin

    2010-01-01

    It's a cliche that rocket engineers and space scientists don t see eye-to-eye. That goes double for rocket engineers working on human spaceflight and scientists working on space telescopes and planetary probes. They work fundamentally different problems but often feel that they are competing for the same pot of money. Put the two groups together for a weekend, and the results could be unscientific or perhaps combustible. Fortunately, that wasn't the case when NASA put heavy lift launch vehicle designers together with astronomers and planetary scientists for two weekend workshops in 2008. The goal was to bring the top people from both groups together to see how the mass and volume capabilities of NASA's Ares V heavy lift launch vehicle could benefit the science community. Ares V is part of NASA's Constellation Program for resuming human exploration beyond low Earth orbit, starting with missions to the Moon. In the current mission scenario, Ares V launches a lunar lander into Earth orbit. A smaller Ares I rocket launches the Orion crew vehicle with up to four astronauts. Orion docks with the lander, attached to the Ares V Earth departure stage. The stage fires its engine to send the mated spacecraft to the Moon. Standing 360 feet high and weighing 7.4 million pounds, NASA's new heavy lifter will be bigger than the 1960s-era Saturn V. It can launch almost 60 percent more payload to translunar insertion together with the Ares I and 35 percent more mass to low Earth orbit than the Saturn V. This super-sized capability is, in short, designed to send more people to more places to do more things than the six Apollo missions.

  3. Thermal balance of a four stroke SI engine operating on hydrogen as a supplementary fuel

    International Nuclear Information System (INIS)

    Yueksel, F.; Ceviz, M.A.

    2003-01-01

    This paper investigates the effects of adding constant quantity hydrogen to gasoline-air mixture on SI engine thermal balance and performance. A four stroke, four-cylinder SI engine was used for conducting this study. Thermal balance tests were conducted for engine thermal efficiency, heat loss through the exhaust gases, heat loss to the cooling water and unaccounted losses (i.e. heat lost by lubricating oil, radiation), while performance tests were in respect to the brake power, specific fuel consumption and air ratio. Hydrogen supplementations were used with three different and fixed mass flow rates; 0.129, 0.168 and 0.208 kg h -1 at near three-fourth throttle opening position and variable engine speed ranging from 1000 to 4500 rpm. The results showed that supplementation of hydrogen to gasoline decreases the heat loss to cooling water and unaccounted losses, and the heat loss through the exhaust gas is nearly the same with pure gasoline experiments. Additionally, specific fuel consumption decreases, while the engine thermal efficiency and the air ratio increase. Engine performance parameters such as thermal efficiency and specific fuel consumption improved the level of the ratio of hydrogen mass flow rate to that of gasoline up to 5%

  4. A Collaborative Analysis Tool for Integrated Hypersonic Aerodynamics, Thermal Protection Systems, and RBCC Engine Performance for Single Stage to Orbit Vehicles

    Science.gov (United States)

    Stanley, Thomas Troy; Alexander, Reginald; Landrum, Brian

    2000-01-01

    Presented is a computer-based tool that connects several disciplines that are needed in the complex and integrated design of high performance reusable single stage to orbit (SSTO) vehicles. Every system is linked to every other system, as is the case of SSTO vehicles with air breathing propulsion, which is currently being studied by NASA. An RBCC propulsion system integrates airbreathing and rocket propulsion into a single engine assembly enclosed within a cowl or duct. A typical RBCC propulsion system operates as a ducted rocket up to approximately Mach 3. Then there is a transition to a ramjet mode for supersonic-to-hypersonic acceleration. Around Mach 8 the engine transitions to a scramjet mode. During the ramjet and scramjet modes, the integral rockets operate as fuel injectors. Around Mach 10-12 (the actual value depends on vehicle and mission requirements), the inlet is physically closed and the engine transitions to an integral rocket mode for orbit insertion. A common feature of RBCC propelled vehicles is the high degree of integration between the propulsion system and airframe. At high speeds the vehicle forebody is fundamentally part of the engine inlet, providing a compression surface for air flowing into the engine. The compressed air is mixed with fuel and burned. The combusted mixture must be expanded to an area larger than the incoming stream to provide thrust. Since a conventional nozzle would be too large, the entire lower after body of the vehicle is used as an expansion surface. Because of the high external temperatures seen during atmospheric flight, the design of an airbreathing SSTO vehicle requires delicate tradeoffs between engine design, vehicle shape, and thermal protection system (TPS) sizing in order to produce an optimum system in terms of weight (and cost) and maximum performance. To adequately determine the performance of the engine/vehicle, the Hypersonic Flight Inlet Model (HYFIM) module was designed to interface with the RBCC

  5. Robustly Engineering Thermal Conductivity of Bilayer Graphene by Interlayer Bonding

    Science.gov (United States)

    Zhang, Xiaoliang; Gao, Yufei; Chen, Yuli; Hu, Ming

    2016-01-01

    Graphene and its bilayer structure are the two-dimensional crystalline form of carbon, whose extraordinary electron mobility and other unique features hold great promise for nanoscale electronics and photonics. Their realistic applications in emerging nanoelectronics usually call for thermal transport manipulation in a controllable and precise manner. In this paper we systematically studied the effect of interlayer covalent bonding, in particular different interlay bonding arrangement, on the thermal conductivity of bilayer graphene using equilibrium molecular dynamics simulations. It is revealed that, the thermal conductivity of randomly bonded bilayer graphene decreases monotonically with the increase of interlayer bonding density, however, for the regularly bonded bilayer graphene structure the thermal conductivity possesses unexpectedly non-monotonic dependence on the interlayer bonding density. The results suggest that the thermal conductivity of bilayer graphene depends not only on the interlayer bonding density, but also on the detailed topological configuration of the interlayer bonding. The underlying mechanism for this abnormal phenomenon is identified by means of phonon spectral energy density, participation ratio and mode weight factor analysis. The large tunability of thermal conductivity of bilayer graphene through rational interlayer bonding arrangement paves the way to achieve other desired properties for potential nanoelectronics applications involving graphene layers. PMID:26911859

  6. Backyard rockets learn to make and launch rockets, missiles, cannons, and other projectiles

    CERN Document Server

    com, Instructables; Warren, Mike

    2014-01-01

    Originating from Instructables, a popular project-based community made up of all sorts of characters with wacky hobbies and a desire to pass on their wisdom to others, Backyard Rockets is made up of projects from a medley of authors who have collected and shared a treasure trove of rocket-launching plans and the knowledge to make their projects soar! Backyard Rockets gives step-by-step instructions, with pictures to guide the way, on how to launch your very own project into the sky. All of these authors have labored over their endeavors to pass their knowledge on and make it easier for others to attempt. Discover how to create the following projects: Teeny, Tiny Rocket Engine Ultimate Straw Rocket Rocket Eggstronaut Pocket Rocket Launcher Iron Man Model Rocket Model Rocket with Camera Rocket-Powered Matchbox Cars – Extreme And much more! The Instructables community has provided a compendium of rocket savvy from innovators who have paved the way for other curious minds. In addition to rockets, fireworks, and ...

  7. Development of a Diesel Engine Thermal Overload Monitoring System with Applications and Test Results

    Directory of Open Access Journals (Sweden)

    Sangram Kishore Nanda

    2017-06-01

    Full Text Available In this research, the development of a diesel engine thermal overload monitoring system is presented with applications and test results. The designed diesel engine thermal overload monitoring system consists of two set of sensors, i.e., a lambda sensor to measure the oxygen concentration and a fast response thermocouple to measure the temperature of the gas leaving the cylinder. A medium speed Ruston diesel engine is instrumented to measure the required engine process parameters, measurements are taken at constant load and variable fuel delivery i.e., normal and excessive injection. It is indicated that with excessive injection, the test engine is of high risk to be operated at thermal overload condition. Further tests were carried out on a Sulzer 7RTA84T engine to explore the influence of engine operating at thermal overload condition on exhaust gas temperature and oxygen concentration in the blow down gas. It is established that a lower oxygen concentration in the blow down gas corresponds to a higher exhaust gas temperature. The piston crown wear rate will then be much higher due to the high rate of heat transfer from a voluminous flame.

  8. Workshop on the applications of new computer tools to thermal engineering; Applications a la thermique des nouveaux outils informatiques

    Energy Technology Data Exchange (ETDEWEB)

    NONE

    1996-12-31

    This workshop on the applications of new computer tools to thermal engineering has been organized by the French society of thermal engineers. Seven papers have been presented, from which two papers dealing with thermal diffusivity measurements in materials and with the optimization of dryers have been selected for ETDE. (J.S.)

  9. Long duration blade loss simulations including thermal growths for dual-rotor gas turbine engine

    Science.gov (United States)

    Sun, Guangyoung; Palazzolo, Alan; Provenza, A.; Lawrence, C.; Carney, K.

    2008-09-01

    This paper presents an approach for blade loss simulation including thermal growth effects for a dual-rotor gas turbine engine supported on bearing and squeeze film damper. A nonlinear ball bearing model using the Hertzian formula predicts ball contact load and stress, while a simple thermal model estimates the thermal growths of bearing components during the blade loss event. The modal truncation augmentation method combined with a proposed staggered integration scheme is verified through simulation results as an efficient tool for analyzing a flexible dual-rotor gas turbine engine dynamics with the localized nonlinearities of the bearing and damper, with the thermal growths and with a flexible casing model. The new integration scheme with enhanced modeling capability reduces the computation time by a factor of 12, while providing a variety of solutions with acceptable accuracy for durations extending over several thermal time constants.

  10. A Hydrogen Containment Process for Nuclear Thermal Engine Ground testing

    Science.gov (United States)

    Wang, Ten-See; Stewart, Eric; Canabal, Francisco

    2016-01-01

    The objective of this study is to propose a new total hydrogen containment process to enable the testing required for NTP engine development. This H2 removal process comprises of two unit operations: an oxygen-rich burner and a shell-and-tube type of heat exchanger. This new process is demonstrated by simulation of the steady state operation of the engine firing at nominal conditions.

  11. Calculation of thermal conductivity for new materials used in intake systems of internal combustion engines

    Science.gov (United States)

    Birtok-Bǎneasǎ, Corneliu; RaÅ£iu, Sorin Aurel; HepuÅ£, Teodor

    2017-07-01

    This paper presents a method for reduce thermal losses in the intake system of an internal combustion engine, whit improvement of airflow and thermal protection. The method consists in insulating the intake with a new kind of material. The present paper focuses on calculation of thermal conductivity for a new material developed by the authors, using the heat flux plate method. This experimental method consists in placing the sample of the new material in a calorimetric chamber and heating from underside. As the heat flux which passes through the sample material is constant and knowing the values of the temperatures for the both sides of sample, the sample material thermal conductivity is determined.

  12. Materials Characterization of Additively Manufactured Components for Rocket Propulsion

    Science.gov (United States)

    Carter, Robert; Draper, Susan; Locci, Ivan; Lerch, Bradley; Ellis, David; Senick, Paul; Meyer, Michael; Free, James; Cooper, Ken; Jones, Zachary

    2015-01-01

    To advance Additive Manufacturing (AM) technologies for production of rocket propulsion components the NASA Glenn Research Center (GRC) is applying state of the art characterization techniques to interrogate microstructure and mechanical properties of AM materials and components at various steps in their processing. The materials being investigated for upper stage rocket engines include titanium, copper, and nickel alloys. Additive manufacturing processes include laser powder bed, electron beam powder bed, and electron beam wire fed processes. Various post build thermal treatments, including Hot Isostatic Pressure (HIP), have been studied to understand their influence on microstructure, mechanical properties, and build density. Micro-computed tomography, electron microscopy, and mechanical testing in relevant temperature environments has been performed to develop relationships between build quality, microstructure, and mechanical performance at temperature. A summary of GRC's Additive Manufacturing roles and experimental findings will be presented.

  13. Material Characterization of Additively Manufactured Components for Rocket Propulsion

    Science.gov (United States)

    Carter, Robert; Draper, Susan; Locci, Ivan; Lerch, Bradley; Ellis, David; Senick, Paul; Meyer, Michael; Free, James; Cooper, Ken; Jones, Zachary

    2015-01-01

    To advance Additive Manufacturing (AM) technologies for production of rocket propulsion components the NASA Glenn Research Center (GRC) is applying state of the art characterization techniques to interrogate microstructure and mechanical properties of AM materials and components at various steps in their processing. The materials being investigated for upper stage rocket engines include titanium, copper, and nickel alloys. Additive manufacturing processes include laser powder bed, electron beam powder bed, and electron beam wire fed processes. Various post build thermal treatments, including Hot Isostatic Pressure (HIP), have been studied to understand their influence on microstructure, mechanical properties, and build density. Micro-computed tomography, electron microscopy, and mechanical testing in relevant temperature environments has been performed to develop relationships between build quality, microstructure, and mechanical performance at temperature. A summary of GRCs Additive Manufacturing roles and experimental findings will be presented.

  14. Thermal barrier coatings for the space shuttle main engine turbine blades

    Science.gov (United States)

    Bhat, B. N.; Gilmore, H. L.; Holmes, R. R.

    1985-01-01

    The Space Shuttle Main Engine (SSME) turbopump turbine blades experience extremely severe thermal shocks during start-up and shut-down. For instance, the high pressure fuel turbopump turbine which burns liquid hydrogen operates at approximately 1500 F, but is shut down fuel rich with turbine blades quenced in liquid hydrogen. This thermal shock is a major contributor to blade cracking. The same thermal shock cause the protective ZrO2 thermal barrier coatings to spall or flake off, leaving only the NiCrAlY bond coating which provides only a minimum thermal protection. The turbine blades are therefore life limited to about 3000 sec for want of a good thermal barrier. A suitable thermal barrier coating (TBC) is being developed for the SSME turbine blades. Various TBCs developed for the gas turbine engines were tested in a specially built turbine blade tester. This tester subjects the coated blades to thermal and pressure cycles similar to those during actual operation of the turbine. The coatings were applied using a plasma spraying techniques both under atmospheric conditions and in vacuum. Results are presented. In general vacuum plasma sprayed coatings performed much better than those sprayed under atmospheric conditions. A 50 to 50 blend of Cr2O3 and NiCrAlY, vacuum plasma sprayed on SSME turbopump turbine blades appear to provide significant improvements in coating durability and thermal protection.

  15. Implementing a Flip-Flop Teaching Model in Thermal Physics for Engineering Students

    Directory of Open Access Journals (Sweden)

    Dr. Emil C. Alcantara

    2015-11-01

    Full Text Available Implementing flip-flop teaching in a physics classroom allows students to learn concepts outside of the classroom and apply what they learn in the classroom, working with other students and getting immediate feedback from the instructor. The purpose of this study was to determine the effect of flip-flop teaching in the performance of engineering students in introductory physics particularly in thermal physics. The study employed descriptive and quasi-experimental method to describe and compare the performance of engineering students in thermal physics when grouped according to sex and types of instruction. Three physics classes consisting of 125 sophomore engineering students at the Batangas State University during the second semester of the SY 2013-2014 were handled by the researcher and selected purposively as participants of the study. It was found out that the variation in the performances of male and female students in the conceptual questions, in the problem solving questions, and overall performance in thermal physics are not significantly different. Male and female students have an overall satisfactory performance in thermal physics. The study also revealed that the variation in the performances of the students in the conceptual questions, in the problem solving questions, and overall performance in thermal physics when grouped according to the types of instruction are not significantly different. Engineering students taught in a traditional physics classroom, in a flipped physics classroom, and in an enhanced-flipped physics classroom are more likely to have similar performances in thermal physics.

  16. The kinematic Stirling engine as an energy conversion subsystem for paraboloidal dish solar thermal plants

    Science.gov (United States)

    Bowyer, J. M.

    1984-01-01

    The potential of a suitably designed and economically manufactured Stirling engine as the energy conversion subsystem of a paraboloidal dish-Stirling solar thermal power module was estimated. Results obtained by elementary cycle analyses were shown to match quite well the performance characteristics of an advanced kinematic Stirling engine, the United Stirling P-40, as established by current prototypes of the engine and by a more sophisticated analytic model of its advanced derivative. In addition to performance, brief consideration was given to other Stirling engine criteria such as durability, reliability, and serviceability. Production costs were not considered here.

  17. Source data for modeling of thermal engineering calculations

    Directory of Open Access Journals (Sweden)

    Charvátová Pavlína

    2018-01-01

    Full Text Available Increasing demands on thermal insulation. Their more accurate assessment by computers lead to increasingly bigger differences between computational models and reality. The result is an increasingly problematic optimization of building design. One of the key initial parameters is climatological data.

  18. Thermal Development Test of the NEXT PM1 Ion Engine

    Science.gov (United States)

    Anderson, John R.; Snyder, John S.; VanNoord, Jonathan L.; Soulas, George C.

    2010-01-01

    NASA's Evolutionary Xenon Thruster (NEXT) is a next-generation high-power ion propulsion system under development by NASA as a part of the In-Space Propulsion Technology Program. NEXT is designed for use on robotic exploration missions of the solar system using solar electric power. Potential mission destinations that could benefit from a NEXT Solar Electric Propulsion (SEP) system include inner planets, small bodies, and outer planets and their moons. This range of robotic exploration missions generally calls for ion propulsion systems with deep throttling capability and system input power ranging from 0.6 to 25 kW, as referenced to solar array output at 1 Astronomical Unit (AU). Thermal development testing of the NEXT prototype model 1 (PM1) was conducted at JPL to assist in developing and validating a thruster thermal model and assessing the thermal design margins. NEXT PM1 performance prior to, during and subsequent to thermal testing are presented. Test results are compared to the predicted hot and cold environments expected missions and the functionality of the thruster for these missions is discussed.

  19. Development and Performance of the 10 kN Hybrid Rocket Motor for the Stratos II Sounding Rocket

    NARCIS (Netherlands)

    Werner, R.M.; Knop, T.R.; Wink, J; Ehlen, J; Huijsman, R; Powell, S; Florea, R.; Wieling, W; Cervone, A.; Zandbergen, B.T.C.

    2016-01-01

    This paper presents the development work of the 10 kN hybrid rocket motor DHX-200 Aurora. The DHX-200 Aurora was developed by Delft Aerospace Rocket Engineering (DARE) to power the Stratos II and Stratos II+ sounding rocket, with the later one being launched in October 2015. Stratos II and Stratos

  20. Infrared nondestructive measurement of thermal resistance between liner and engine block: design of experiment

    Energy Technology Data Exchange (ETDEWEB)

    Laloue, P.; L' Ecolier, J.; Nigon, F. [PSA Peugeot Citroen, Laboratoire Optique et Thermique, 45 rue Jean-Pierre Timbaud, 78 300 Poissy (France); Bissieux, C.; Henry, J.-F.; Pron, H. [Universite de Reims, Unite de Thermique et Analyse Physique, EA 3802, Laboratoire de Thermophysique, UFR Sciences, Moulin de la Housse, BP 1039, 51 687 Reims Cedex 2 (France)

    2008-03-15

    Thermal resistances between liners and engine blocks are nondestructively studied by photothermal infrared thermography. Under controlled sinusoidal light irradiation, the thermal response of the sample is measured by means of an infrared camera. A numerical lock-in procedure yields amplitude and absolute phase maps of the thermal field periodic component. Then, apart from classical qualitative detection of air layers, a quantitative characterization of thermal resistance becomes available. An analytical modeling, associated with an inverse procedure using the Gauss-Newton parameter estimation method, allows to identify the thermal resistance on academic samples representative of the liner-engine block interface. Simply joined cast iron and aluminum plates present thermal resistances about 2 x 10{sup -3} K m{sup 2} W{sup -1}. The implementation of a numerical modeling allows to study two-dimensional defects. When the samples are pressed on their periphery, thus straightened, contact resistances ranging from 2 x 10{sup -4} to 7 x 10{sup -4} K m{sup 2} W{sup -1} have been measured. Then, the method is applied to liner-engine block interfaces where the thermal resistances fall to about 2 x 10{sup -5} K m{sup 2} W{sup -1}, matching the values obtained when a cast iron plate is locally pressed against an aluminum plate. (author)

  1. Proceedings ICTEA 2007, the 3. international conference on thermal engineering : theory and applications

    International Nuclear Information System (INIS)

    Akash, B.; Saghir, M.Z.

    2007-01-01

    This conference provided an opportunity to share research trends in thermal energy. It focused on the application of experimental, analytical or theoretical thermal and energy engineering. New technologies that improve the energy efficiency of engines, reduce exhaust emission levels and explore energy alternatives were highlighted along with market information and consumer education programs. A broad range of topics were addressed, including heat transfer; thermodiffusion; fluid mechanics; new and renewable energy technologies; environmental engineering; heat transfer with non-Newtonian fluid flow; polymer processing technology; energy management; solar thermal energy systems; air-conditioning and refrigeration; PV solar systems; and, energy conversion. The conference featured 152 presentations, of which 81 have been catalogued separately for inclusion in this database

  2. Evaluation of thermal efficiency and energy conversion of thermoacoustic Stirling engines

    International Nuclear Information System (INIS)

    Zhong Junhu; Zheng Yuli; Qing Li; Qiang Li

    2010-01-01

    Thermodynamic cycle transferring heat and work was executed in thermoacoustic engines, when the acoustic resonators substituted the moving mechanical components of the traditional heat engines. Based on the traveling-wave phasing and reversible heat transfer, thermoacoustic Stirling engines could achieve 70% of the Carnot efficiency theoretically, if the inevitable viscous dissipation in resonators was also counted as exported power. It should be pointed out an error on this efficiency evaluation in the previous literatures. More than 70% of the acoustic power production was often consumed by the side-branch resonator that was the essential configuration to build up a thermoacoustic Stirling engine. According to the simulation results and some experimental data, the actual available acoustic power consumed by the acoustic loads was restricted by the operating peak-to-mean pressure ratio, i.e. |p 1 /p m |. When the peak-to-mean pressure ratio operated on 4-6.5%, the thermal efficiency and power density of the available acoustic power reached higher levels. But the available acoustic power would approach zero when |p 1 /p m | attained 10%. It was approved that the turbulence oscillation occurred on the higher |p 1 /p m | (usually >4%) was the main reason of the excess dissipation in the side-branch resonator. This character of the available power limited the wide application of thermoacoustic Stirling engines. The evaluation of thermal efficiency and energy conversion also indicated the improving direction of thermoacoustic Stirling engines. Generators driven by the thermoacoustic Stirling engines were an effective way, due to the elimination of the side-branch resonator. To achieve a high power density and a high pressure ratio on the higher available power efficiency level, the standing-wave thermoacoustic engines might outvie the traveling-wave thermoacoustic engines. To enjoy the best features of standing-wave engines and traveling-wave engines simultaneously

  3. Interfacial Engineering of Silicon Carbide Nanowire/Cellulose Microcrystal Paper toward High Thermal Conductivity.

    Science.gov (United States)

    Yao, Yimin; Zeng, Xiaoliang; Pan, Guiran; Sun, Jiajia; Hu, Jiantao; Huang, Yun; Sun, Rong; Xu, Jian-Bin; Wong, Ching-Ping

    2016-11-16

    Polymer composites with high thermal conductivity have attracted much attention, along with the rapid development of electronic devices toward higher speed and better performance. However, high interfacial thermal resistance between fillers and matrix or between fillers and fillers has been one of the primary bottlenecks for the effective thermal conduction in polymer composites. Herein, we report on engineering interfacial structure of silicon carbide nanowire/cellulose microcrystal paper by generating silver nanostructures. We show that silver nanoparticle-deposited silicon carbide nanowires as fillers can effectively enhance the thermal conductivity of the matrix. The in-plane thermal conductivity of the resultant composite paper reaches as high as 34.0 W/m K, which is one order magnitude higher than that of conventional polymer composites. Fitting the measured thermal conductivity with theoretical models qualitatively demonstrates that silver nanoparticles bring the lower interfacial thermal resistances both at silicon carbide nanowire/cellulose microcrystal and silicon carbide nanowire/silicon carbide nanowire interfaces. This interfacial engineering approach provides a powerful tool for sophisticated fabrication of high-performance thermal-management materials.

  4. Numerical simulation of thermal loading produced by shaped high power laser onto engine parts

    International Nuclear Information System (INIS)

    Song Hongwei; Li Shaoxia; Zhang Ling; Yu Gang; Zhou Liang; Tan Jiansong

    2010-01-01

    Recently a new method for simulating the thermal loading on pistons of diesel engines was reported. The spatially shaped high power laser is employed as the heat source, and some preliminary experimental and numerical work was carried out. In this paper, a further effort was made to extend this simulation method to some other important engine parts such as cylinder heads. The incident Gaussian beam was transformed into concentric multi-circular patterns of specific intensity distributions, with the aid of diffractive optical elements (DOEs). By incorporating the appropriate repetitive laser pulses, the designed transient temperature fields and thermal loadings in the engine parts could be simulated. Thermal-structural numerical models for pistons and cylinder heads were built to predict the transient temperature and thermal stress. The models were also employed to find the optimal intensity distributions of the transformed laser beam that could produce the target transient temperature fields. Comparison of experimental and numerical results demonstrated that this systematic approach is effective in simulating the thermal loading on the engine parts.

  5. Nano-Material and Structural Engineering for Thermal Highways

    Science.gov (United States)

    2013-06-14

    engine self-lubrication, fuel cell supply and water removal, and drug delivery or biofluid sample extraction. The heating and cooling process of the...with the relation of SE=-10log(T). In order to obtain the IR reflectance, plasma frequency and real and imaginary parts of permittivity , the Drude...lubrication, fuel cell supply and water removal, and drug delivery or biofluid sample extraction. Another application that has garnered some interest is the

  6. Thermal Loss Determination for a Small Internal Combustion Engine

    Science.gov (United States)

    2014-03-27

    drivetrain shocks. A timing belt pulley system was mounted on pillow block bearings and was used to transfer power to the dynamometer while minimizing...efficiencies of less than 20% are common. The first measurement method was an energy balance between the fuel energy entering the system and the various avenues...for energy to leave the system . The second method used an enclosure around the engine and measured the enthalpy increase of the air flowing past the

  7. Development of a numerical tool to study the mixing phenomenon occurring during mode one operation of a multi-mode ejector-augmented pulsed detonation rocket engine

    Science.gov (United States)

    Dawson, Joshua

    A novel multi-mode implementation of a pulsed detonation engine, put forth by Wilson et al., consists of four modes; each specifically designed to capitalize on flow features unique to the various flow regimes. This design enables the propulsion system to generate thrust through the entire flow regime. The Multi-Mode Ejector-Augmented Pulsed Detonation Rocket Engine operates in mode one during take-off conditions through the acceleration to supersonic speeds. Once the mixing chamber internal flow exceeds supersonic speed, the propulsion system transitions to mode two. While operating in mode two, supersonic air is compressed in the mixing chamber by an upstream propagating detonation wave and then exhausted through the convergent-divergent nozzle. Once the velocity of the air flow within the mixing chamber exceeds the Chapman-Jouguet Mach number, the upstream propagating detonation wave no longer has sufficient energy to propagate upstream and consequently the propulsive system shifts to mode three. As a result of the inability of the detonation wave to propagate upstream, a steady oblique shock system is established just upstream of the convergent-divergent nozzle to initiate combustion. And finally, the propulsion system progresses on to mode four operation, consisting purely of a pulsed detonation rocket for high Mach number flight and use in the upper atmosphere as is needed for orbital insertion. Modes three and four appear to be a fairly significant challenge to implement, while the challenge of implementing modes one and two may prove to be a more practical goal in the near future. A vast number of potential applications exist for a propulsion system that would utilize modes one and two, namely a high Mach number hypersonic cruise vehicle. There is particular interest in the dynamics of mode one operation, which is the subject of this research paper. Several advantages can be obtained by use of this technology. Geometrically the propulsion system is fairly

  8. Calculation of the Thermal Loading of the Cylinder-Piston Group of the Automobile Engine

    Science.gov (United States)

    Barchenko, F. B.; Bakulin, V. N.

    2017-05-01

    We propose a mathematical model for calculating thermal loods of parts of the cylinder-piston group of the automobile engine operating under unstable conditions in its complete life cycle. Methods have been described for calculating the boundary conditions to determine the thermal state of the parts of the cylinder-piston group of such an engine with the use of theoretical formulas, empirical and semiempirical relations, and tabulated data. In modeling, we calculated the work of all systems of the engine (pumps, pipelines, heat exchangers) influencing directly or indirectly the thermal state of its cylinder-piston group. The nonstationary thermal state was calculated once in the operating cycle of the engine with the use of the cycle-averaged values of the local heat transfer coefficients and the resulting temperature of the medium. The personal computer counting time for one time step of a transport diesel engine of typical design with a number of units of the order of 500 was 5 s.

  9. Systems optimisation of an active thermal management system during engine warm-up

    OpenAIRE

    Burke, Richard D.; Lewis, Andrew J.; Akehurst, Sam; Brace, Chris J.; Pegg, Ian; Stark, Roland

    2012-01-01

    Active thermal management systems offer a potential for small improvements in fuel consumption that will contribute to upcoming legislation on carbon dioxide emissions. These systems offer new degrees of freedom for engine calibration; however, their full potential will only be exploited if a systems approach to their calibration is adopted, in conjunction with other engine controls. In this work, a design-of-experiments approach is extended to allow its application to transient drive cycles ...

  10. Simulation, design and thermal analysis of a solar Stirling engine using MATLAB

    International Nuclear Information System (INIS)

    Shazly, J.H.; Hafez, A.Z.; El Shenawy, E.T.; Eteiba, M.B.

    2014-01-01

    Highlights: • Modeling and simulation for a prototype of the solar-powered Stirling engine. • The solar-powered Stirling engine working at the low temperature range. • Estimating output power from the solar Stirling engine using Matlab program. • Solar radiation simulation program presents a solar radiation data using MATLAB. - Abstract: This paper presents the modeling and simulation for a prototype of the solar-powered Stirling engine working at the low temperature range. A mathematical model for the thermal analysis of the solar-powered low temperature Stirling engine with heat transfer is developed using Matlab program. The model takes into consideration the effect of the absorber temperature on the thermal analysis like as radiation and convection heat transfer between the absorber and the working fluid as well as radiation and convection heat transfer between the lower temperature plate and the working fluid. Hence, the present analysis provides a theoretical guidance for designing and operating of the solar-powered low temperature Stirling engine system, as well as estimating output power from the solar Stirling engine using Matlab program. This study attempts to demonstrate the potential of the low temperature Stirling engine as an option for the prime movers for Photovoltaic tracking systems. The heat source temperature is 40–60 °C as the temperature available from the sun directly

  11. Grain boundary engineering to enhance thermal stability of electrodeposited nickel

    DEFF Research Database (Denmark)

    Alimadadi, Hossein

    by miniaturization of the grains down to nano-meter scale. However, this augments the total grain boundary energy stored in the material, hence, making the material less thermally stable. Coherent twin boundaries are of very low energy and mobility compared to all other boundaries in a FCC material. Accordingly...... interest. The evolution of microstructure in as-deposited and annealed condition was investigated with a combination of complementary microscopic techniques, electron backscatter diffraction (EBSD), electron channelling contrast imaging (ECCI), ion channelling contrast imaging (ICCI), and, for the as...

  12. A technology data base for the design of 500 to 5000-lb thrust class liquid rocket engines utilizing hydrogen and oxygen as propellants

    Science.gov (United States)

    Schoenman, L.

    1982-01-01

    This paper presents an overview of the results of experimental evaluations of candidate designs for igniters, injectors, and propellant-cooled thrust chambers applicable to restartable high-performance, high-reliability upper-stage engines and to pulsing-type reaction control engines (RCE). Injection element types best suited for liquid, gas, and liquid/gas phase propellant supply are identified. The resulting interactions between element type, combustion efficiency, and chamber wall heating are compared. The distinction between thrust chamber design requirements for upper stage vs RCE applications as measured by cycle life requirements is translated into design configurations consisting of all-film-cooled, all-regeneratively-cooled, and composites of the two cooling approaches. The validity of the design approaches is confirmed by data from engine durability testing involving over 90,000 starts and 9,000 thermal cycles on RCE-type designs and multiple long-duration burns (up to 2,000 sec) on regeneratively cooled upper-stage designs.

  13. Solution of neutronic and thermal-hydraulic problems on an engineering work station

    International Nuclear Information System (INIS)

    Zee, S.K.; Sills, E.D.; Turinsky, P.J.; Doster, J.M.

    1986-01-01

    Interest is in developing neutronic and thermal-hydraulic computer programs that execute efficiently on advanced engineering work stations. Engineering work stations are characterized by a 32-bit arithmetic processor, graphics capabilities, and networking capabilities. These attributes allow an engineer to solve substantive problems in a graphical interactive environment with shared resources available via networking. An advanced engineering work station is further characterized as having computational capability comparable to a mainframe, achieved via a parallel computer architecture obtained by both multi-central processing units (CPUs) and vector pipelines. In this paper, the authors present timing studies completed on an engineering work station, and then extrapolate performance on an advanced engineering work station using results from a supercomputer with parallel architecture. In this paper, the authors report on two codes, a neutronic code and a LWR system's thermal-hydraulic code. The neutronic code solves the two-group, two-dimensional (x-y) neutron diffusion equations using the finite difference method. The system's thermal-hydraulic codes solves the mixture drift-flux representation of the tube-stream form of the Navier-Stokes equations (four-equation model)

  14. Microwave Thermal Propulsion

    Science.gov (United States)

    Parkin, Kevin L. G.; Lambot, Thomas

    2017-01-01

    We have conducted research in microwave thermal propulsion as part of the space exploration access technologies (SEAT) research program, a cooperative agreement (NNX09AF52A) between NASA and Carnegie Mellon University. The SEAT program commenced on the 19th of February 2009 and concluded on the 30th of September 2015. The DARPA/NASA Millimeter-wave Thermal Launch System (MTLS) project subsumed the SEAT program from May 2012 to March 2014 and one of us (Parkin) served as its principal investigator and chief engineer. The MTLS project had no final report of its own, so we have included the MTLS work in this report and incorporate its conclusions here. In the six years from 2009 until 2015 there has been significant progress in millimeter-wave thermal rocketry (a subset of microwave thermal rocketry), most of which has been made under the auspices of the SEAT and MTLS programs. This final report is intended for multiple audiences. For researchers, we present techniques that we have developed to simplify and quantify the performance of thermal rockets and their constituent technologies. For program managers, we detail the facilities that we have built and the outcomes of experiments that were conducted using them. We also include incomplete and unfruitful lines of research. For decision-makers, we introduce the millimeter-wave thermal rocket in historical context. Considering the economic significance of space launch, we present a brief but significant cost-benefit analysis, for the first time showing that there is a compelling economic case for replacing conventional rockets with millimeter-wave thermal rockets.

  15. Application of thermal barrier coating for improving the suitability of Annona biodiesel in a diesel engine

    Directory of Open Access Journals (Sweden)

    Ramalingam Senthil

    2016-01-01

    Full Text Available The Annona biodiesel was produced from Annona oil through transesterification process. The aim of the present study is to analyze the performance and emission characteristics of a single cylinder, direct injection, compression ignition engine using a annona methyl ester as a fuel. They are blended together with the Neat diesel fuel such as 20%, 40%, 60%, 80%, and Neat biodiesel. The performance, emission and combustion characteristics are evaluated by operating the engine at different loads. The performance parameters such as brake thermal efficiency, brake specific fuel consumption. The emission constituents such as carbon monoxide, unburned hydrocarbons, oxides of nitrogen, and smoke were recorded. Then the piston and both exhaust and intake valves of the test engine were coated with 100 µm of NiCrAl as lining layer. Later the same parts were coated with 400 µm material of coating that was the mixture of 88% of ZrO2, 4% of MgO, and 8% of Al2O3. After the engine coating process, the same fuels is tested in the engine at the same engine operation. The same performance and emission parameters were evaluated. Finally, these parameters are compared with uncoated engine in order to find out the changes in the performance and emission parameters of the coated engine. It is concluded that the coating engine resulting in better performance, especially in considerably lower brake specific fuel consumption values. The engine emissions are lowered both through coating and annona methyl ester biodiesel expect the nitrogen oxides emission.

  16. An investigation of enhanced capability thermal barrier coating systems for diesel engine components

    Science.gov (United States)

    Holtzman, R. L.; Layne, J. L.; Schechter, B.

    1984-01-01

    Material systems and processes for the development of effective and durable thermal barriers for heavy duty diesel engines were investigated. Seven coating systems were evaluated for thermal conductivity, erosion resistance, corrosion/oxidation resistance, and thermal shock resistance. An advanced coating system based on plasma sprayed particle yttria stabilized zirconia (PS/HYSZ) was judged superior in these tests. The measured thermal conductivity of the selected coating was 0.893 W/m C at 371 C. The PS/HYSZ coating system was applied to the piston crown, fire deck and valves of a single cylinder low heat rejection diesel engine. The coated engine components were tested for 24 hr at power levels from 0.83 MPa to 1.17 MPa brake mean effective pressure. The component coatings survived the engine tests with a minimum of distress. The measured fire deck temperatures decreased 86 C (155 F) on the intake side and 42 C (75 F) on the exhaust side with the coating applied.

  17. Loadings in thermal barrier coatings of jet engine turbine blades an experimental research and numerical modeling

    CERN Document Server

    Sadowski, Tomasz

    2016-01-01

    This book discusses complex loadings of turbine blades and protective layer Thermal Barrier Coating (TBC), under real working airplane jet conditions. They obey both multi-axial mechanical loading and sudden temperature variation during starting and landing of the airplanes. In particular, two types of blades are analyzed: stationary and rotating, which are widely applied in turbine engines produced by airplane factories.

  18. Thermoacoustic model of a modified free piston Stirling engine with a thermal buffer tube

    International Nuclear Information System (INIS)

    Yang, Qin; Luo, Ercang; Dai, Wei; Yu, Guoyao

    2012-01-01

    This article presents a modified free-piston Stirling heat engine configuration in which a thermal buffer tube is added to sandwich between the hot and cold heat exchangers. Such a modified configuration may lead to an easier fabrication and lighter weight of a free piston. To analyze the thermodynamic performance of the modified free piston Stirling heat engine, thermoacoustic theory is used. In the thermoacoustic modelling, the regenerator, the free piston, and the thermal buffer tube are given at first. Then, based on linear thermoacoustic network theory, the thermal and thermodynamic networks are presented to characterize acoustic pressure and volume flow rate distributions at different interfaces, and the global performance such as the power output, the heat input and the thermal efficiency. A free piston Stirling heat engine with several hundreds of watts mechanical power output is selected as an example. The typical operating and structure parameters are as follows: frequency around 50 Hz, mean pressure around 3.0 MPa, and a diameter of free piston around 50 mm. From the analysis, it was found that the modified free-piston Stirling heat engine has almost the same thermodynamic performance as the original design, which indicates that the modified configuration is worthy to develop in future because of its mechanical simplicity and reliability.

  19. Thermal Management as a Force Multiplier within the Research, Development, and Engineering Command (RDECOM)

    Science.gov (United States)

    2012-08-01

    we mean energy that has low availability to do work (low exergy ). The closer a system is to the condition of its surroundings in terms of...temperature (4). Similarly, inefficiencies in a turbine engine for an Abrams tank means that exhaust gasses can contain megawatts of thermal energy. Thus

  20. The performance of solar thermal electric power systems employing small heat engines

    Science.gov (United States)

    Pons, R. L.

    1980-01-01

    The paper presents a comparative analysis of small (10 to 100 KWe) heat engines for use with a solar thermal electric system employing the point-focusing, distributed receiver (PF-DR) concept. Stirling, Brayton, and Rankine cycle engines are evaluated for a nominal overall system power level of 1 MWe, although the concept is applicable to power levels up to at least 10 MWe. Multiple concentrators are electrically connected to achieve the desired plant output. Best performance is achieved with the Stirling engine, resulting in a system Levelized Busbar Energy Cost of just under 50 mills/kWH and a Capital Cost of $900/kW, based on the use of mass-produced components. Brayton and Rankine engines show somewhat less performance but are viable alternatives with particular benefits for special applications. All three engines show excellent performance for the small community application.

  1. Thermal conductivity engineering of bulk and one-dimensional Si-Ge nanoarchitectures.

    Science.gov (United States)

    Kandemir, Ali; Ozden, Ayberk; Cagin, Tahir; Sevik, Cem

    2017-01-01

    Various theoretical and experimental methods are utilized to investigate the thermal conductivity of nanostructured materials; this is a critical parameter to increase performance of thermoelectric devices. Among these methods, equilibrium molecular dynamics (EMD) is an accurate technique to predict lattice thermal conductivity. In this study, by means of systematic EMD simulations, thermal conductivity of bulk Si-Ge structures (pristine, alloy and superlattice) and their nanostructured one dimensional forms with square and circular cross-section geometries (asymmetric and symmetric) are calculated for different crystallographic directions. A comprehensive temperature analysis is evaluated for selected structures as well. The results show that one-dimensional structures are superior candidates in terms of their low lattice thermal conductivity and thermal conductivity tunability by nanostructuring, such as by diameter modulation, interface roughness, periodicity and number of interfaces. We find that thermal conductivity decreases with smaller diameters or cross section areas. Furthermore, interface roughness decreases thermal conductivity with a profound impact. Moreover, we predicted that there is a specific periodicity that gives minimum thermal conductivity in symmetric superlattice structures. The decreasing thermal conductivity is due to the reducing phonon movement in the system due to the effect of the number of interfaces that determine regimes of ballistic and wave transport phenomena. In some nanostructures, such as nanowire superlattices, thermal conductivity of the Si/Ge system can be reduced to nearly twice that of an amorphous silicon thermal conductivity. Additionally, it is found that one crystal orientation, [Formula: see text]100[Formula: see text], is better than the [Formula: see text]111[Formula: see text] crystal orientation in one-dimensional and bulk SiGe systems. Our results clearly point out the importance of lattice thermal conductivity

  2. Resonant enhancement in nanostructured thermoelectric performance via electronic thermal conductivity engineering

    Science.gov (United States)

    Patil, Urvesh; Muralidharan, Bhaskaran

    2017-01-01

    The use of an asymmetric broadening in the transport distribution, a characteristic of resonant structures, is proposed as a route to engineer a decrease in electronic thermal conductivity thereby enhancing the electronic figure of merit in nanostructured thermoelectrics. Using toy models, we first demonstrate that a decrease in thermal conductivity resulting from such an asymmetric broadening may indeed lead to an electronic figure of merit well in excess of 1000 in an idealized situation and in excess of 10 in a realistic situation. We then substantiate with realistic resonant structures designed using graphene nano-ribbons by employing a tight binding framework with edge correction that match density functional theory calculations under the local density approximation. The calculated figure of merit exceeding 10 in such realistic structures further reinforces the concept and sets a promising direction to use nano-ribbon structures to engineer a favorable decrease in the electronic thermal conductivity.

  3. FAILURE MECHANISMS OF THERMAL BARRIER COATINGS INTERNAL COMBUSTION ENGINES AND llMPROVEMENTS

    Directory of Open Access Journals (Sweden)

    ADNAN PARLAK

    2003-04-01

    Full Text Available MechanicaJ properties of high performance ceramics have been improved to the point where their use in heat engines is possible. The high temperature strength and low thermal expansion properties of bigh performance ceramics offer an advantage over metals in the development of non-water cooling engine. However, because bard environment in diesel engine combustion chamber, solving the problem of durabiUty of TBC is important. DurabiUty of thermal barrier coatings(TBC is liınited by two main failure mechanisms: Therınal expansion nlİsmatch betwcen bond coat and top coat and bond coat oxidation. Both of these can cause failure of the ceramic top coat. Developments of recent years sholv that bond coats \\Vith higher oxidation resistance tend to have better coating system cyclic lives

  4. Engineering process instructions and development summary MC3642 thermal battery

    Energy Technology Data Exchange (ETDEWEB)

    Jacobs, D.

    1981-06-01

    The MC3642 is a dual channel thermal battery used on the DE1010/W85 Command Disable Controller. It utilizes the CalCaCrO{sub 4} electrochemical system. The electrical requirements of this battery are as follows: RISE TIME PEAK VOLTAGE ACTIVE LIFE LOAD Channel 1 - 1.0 Sec. Max. 34 Volts 10 Sec. Min. 40.0 Ohms to 20 Volts above 20 Volts Channel 2 - .350 Sec. Max. 42 Volts 10 MSec. Min. 6.5 Ohms to 23 Volts above 23 Volts The battery consists of 14 cells connected in series (Channel 2) and 12 cells connected in series (Channel 1). Each cell is composed of an anode fabricated from a bimetallic sheet (0.005{double_prime} thick calcium on 0.005{double_prime} thick iron substrate), a depolarizer-electrolyte-binder (DEB) pellet and a heat pellet. Activation is achieved by mechanical primer. Optimum battery performance is achieved with a 35155/10 DEB pellet weighing .80g and a heat pellet, weighing 1.30 grams, of 88/12 heat powder.

  5. Designing a solar powered Stirling heat engine based on multiple criteria: Maximized thermal efficiency and power

    International Nuclear Information System (INIS)

    Ahmadi, Mohammad Hossein; Sayyaadi, Hoseyn; Dehghani, Saeed; Hosseinzade, Hadi

    2013-01-01

    Highlights: • Thermodynamic model of a solar-dish Stirling engine was presented. • Thermal efficiency and output power of the engine were simultaneously maximized. • A final optimal solution was selected using several decision-making methods. • An optimal solution with least deviation from the ideal design was obtained. • Optimal solutions showed high sensitivity against variation of system parameters. - Abstract: A solar-powered high temperature differential Stirling engine was considered for optimization using multiple criteria. A thermal model was developed so that the output power and thermal efficiency of the solar Stirling system with finite rate of heat transfer, regenerative heat loss, conductive thermal bridging loss, finite regeneration process time and imperfect performance of the dish collector could be obtained. The output power and overall thermal efficiency were considered for simultaneous maximization. Multi-objective evolutionary algorithms (MOEAs) based on the NSGA-II algorithm were employed while the solar absorber temperature and the highest and lowest temperatures of the working fluid were considered the decision variables. The Pareto optimal frontier was obtained and a final optimal solution was also selected using various decision-making methods including the fuzzy Bellman–Zadeh, LINMAP and TOPSIS. It was found that multi-objective optimization could yield results with a relatively low deviation from the ideal solution in comparison to the conventional single objective approach. Furthermore, it was shown that, if the weight of thermal efficiency as one of the objective functions is considered to be greater than weight of the power objective, lower absorber temperature and low temperature ratio should be considered in the design of the Stirling engine

  6. Effects of water-emulsified fuel on a diesel engine generator's thermal efficiency and exhaust.

    Science.gov (United States)

    Syu, Jin-Yuan; Chang, Yuan-Yi; Tseng, Chao-Heng; Yan, Yeou-Lih; Chang, Yu-Min; Chen, Chih-Chieh; Lin, Wen-Yinn

    2014-08-01

    Water-emulsified diesel has proven itself as a technically sufficient improvement fuel to improve diesel engine fuel combustion emissions and engine performance. However, it has seldom been used in light-duty diesel engines. Therefore, this paper focuses on an investigation into the thermal efficiency and pollution emission analysis of a light-duty diesel engine generator fueled with different water content emulsified diesel fuels (WD, including WD-0, WD-5, WD-10, and WD-15). In this study, nitric oxide, carbon monoxide, hydrocarbons, and carbon dioxide were analyzed by a vehicle emission gas analyzer and the particle size and number concentration were measured by an electrical low-pressure impactor. In addition, engine loading and fuel consumption were also measured to calculate the thermal efficiency. Measurement results suggested that water-emulsified diesel was useful to improve the thermal efficiency and the exhaust emission of a diesel engine. Obviously, the thermal efficiency was increased about 1.2 to 19.9%. In addition, water-emulsified diesel leads to a significant reduction of nitric oxide emission (less by about 18.3 to 45.4%). However the particle number concentration emission might be increased if the loading of the generator becomes lower than or equal to 1800 W. In addition, exhaust particle size distributions were shifted toward larger particles at high loading. The consequence of this research proposed that the water-emulsified diesel was useful to improve the engine performance and some of exhaust emissions, especially the NO emission reduction. Implications: The accumulated test results provide a good basis to resolve the corresponding pollutants emitted from a light-duty diesel engine generator. By measuring and analyzing transforms of exhaust pollutant from this engine generator, the effects of water-emulsified diesel fuel and loading on emission characteristics might be more clear. Understanding reduction of pollutant emissions during the use

  7. Small rocket research and technology

    Science.gov (United States)

    Schneider, Steven; Biaglow, James

    1993-11-01

    Small chemical rockets are used on nearly all space missions. The small rocket program provides propulsion technology for civil and government space systems. Small rocket concepts are developed for systems which encompass reaction control for launch and orbit transfer systems, as well as on-board propulsion for large space systems and earth orbit and planetary spacecraft. Major roles for on-board propulsion include apogee kick, delta-V, de-orbit, drag makeup, final insertions, north-south stationkeeping, orbit change/trim, perigee kick, and reboost. The program encompasses efforts on earth-storable, space storable, and cryogenic propellants. The earth-storable propellants include nitrogen tetroxide (NTO) as an oxidizer with monomethylhydrazine (MMH) or anhydrous hydrazine (AH) as fuels. The space storable propellants include liquid oxygen (LOX) as an oxidizer with hydrazine or hydrocarbons such as liquid methane, ethane, and ethanol as fuels. Cryogenic propellants are LOX or gaseous oxygen (GOX) as oxidizers and liquid or gaseous hydrogen as fuels. Improved performance and lifetime for small chemical rockets are sought through the development of new predictive tools to understand the combustion and flow physics, the introduction of high temperature materials to eliminate fuel film cooling and its associated combustion inefficiency, and improved component designs to optimize performance. Improved predictive technology is sought through the comparison of both local and global predictions with experimental data. Results indicate that modeling of the injector and combustion process in small rockets needs improvement. High temperature materials require the development of fabrication processes, a durability data base in both laboratory and rocket environments, and basic engineering property data such as strength, creep, fatigue, and work hardening properties at both room and elevated temperature. Promising materials under development include iridium-coated rhenium and a

  8. Computational and Experimental Investigation of Liquid Propellant Rocket Combustion Instability

    Data.gov (United States)

    National Aeronautics and Space Administration — Combustion instability has been a problem faced by rocket engine developers since the 1940s. The complicated phenomena has been highly unpredictable, causing engine...

  9. Methylcellulose Based Thermally Reversible Hydrogel System for Tissue Engineering Applications

    Directory of Open Access Journals (Sweden)

    Ram V. Devireddy

    2013-06-01

    Full Text Available The thermoresponsive behavior of a Methylcellulose (MC polymer was systematically investigated to determine its usability in constructing MC based hydrogel systems in cell sheet engineering applications. Solution-gel analyses were made to study the effects of polymer concentration, molecular weight and dissolved salts on the gelation of three commercially available MCs using differential scanning calorimeter and rheology. For investigation of the hydrogel stability and fluid uptake capacity, swelling and degradation experiments were performed with the hydrogel system exposed to cell culture solutions at incubation temperature for several days. From these experiments, the optimal composition of MC-water-salt that was able to produce stable hydrogels at or above 32 °C, was found to be 12% to 16% of MC (Mol. wt. of 15,000 in water with 0.5× PBS (~150mOsm. This stable hydrogel system was then evaluated for a week for its efficacy to support the adhesion and growth of specific cells in culture; in our case the stromal/stem cells derived from human adipose tissue derived stem cells (ASCs. The results indicated that the addition (evenly spread of ~200 µL of 2 mg/mL bovine collagen type -I (pH adjusted to 7.5 over the MC hydrogel surface at 37 °C is required to improve the ASC adhesion and proliferation. Upon confluence, a continuous monolayer ASC sheet was formed on the surface of the hydrogel system and an intact cell sheet with preserved cell–cell and cell–extracellular matrix was spontaneously and gradually detached when the grown cell sheet was removed from the incubator and exposed to room temperature (~30 °C within minutes.

  10. Methylcellulose based thermally reversible hydrogel system for tissue engineering applications.

    Science.gov (United States)

    Thirumala, Sreedhar; Gimble, Jeffrey M; Devireddy, Ram V

    2013-06-25

    The thermoresponsive behavior of a Methylcellulose (MC) polymer was systematically investigated to determine its usability in constructing MC based hydrogel systems in cell sheet engineering applications. Solution-gel analyses were made to study the effects of polymer concentration, molecular weight and dissolved salts on the gelation of three commercially available MCs using differential scanning calorimeter and rheology. For investigation of the hydrogel stability and fluid uptake capacity, swelling and degradation experiments were performed with the hydrogel system exposed to cell culture solutions at incubation temperature for several days. From these experiments, the optimal composition of MC-water-salt that was able to produce stable hydrogels at or above 32 °C, was found to be 12% to 16% of MC (Mol. wt. of 15,000) in water with 0.5× PBS (~150mOsm). This stable hydrogel system was then evaluated for a week for its efficacy to support the adhesion and growth of specific cells in culture; in our case the stromal/stem cells derived from human adipose tissue derived stem cells (ASCs). The results indicated that the addition (evenly spread) of ~200 µL of 2 mg/mL bovine collagen type -I (pH adjusted to 7.5) over the MC hydrogel surface at 37 °C is required to improve the ASC adhesion and proliferation. Upon confluence, a continuous monolayer ASC sheet was formed on the surface of the hydrogel system and an intact cell sheet with preserved cell-cell and cell-extracellular matrix was spontaneously and gradually detached when the grown cell sheet was removed from the incubator and exposed to room temperature (~30 °C) within minutes.

  11. Replacement of chemical rocket launchers by beamed energy propulsion.

    Science.gov (United States)

    Fukunari, Masafumi; Arnault, Anthony; Yamaguchi, Toshikazu; Komurasaki, Kimiya

    2014-11-01

    Microwave Rocket is a beamed energy propulsion system that is expected to reach space at drastically lower cost. This cost reduction is estimated by replacing the first-stage engine and solid rocket boosters of the Japanese H-IIB rocket with Microwave Rocket, using a recently developed thrust model in which thrust is generated through repetitively pulsed microwave detonation with a reed-valve air-breathing system. Results show that Microwave Rocket trajectory, in terms of velocity versus altitude, can be designed similarly to the current H-IIB first stage trajectory. Moreover, the payload ratio can be increased by 450%, resulting in launch-cost reduction of 74%.

  12. Nuclear Thermal Propulsion: Past, Present, and a Look Ahead

    Science.gov (United States)

    Borowski, Stanley K.

    2014-01-01

    NTR: High thrust high specific impulse (2 x LOXLH2 chemical) engine uses high power density fission reactor with enriched uranium fuel as thermal power source. Reactor heat is removed using H2 propellant which is then exhausted to produce thrust. Conventional chemical engine LH2 tanks, turbo pumps, regenerative nozzles and radiation-cooled shirt extensions used -- NTR is next evolutionary step in high performance liquid rocket engines.

  13. Local strain field engineering on interfacial thermal resistance of graphene nanoribbon

    Science.gov (United States)

    Xue, Yixuan; Chen, Yang; Cai, Kun; Liu, Zi-Yu; Zhang, Yingyan; Wei, Ning

    2018-01-01

    Strain engineering shows distinct advantages in thermal management by tuning thermal resistance in a wide range. Till now, most of the relative studies were concentrated in uniform deformation, wherein the effects of the localized strain field are rarely exploited. Herein, by using non-equilibrium molecular dynamics simulations, we explore the local strain field engineering effects on the interfacial thermal resistance (ITR) of graphene nanoribbons (GNRs). The model of GNRs employed in this work contains extended drag threads, which are used to create a local strain field. Our simulation results show that the ITR has a quasi-linear relationship with the local tensile strain. GNRs are very sensitive to the local strain field in terms of ITR with a maximum enhancement factor of ˜1.5 at the strain of 10%. The ITR is found to depend linearly on the local strain. This phenomenon is thoroughly explained by micro-structure deformation, heat flux scattering, and phonon density of state overlapping. Our findings here offer a simple yet useful tool in modulating the thermal properties of graphene and other two-dimensional materials by using local strain engineering.

  14. Thermally enhanced photoluminescence for energy harvesting: from fundamentals to engineering optimization

    Science.gov (United States)

    Kruger, N.; Kurtulik, M.; Revivo, N.; Manor, A.; Sabapathy, T.; Rotschild, C.

    2018-05-01

    The radiance of thermal emission, as described by Planck’s law, depends only on the emissivity and temperature of a body, and increases monotonically with the temperature rise at any emitted wavelength. Non-thermal radiation, such as photoluminescence (PL), is a fundamental light–matter interaction that conventionally involves the absorption of an energetic photon, thermalization, and the emission of a redshifted photon. Such a quantum process is governed by rate conservation, which is contingent on the quantum efficiency. In the past, the role of rate conservation for significant thermal excitation had not been studied. Recently, we presented the theory and an experimental demonstration that showed, in contrast to thermal emission, that the PL rate is conserved when the temperature increases while each photon is blueshifted. A further rise in temperature leads to an abrupt transition to thermal emission where the photon rate increases sharply. We also demonstrated how such thermally enhanced PL (TEPL) generates orders of magnitude more energetic photons than thermal emission at similar temperatures. These findings show that TEPL is an ideal optical heat pump that can harvest thermal losses in photovoltaics with a maximal theoretical efficiency of 70%, and practical concepts potentially reaching 45% efficiency. Here we move the TEPL concept onto the engineering level and present Cr:Nd:YAG as device grade PL material, absorbing solar radiation up to 1 μm wavelength and heated by thermalization of energetic photons. Its blueshifted emission, which can match GaAs cells, is 20% of the absorbed power. Based on a detailed balance simulation, such a material coupled with proper photonic management can reach 34% power conversion efficiency. These results raise confidence in the potential of TEPL becoming a disruptive technology in photovoltaics.

  15. Fundamental limitations of non-thermal plasma processing for internal combustion engine NOx control

    International Nuclear Information System (INIS)

    Penetrante, B.M.

    1993-01-01

    This paper discusses the physics and chemistry of non-thermal plasma processing for post-combustion NO x control in internal combustion engines. A comparison of electron beam and electrical discharge processing is made regarding their power consumption, radical production, NO x removal mechanisms, and by product formation. Can non-thermal deNO x operate efficiently without additives or catalysts? How much electrical power does it cost to operate? What are the by-products of the process? This paper addresses these fundamental issues based on an analysis of the electron-molecule processes and chemical kinetics

  16. Rockets two classic papers

    CERN Document Server

    Goddard, Robert

    2002-01-01

    Rockets, in the primitive form of fireworks, have existed since the Chinese invented them around the thirteenth century. But it was the work of American Robert Hutchings Goddard (1882-1945) and his development of liquid-fueled rockets that first produced a controlled rocket flight. Fascinated by rocketry since boyhood, Goddard designed, built, and launched the world's first liquid-fueled rocket in 1926. Ridiculed by the press for suggesting that rockets could be flown to the moon, he continued his experiments, supported partly by the Smithsonian Institution and defended by Charles Lindbergh. T

  17. Energy, Entropy and Exergy Concepts and Their Roles in Thermal Engineering

    Directory of Open Access Journals (Sweden)

    Yunus A. Cengel

    2001-08-01

    Full Text Available Abstract: Energy, entropy and exergy concepts come from thermodynamics and are applicable to all fields of science and engineering. Therefore, this article intends to provide background for better understanding of these concepts and their differences among various classes of life support systems with a diverse coverage. It also covers the basic principles, general definitions and practical applications and implications. Some illustrative examples are presented to highlight the importance of the aspects of energy, entropy and exergy and their roles in thermal engineering.

  18. Progress Toward Quality Assurance Standards for Advanced Hydrocarbon Fuels Based on Thermal Performance Testing and Chemometric Modeling

    Science.gov (United States)

    2015-12-15

    hydrocarbon-fueled liquid rocket engines , combustion enthalpy is transferred at high rates to thrust chamber surfaces, which are maintained at acceptably...thermal management of engine and vehicle structures , and more precisely the fuel’s ability to absorb heat without detrimentally affecting cooling...Specifically, reliable and predictable thermal management of engine and vehicle structures , and more precisely the fuel’s ability to absorb heat without

  19. A new closed-form thermodynamic model for thermal simulation of spark ignition internal combustion engines

    International Nuclear Information System (INIS)

    Barjaneh, Afshin; Sayyaadi, Hoseyn

    2015-01-01

    Highlights: • A new closed-form thermal model was developed for SI engines. • Various irreversibilities of real engines were integrated into the model. • The accuracy of the model was examined on two real SI engines. • The superiority of the model to previous closed-form models was shown. • Accuracy and losses were studied over the operating range of engines. - Abstract: A closed form model based on finite speed thermodynamics, FST, modified to consider various losses was developed on Otto cycle. In this regard, the governing equations of the finite speed thermodynamics were developed for expansion/compression processes while heat absorption/rejection of the Otto cycle was determined based on finite time thermodynamics, FTT. In addition, other irreversibility including power loss caused by heat transfer through the cylinder walls and irreversibility due to throttling process was integrated into the model. The developed model was verified by implementing on two different spark ignition internal combustion engines and the results of modeling were compared with experimental results as well as FTT model. It was found that the developed model was not only very simple in use like a closed form thermodynamic model, but also it models a real spark ignition engine with reasonable accuracy. The error in predicting the output power at rated operating range of the engine was 39%, while in the case of the FTT model, this figure was 167.5%. This comparison for predicting thermal efficiency was +7% error (as difference) for the developed model compared to +39.4% error of FTT model.

  20. Strain engineering of phonon thermal transport properties in monolayer 2H-MoTe2.

    Science.gov (United States)

    Shafique, Aamir; Shin, Young-Han

    2017-12-06

    The effect of strain on the phonon properties such as phonon group velocity, phonon anharmonicity, phonon lifetime, and lattice thermal conductivity of monolayer 2H-MoTe 2 is studied by solving the Boltzmann transport equation based on first principles calculations. The phonon thermal transport properties of the unstrained monolayer 2H-MoTe 2 are compared to those of the strained case under different biaxial tensile strains. One of the common features of two-dimensional materials is the quadratic nature near the Γ point of the out-of-plane phonon flexural mode that disappears by applying tensile strain. We find that the lattice thermal conductivity of the monolayer 2H-MoTe 2 is very sensitive to strain, and the lattice thermal conductivity is reduced by approximately 2.5 times by applying 8% biaxial tensile strain due to the reduction in phonon group velocities and phonon lifetime. We also analyze how the contribution of each mode to lattice thermal conductivity changes with tensile strain. These results highlight that tensile strain is a key parameter in engineering phonon thermal transport properties in monolayer 2H-MoTe 2 .

  1. Mechanical and thermal analysis of the internal combustion engine piston using Ansys

    Science.gov (United States)

    Cioată, V. G.; Kiss, I.; Alexa, V.; Raţiu, S. A.

    2017-01-01

    The piston is one of the most important components of the internal combustion engine. Piston fail mainly due to mechanical stresses and thermal stresses. In this paper is determined by using the finite element method, stress and displacement distribution due the flue gas pressure and temperature, separately and combined. The FEA is performed by CAD and CAE software. The results are compared with those obtained by the analytical method and conclusions have been drawn.

  2. Performance and emission characteristics of the thermal barrier coated SI engine by adding argon inert gas to intake mixture.

    Science.gov (United States)

    Karthikeya Sharma, T

    2015-11-01

    Dilution of the intake air of the SI engine with the inert gases is one of the emission control techniques like exhaust gas recirculation, water injection into combustion chamber and cyclic variability, without scarifying power output and/or thermal efficiency (TE). This paper investigates the effects of using argon (Ar) gas to mitigate the spark ignition engine intake air to enhance the performance and cut down the emissions mainly nitrogen oxides. The input variables of this study include the compression ratio, stroke length, and engine speed and argon concentration. Output parameters like TE, volumetric efficiency, heat release rates, brake power, exhaust gas temperature and emissions of NOx, CO2 and CO were studied in a thermal barrier coated SI engine, under variable argon concentrations. Results of this study showed that the inclusion of Argon to the input air of the thermal barrier coated SI engine has significantly improved the emission characteristics and engine's performance within the range studied.

  3. Thermal modeling of a novel thermosyphonic waste heat absorption system for internal combustion engines

    International Nuclear Information System (INIS)

    Nwosu, Paul Nwachukwu; Nuutinen, Mika; Larmi, Martti

    2014-01-01

    This paper investigates a thermal system that absorbs waste heat from an internal combustion (IC) engine in order to raise the temperature of a working fluid to a saturated state using thermosyphonic flow, non-intrusive of the engine operations. The absorbed heat is rejected to an enclosed space, suitable for in-transit drying. The thermal system comprises a cross-flow heat exchanger connected to a radiator which preheats the working fluid from an insulated (storage) tank. The preheated fluid flows through a radiant heat absorber which absorbs radiant heat from the exhaust manifold. To ensure that the system efficiently performs, a temperature differential is maintained by the heated space while the fluid is cyclically delivered to the tank. The system’s operations are described using a novel flow cycle, and the results indicate a significant heat recovery potential. - Highlights: • This paper investigates a thermal system that absorbs waste heat from an internal combustion (IC) engine. • The absorbed heat is used to raise the temperature of a working fluid employing thermosyphonic flow. • The preheated fluid flows through a radiant heat absorber which absorbs radiant heat from the exhaust manifold. • To ensure that the system efficiently performs, a temperature differential is maintained by a heated space. • The system's operations are described using a novel flow cycle

  4. Thermal conductivity engineering in width-modulated silicon nanowires and thermoelectric efficiency enhancement

    Science.gov (United States)

    Zianni, Xanthippi

    2018-03-01

    Width-modulated nanowires have been proposed as efficient thermoelectric materials. Here, the electron and phonon transport properties and the thermoelectric efficiency are discussed for dimensions above the quantum confinement regime. The thermal conductivity decreases dramatically in the presence of thin constrictions due to their ballistic thermal resistance. It shows a scaling behavior upon the width-modulation rate that allows for thermal conductivity engineering. The electron conductivity also decreases due to enhanced boundary scattering by the constrictions. The effect of boundary scattering is weaker for electrons than for phonons and the overall thermoelectric efficiency is enhanced. A ZT enhancement by a factor of 20-30 is predicted for width-modulated nanowires compared to bulk silicon. Our findings indicate that width-modulated nanostructures are promising for developing silicon nanostructures with high thermoelectric efficiency.

  5. Proceedings of ICTEA 2006, the 2. international conference on thermal engineering : theory and applications

    International Nuclear Information System (INIS)

    Haik, Y; Saghir, Z.

    2006-01-01

    This international conference provided a venue for the exchange of research and the discussion of ideas related to thermal engineering. Participants at the conference discussed emerging research trends in thermal energy and presented new technologies and advances in computerized simulations and thermodynamic analyses related to thermal energy. Recent developments in solar cell technology, waste heat utilization, and energy management were presented. New developments in biomass combustion technologies were also described. The conference was divided into 22 sessions that discussed materials and polymers; computational fluid dynamics; energy management; solar energy; natural convection; experimental fluid flow; experimental combustion; multi-phase; environment; solar renewables; computational fluid dynamics and combustion; porous media; and micro and nano media. The conference featured 118 presentations, of which 63 have been catalogued separately for inclusion in this database. refs., tabs., figs

  6. THERMAL DISPLACEMENT OF CRANKSHAFT AXIS OF SLOW-SPEED MARINE ENGINE

    Directory of Open Access Journals (Sweden)

    Lech Murawski

    2016-08-01

    Full Text Available The paper presents analysis of displacement of a crankshaft axis caused by temperature of marine, slow-speed main engine. Information of thermal displacement of a power transmission system axis is significant during a shaft line alignment and a crankshaft springing analysis. Warmed-up main engine is a source of deformations of an engine body as well as a ship hull in the area of an engine room and hence axis of a crankshaft and a shaftline. Engines' producers recommend the model of parallel displacement of the crankshaft axis under the influence of an engine heat. The model gives us the value (one number! of the crankshaft axis displacement in the hot propulsion system's condition. This model may be too simple in some cases. Presented numerical analyses are based on temperature measurements of the main engine body and the ship hull during a sea voyage. The paper presents computations of MAN B&W K98MC type engine (power: 40000 kW, revolutions: 94 rpm mounted on 4500 TEU container ship (length: 290 m. Propulsion system is working in nominal, steady-state conditions; it is the basic assumption during the analyses. Numerical analyses were preformed with usage of Nastran software based on Finite Element Method. The FEM model of the engine body comprised over 800 thousand degree of freedom. Stiffness of the ship hull (mainly double bottom with the foundation was modelled by a simple cuboid. Material properties of that cuboid were determined on the base of separately performed calculations.

  7. Experimentally Studied Thermal Piston-head State of the Internal-Combustion Engine with a Thermal Layer Formed by Micro-Arc Oxidation Method

    Directory of Open Access Journals (Sweden)

    N. Yu. Dudareva

    2015-01-01

    Full Text Available The paper presents results of experimental study to show the efficiency of reducing thermal tension of internal combustion engine (ICE pistons through forming a thermal barrier coating on the piston-head. During the engine operation the piston is under the most thermal stress. High temperatures in the combustion chamber may lead to the piston-head burnout and destruction and engine failure.Micro-arc oxidation (MAO method was selected as the technology to create a thermal barrier coating. MAO technology allows us to form the ceramic coating with a thickness of 400μm on the surface of aluminum alloy, which have high heat resistance, and have good adhesion to the substrate even under thermal cycling stresses.Deliverables of MAO method used to protect pistons described in the scientific literature are insufficient, as they are either calculated or experimentally obtained at the special plants (units, which do not reproduce piston operation in a real engine. This work aims to fill this gap. The aim of the work is an experimental study of the thermal protective ability of MAO-layer formed on the piston-head with simulation of thermal processes of the real engine.The tests were performed on a specially designed and manufactured stand free of motor, which reproduces operation conditions maximum close to those of the real engine. The piston is heated by a fire source - gas burner with isobutene balloon, cooling is carried out by the water circulation system through the water-cooling jacket.Tests have been conducted to compare the thermal state of the regular engine piston without thermal protection and the piston with a heat layer formed on the piston-head by MAO method. The study findings show that the thermal protective MAO-layer with thickness of 100μm allows us to reduce thermal tension of piston on average by 8,5 %. Thus at high temperatures there is the most pronounced effect that is important for the uprated engines.The obtained findings can

  8. Experimental analysis on thermally coated diesel engine with neem oil methyl ester and its blends

    Science.gov (United States)

    Karthickeyan, V.

    2018-01-01

    Depletion of fossil fuel has created a threat to the nation's energy policy, which in turn led to the development of new source renewable of energy. Biodiesel was considered as the most promising alternative to the traditional fossil fuel. In the present study, raw neem oil was considered as a principle source for the production of biodiesel and converted into Neem Oil Methyl Ester (NOME) using two stage transesterification process. The chemical compositions of NOME was analysed using Fourier Transform Infra-Red Spectroscopy (FTIR) and Gas Chromatography- Mass Spectrometry (GC-MS). Baseline readings were recorded with diesel, 25NOME (25% NOME with 75% diesel) and 50NOME (50% NOME with 50% diesel) in a direct injection, four stroke, water cooled diesel engine. Thermal Barrier Coating (TBC) was considered as a better technique for emission reduction in direct injection diesel engine. In the present study, Partially Stabilized Zirconia (PSZ) was used as a TBC material to coat the combustion chamber components like cylinder head, piston head and intake and exhaust valves. In coated engine, 25NOME showed better brake thermal efficiency and declined brake specific fuel consumption than 50NOME. Decreased exhaust emissions like CO, HC and smoke were observed with 25NOME in coated engine except NOx. [Figure not available: see fulltext.

  9. Laser High-Cycle Thermal Fatigue of Pulse Detonation Engine Combustor Materials Tested

    Science.gov (United States)

    Zhu, Dong-Ming; Fox, Dennis S.; Miller, Robert A.

    2001-01-01

    Pulse detonation engines (PDE's) have received increasing attention for future aerospace propulsion applications. Because the PDE is designed for a high-frequency, intermittent detonation combustion process, extremely high gas temperatures and pressures can be realized under the nearly constant-volume combustion environment. The PDE's can potentially achieve higher thermodynamic cycle efficiency and thrust density in comparison to traditional constant-pressure combustion gas turbine engines (ref. 1). However, the development of these engines requires robust design of the engine components that must endure harsh detonation environments. In particular, the detonation combustor chamber, which is designed to sustain and confine the detonation combustion process, will experience high pressure and temperature pulses with very short durations (refs. 2 and 3). Therefore, it is of great importance to evaluate PDE combustor materials and components under simulated engine temperatures and stress conditions in the laboratory. In this study, a high-cycle thermal fatigue test rig was established at the NASA Glenn Research Center using a 1.5-kW CO2 laser. The high-power laser, operating in the pulsed mode, can be controlled at various pulse energy levels and waveform distributions. The enhanced laser pulses can be used to mimic the time-dependent temperature and pressure waves encountered in a pulsed detonation engine. Under the enhanced laser pulse condition, a maximum 7.5-kW peak power with a duration of approximately 0.1 to 0.2 msec (a spike) can be achieved, followed by a plateau region that has about one-fifth of the maximum power level with several milliseconds duration. The laser thermal fatigue rig has also been developed to adopt flat and rotating tubular specimen configurations for the simulated engine tests. More sophisticated laser optic systems can be used to simulate the spatial distributions of the temperature and shock waves in the engine. Pulse laser high

  10. Modeling of Thermal Cycle CI Engine with Multi-Stage Fuel Injection

    Directory of Open Access Journals (Sweden)

    Arkadiusz Jamrozik

    2017-09-01

    Full Text Available This work presents a complete thermal cycle modeling of a four-stroke diesel engine with a three-dimensional simulation program CFD - AVL Fire. The object of the simulation was the S320 Andoria engine. The purpose of the study was to determine the effect of fuel dose distribution on selected parameters of the combustion process. As a result of the modeling, time spatial pressure distributions, rate of pressure increase, heat release rate and NO and soot emission were obtained for 3 injection strategies: no division, one pilot dose and one main dose and two pilot doses and one main dose. It has been found that the use of pilot doses on the one hand reduces engine hardness and lowers NO emissions and on the other hand, increases soot emissions.

  11. The influence of thermal regime on gasoline direct injection engine performance and emissions

    Science.gov (United States)

    Leahu, C. I.; Tarulescu, S.

    2016-08-01

    This paper presents the experimental research regarding to the effects of a low thermal regime on fuel consumption and pollutant emissions from a gasoline direct injection (GDI) engine. During the experimental researches, the temperature of the coolant and oil used by the engine were modified 4 times (55, 65, 75 and 85 oC), monitoring the effects over the fuel consumption and emissions (CO2, CO and NOx). The variations in temperature of the coolant and oil have been achieved through AVL coolant and oil conditioning unit, integrated in the test bed. The obtained experimental results reveals the poor quality of exhaust gases and increases of fuel consumption for the gasoline direct injection engines that runs outside the optimal ranges for coolant and oil temperatures.

  12. Aircraft engine-mounted camera system for long wavelength infrared imaging of in-service thermal barrier coated turbine blades

    Science.gov (United States)

    Markham, James; Cosgrove, Joseph; Scire, James; Haldeman, Charles; Agoos, Ian

    2014-12-01

    This paper announces the implementation of a long wavelength infrared camera to obtain high-speed thermal images of an aircraft engine's in-service thermal barrier coated turbine blades. Long wavelength thermal images were captured of first-stage blades. The achieved temporal and spatial resolutions allowed for the identification of cooling-hole locations. The software and synchronization components of the system allowed for the selection of any blade on the turbine wheel, with tuning capability to image from leading edge to trailing edge. Its first application delivered calibrated thermal images as a function of turbine rotational speed at both steady state conditions and during engine transients. In advance of presenting these data for the purpose of understanding engine operation, this paper focuses on the components of the system, verification of high-speed synchronized operation, and the integration of the system with the commercial jet engine test bed.

  13. Stirling engines for low-temperature solar-thermal-electric power generation

    Science.gov (United States)

    der Minassians, Artin

    This dissertation discusses the design and development of a distributed solar-thermal-electric power generation system that combines solar-thermal technology with a moderate-temperature Stirling engine to generate electricity. The conceived system incorporates low-cost materials and utilizes simple manufacturing processes. This technology is expected to achieve manufacturing cost of less than $1/W. Since solar-thermal technology is mature, the analysis, design, and experimental assessment of moderate-temperature Stirling engines is the main focus of this thesis. The design, fabrication, and test of a single-phase free-piston Stirling engine prototype is discussed. This low-power prototype is designed and fabricated as a test rig to provide a clear understanding of the Stirling cycle operation, to identify the key components and the major causes of irreversibility, and to verify corresponding theoretical models. As a component, the design of a very low-loss resonant displacer piston subsystem is discussed. The displacer piston is part of a magnetic circuit that provides both a required stiffness and actuation forces. The stillness is provided by a magnetic spring, which incorporates an array of permanent magnets and has a very linear stiffness characteristic that facilitates the frequency tuning. In this prototype, the power piston is not mechanically linked to the displacer piston and forms a mass-spring resonating subsystem with the engine chamber gas spring and has resonant frequency matched to that of the displacer. The fabricated engine prototype is successfully tested and the experimental results are presented and discussed. Extensive experimentation on individual component subsystems confirms the theoretical models and design considerations, providing a sound basis for higher power Stirling engine designs for residential or commercial deployments. Multi-phase Stirling engine systems are also considered and analyzed. The modal analysis of these machines proves

  14. Characterization and comparative investigation of thermally insulating layers for the turbine and engine construction

    International Nuclear Information System (INIS)

    Steffens, H.D.; Fischer, U.

    1987-01-01

    The aim of the research project was to subject commercially produced thermal insulation layer systems, the use of which seems promising for engine and turbine construction, to standardized characterisation, testing and comparison. Suitable methods and procedures for this had to be developed, in order to be able to derive instructions for optimisation guidelines for the production of improved thermal insulation systems from the results of investigations. In the context of the research project, a computer-controlled thermal shock test rig was first developed, designed and built. This test rig was designed so that important test conditions, such as the heating and cooling speed could be varied reproducibly over wide ranges. Methods and procedures were worked out, which permit a comparative qualitative and quantitative characterisation of layers of thermal insulation. From metallographic investigations, the layer build-up, layer structure, porosity and crack morphology of the layers in the delivered state and after testing could be assessed and compared. X-ray fine structure investigations gave information on the type and quantity of the phases occurring in the ceramic layers. The results of thermal shock tests which were done at different temperature intervals depending on the substrate, could be correlated with the build-up of layers and supplied information on damage mechanisms and the course of failure. (orig.) With 57 figs., 16 tabs., 89 refs [de

  15. Regeneration in an internal combustion engine: Thermal-hydraulic modeling and analysis

    International Nuclear Information System (INIS)

    Thyageswaran, Sridhar

    2016-01-01

    Highlights: • An arrangement is proposed for in-cylinder regeneration in a 4-stroke engine. • Thermodynamic models are formulated for overall cycle analysis. • A design procedure is outlined for micro-channel regenerators. • Partial differential equations are solved for flow inside the regenerator. • Regeneration with lean combustion decreases the idealized cycle efficiency. - Abstract: An arrangement is proposed for a four-stroke internal combustion engine to: (a) recover thermal energy from products of combustion during the exhaust stroke; (b) store that energy as sensible heat in a micro-channel regenerator matrix; and (c) transfer the stored heat to compressed fresh charge that flows through the regenerator during the succeeding mechanical cycle. An extra moveable piston that can be locked at preferred positions and a sequence of valve events enable the regenerator to lose heat to the working fluid during one interval of time but gain heat from the fluid during another interval of time. This paper examines whether or not this scheme for in-cylinder regeneration (ICR) improves the cycle thermal efficiency η I . Models for various thermodynamic processes in the cycle and treatments for unsteady compressible flow and heat transfer inside the regenerator are developed. Digital simulations of the cycle are made. Compared to an idealized engine cycle devoid of regeneration, provisions for ICR seem to deteriorate the thermal efficiency. In an 8:1 compression ratio octane engine simulated with an equivalence ratio of 0.75, η I  = 0.455 with regeneration and η I  = 0.491 without. This study shows that previous claims on efficiency gains via ICR, using highly-simplified models, may be misleading.

  16. Definition and prediction of thermal state of cylinder assembly of an internal combustion engine: a new methodology.

    OpenAIRE

    ZEBBAR, Djallel; DOROKHOV, Alexandre Feodorovish; KOUIDER, Mostefa; ZEBBAR, Souhila; KHERRIS, Sahraoui

    2018-01-01

    Abstract This study deal with a new approach for the definition and prediction of the thermal state of internal combustion engine elements. In this paper, it is well elaborated for diesel engine cylinder liner. The scientific innovation of the new methodology consists in the use of data from external heat balance to determine the thermal state, without carrying out complex and expensive experiments. It takes into account the different heat transfer components: from gas and heat transfer d...

  17. The flight of uncontrolled rockets

    CERN Document Server

    Gantmakher, F R; Dryden, H L

    1964-01-01

    International Series of Monographs on Aeronautics and Astronautics, Division VII, Volume 5: The Flight of Uncontrolled Rockets focuses on external ballistics of uncontrolled rockets. The book first discusses the equations of motion of rockets. The rocket as a system of changing composition; application of solidification principle to rockets; rotational motion of rockets; and equations of motion of the center of mass of rockets are described. The text looks at the calculation of trajectory of rockets and the fundamentals of rocket dispersion. The selection further focuses on the dispersion of f

  18. A numerical simulation package for analysis of neutronics and thermal fluids of space nuclear power and propulsion systems

    International Nuclear Information System (INIS)

    Anghaie, S.; Feller, G.J.; Peery, S.D.; Parsley, R.C.

    1993-01-01

    A system of computer codes for engineering simulation and in-depth analysis of nuclear and thermal fluid design of nuclear thermal rockets is developed. The computational system includes a neutronic solver package, a thermal fluid solver package and a propellant and materials property package. The Rocket Engine Transient Simulation (ROCETS) system code is incorporated with computational modules specific to nuclear powered engines. ROCETS features a component based performance architecture that interfaces component modules into the user designed configuration, interprets user commands, creates an executable FORTRAN computer program, and executes the program to provide output to the user. Basic design features of the Pratt ampersand Whitney XNR2000 nuclear rocket concept and its operational performance are analyzed and simulated

  19. The Effects of Thermal Barrier Coatings on Diesel Engine Performance and Emission

    Science.gov (United States)

    Das, D.; Majumdar, G.; Sen, R. S.; Ghosh, B. B.

    2014-01-01

    The purpose of this paper is to determine the effect of coating thickness on performance and emission of a diesel engine including comparisons with results from an uncoated piston diesel engine. Primarily three piston crowns were coated with Al2O3 (bond coat) of 100 μm thickness each by using Plasma spray coating technique. Then these piston crowns were coated with partially stabilized zirconia with a thickness of 250, 350, 450 μm respectively by using the same technique over the bond coat. These pistons inserted into the cylinder of a diesel engine one by one to collect the combustion and emission data. Then these data were compared with standard diesel engine. It was observed that the thermal efficiency increased with increasing load levels, whereas specific fuel consumption reduced with increasing load. However, it was observed that harmful gases and particulates like CO, smoke and HC were reduced in case of all types of coated piston engine with the increase of load. Increased amount of NOX emission was reported during the experimentation.

  20. Thirteenth symposium on energy engineering sciences: Proceedings. Fluid/thermal processes, systems analysis and control

    Energy Technology Data Exchange (ETDEWEB)

    NONE

    1995-11-01

    The DOE Office of Basic Energy Sciences, of which Engineering Research is a component program, is responsible for the long-term mission-oriented research in the Department. Consistent with the DOE/BES mission, the Engineering Research Program is charged with the identification, initiation, and management of fundamental research on broad, generic topics addressing energy-related engineering problems. Its stated goals are: (1) to improve and extend the body of knowledge underlying current engineering practice so as to create new options for enhancing energy savings and production, for prolonging useful life of energy-related structures and equipment, and for developing advanced manufacturing technologies and materials processing with emphasis on reducing costs with improved industrial production and performance quality; and (2) to expand the store of fundamental concepts for solving anticipated and unforeseen engineering problems in the energy technologies. The meeting covered the following areas: (1) fluid mechanics 1--fundamental properties; (2) fluid mechanics 2--two phase flow; (3) thermal processes; (4) fluid mechanics 3; (5) process analysis and control; (6) fluid mechanics 4--turbulence; (7) fluid mechanics 5--chaos; (8) materials issues; and (9) plasma processes. Selected papers are indexed separately for inclusion in the Energy Science and Technology Database.

  1. Design and Performance Optimizations of Advanced Erosion-Resistant Low Conductivity Thermal Barrier Coatings for Rotorcraft Engines

    Science.gov (United States)

    Zhu, Dongming; Miller, Robert A.; Kuczmarski, Maria A.

    2012-01-01

    Thermal barrier coatings will be more aggressively designed to protect gas turbine engine hot-section components in order to meet future rotorcraft engine higher fuel efficiency and lower emission goals. For thermal barrier coatings designed for rotorcraft turbine airfoil applications, further improved erosion and impact resistance are crucial for engine performance and durability, because the rotorcraft are often operated in the most severe sand erosive environments. Advanced low thermal conductivity and erosion-resistant thermal barrier coatings are being developed, with the current emphasis being placed on thermal barrier coating toughness improvements using multicomponent alloying and processing optimization approaches. The performance of the advanced thermal barrier coatings has been evaluated in a high temperature erosion burner rig and a laser heat-flux rig to simulate engine erosion and thermal gradient environments. The results have shown that the coating composition and architecture optimizations can effectively improve the erosion and impact resistance of the coating systems, while maintaining low thermal conductivity and cyclic oxidation durability

  2. Harvesting of PEM fuel cell heat energy for a thermal engine in an underwater glider

    Energy Technology Data Exchange (ETDEWEB)

    Wang, Shuxin; Xie, Chungang; Wang, Yanhui; Zhang, Lianhong; Jie, Weiping [School of Mechanical Engineering, Tianjin University, Tianjin 300072 (China); Hu, S. Jack [Department of Mechanical Engineering, The University of Michigan, Ann Arbor, MI 48109-2125 (United States)

    2007-06-20

    The heat generated by a proton exchange membrane fuel cell (PEMFC) is generally removed from the cell by a cooling system. Combining heat energy and electricity in a PEMFC is highly desirable to achieve higher fuel efficiency. This paper describes the design of a new power system that combines the heat energy and electricity in a miniature PEMFC to improve the overall power efficiency in an underwater glider. The system makes use of the available heat energy for navigational power of the underwater glider while the electricity generated by the miniature PEMFC is used for the glider's sensors and control system. Experimental results show that the performance of the thermal engine can be obviously improved due to the high quality heat from the PEMFC compared with the ocean environmental thermal energy. Moreover, the overall fuel efficiency can be increased from 17 to 25% at different electric power levels by harvesting the PEMFC heat energy for an integrated fuel cell and thermal engine system in the underwater glider. (author)

  3. A Robust Model Predictive Control for efficient thermal management of internal combustion engines

    International Nuclear Information System (INIS)

    Pizzonia, Francesco; Castiglione, Teresa; Bova, Sergio

    2016-01-01

    Highlights: • A Robust Model Predictive Control for ICE thermal management was developed. • The proposed control is effective in decreasing the warm-up time. • The control system reduces coolant flow rate under fully warmed conditions. • The control strategy operates the cooling system around onset of nucleate boiling. • Little on-line computational effort is required. - Abstract: Optimal thermal management of modern internal combustion engines (ICE) is one of the key factors for reducing fuel consumption and CO 2 emissions. These are measured by using standardized driving cycles, like the New European Driving Cycle (NEDC), during which the engine does not reach thermal steady state; engine efficiency and emissions are therefore penalized. Several techniques for improving ICE thermal efficiency were proposed, which range from the use of empirical look-up tables to pulsed pump operation. A systematic approach to the problem is however still missing and this paper aims to bridge this gap. The paper proposes a Robust Model Predictive Control of the coolant flow rate, which makes use of a zero-dimensional model of the cooling system of an ICE. The control methodology incorporates explicitly the model uncertainties and achieves the synthesis of a state-feedback control law that minimizes the “worst case” objective function while taking into account the system constraints, as proposed by Kothare et al. (1996). The proposed control strategy is to adjust the coolant flow rate by means of an electric pump, in order to bring the cooling system to operate around the onset of nucleate boiling: across it during warm-up and above it (nucleate or saturated boiling) under fully warmed conditions. The computationally heavy optimization is carried out off-line, while during the operation of the engine the control parameters are simply picked-up on-line from look-up tables. Owing to the little computational effort required, the resulting control strategy is suitable for

  4. Not just rocket science

    Energy Technology Data Exchange (ETDEWEB)

    MacAdam, S.; Anderson, R. [Celan Energy Systems, Rancho Cordova, CA (United States)

    2007-10-15

    The paper explains a different take on oxyfuel combustion. Clean Energy Systems (CES) has integrated aerospace technology into conventional power systems, creating a zero-emission power generation technology that has some advantages over other similar approaches. When using coal as a feedstock, the CES process burns syngas rather than raw coal. The process uses recycled water and steam to moderate the temperature, instead of recycled CO{sub 2}. With no air ingress, the CES process produces very pure CO{sub 2}. This makes it possible to capture over 99% of the CO{sub 2} resulting from combustion. CES uses the combustion products to drive the turbines, rather than indirectly raising steam for steam turbines, as in the oxyfuel process used by companies such as Vattenfall. The core of the process is a high-pressure oxy-combustor adapted from rocket engine technology. This combustor burns gaseous or liquid fuels with gaseous oxygen in the presence of water. Fuels include natural gas, coal or coke-derived synthesis gas, landfill and biodigester gases, glycerine solutions and oil/water emulsion. 2 figs.

  5. Collaboration with and without Coauthorship: Rocket Science Versus Economic Science

    OpenAIRE

    Barnett, William

    2015-01-01

    This essay is about my prior experiences as a rocket scientist on Apollo rocket engines, with comparison to my subsequent experiences at the Federal Reserve, and in academia, with emphasis upon differences in collaboration and scientific methodology. A primary difference is in the emphasis on measurement.

  6. Concept study of a hydrogen containment process during nuclear thermal engine ground testing

    Directory of Open Access Journals (Sweden)

    Ten-See Wang

    Full Text Available A new hydrogen containment process was proposed for ground testing of a nuclear thermal engine. It utilizes two thermophysical steps to contain the hydrogen exhaust. First, the decomposition of hydrogen through oxygen-rich combustion at higher temperature; second, the recombination of remaining hydrogen with radicals at low temperature. This is achieved with two unit operations: an oxygen-rich burner and a tubular heat exchanger. A computational fluid dynamics methodology was used to analyze the entire process on a three-dimensional domain. The computed flammability at the exit of the heat exchanger was less than the lower flammability limit, confirming the hydrogen containment capability of the proposed process. Keywords: Hydrogen decomposition reactions, Hydrogen recombination reactions, Hydrogen containment process, Nuclear thermal propulsion, Ground testing

  7. Defect-Engineered Heat Transport in Graphene: A Route to High Efficient Thermal Rectification

    Science.gov (United States)

    Zhao, Weiwei; Wang, Yanlei; Wu, Zhangting; Wang, Wenhui; Bi, Kedong; Liang, Zheng; Yang, Juekuan; Chen, Yunfei; Xu, Zhiping; Ni, Zhenhua

    2015-07-01

    Low-dimensional materials such as graphene provide an ideal platform to probe the correlation between thermal transport and lattice defects, which could be engineered at the molecular level. In this work, we perform molecular dynamics simulations and non-contact optothermal Raman measurements to study this correlation. We find that oxygen plasma treatment could reduce the thermal conductivity of graphene significantly even at extremely low defect concentration (˜83% reduction for ˜0.1% defects), which could be attributed mainly to the creation of carbonyl pair defects. Other types of defects such as hydroxyl, epoxy groups and nano-holes demonstrate much weaker effects on the reduction where the sp2 nature of graphene is better preserved. With the capability of selectively functionalizing graphene, we propose an asymmetric junction between graphene and defective graphene with a high thermal rectification ratio of ˜46%, as demonstrated by our molecular dynamics simulation results. Our findings provide fundamental insights into the physics of thermal transport in defective graphene, and two-dimensional materials in general, which could help on the future design of functional applications such as optothermal and electrothermal devices.

  8. Thermal performance test of hot gas ducts of helium engineering demonstration loop (HENDEL)

    International Nuclear Information System (INIS)

    Hishida, Makoto; Kunitomi, Kazuhiko; Ioka, Ikuo; Umenishi, Koji; Kondo, Yasuo; Tanaka, Toshiyuki; Shimomura, Hiroaki

    1984-01-01

    A hot gas duct provided with internal thermal insulation is supposed to be used for an experimental very high-temperature gas-cooled reactor (VHTR) which has been developed by the Japan Atomic Energy Research Institute (JAERI). This type of hot gas duct has not been used so far in industrial facilities, and only a couple of tests on such a large-scale model of hot gas duct have been conducted. The present test was to investigate the thermal performance of the hot gas ducts which are installed as parts of a helium engineering demonstration loop (HENDEL) of JAERI. Uniform temperature and heat flux distributions at the surface of the duct were observed, the experimental correlation being obtained for the effective thermal conductivity of the internal thermal insulation layer. The measured temperature distribution of the pressure tube was in good agreement with the calculation by a TRUMP heat transfer computer code. The temperature distribution of the inner tube of VHTR hot gas duct was evaluated, and no hot spot was detected. These results would be very valuable for the design and development of VHTR. (author)

  9. Molecularly Engineered Azobenzene Derivatives for High Energy Density Solid-State Solar Thermal Fuels.

    Science.gov (United States)

    Cho, Eugene N; Zhitomirsky, David; Han, Grace G D; Liu, Yun; Grossman, Jeffrey C

    2017-03-15

    Solar thermal fuels (STFs) harvest and store solar energy in a closed cycle system through conformational change of molecules and can release the energy in the form of heat on demand. With the aim of developing tunable and optimized STFs for solid-state applications, we designed three azobenzene derivatives functionalized with bulky aromatic groups (phenyl, biphenyl, and tert-butyl phenyl groups). In contrast to pristine azobenzene, which crystallizes and makes nonuniform films, the bulky azobenzene derivatives formed uniform amorphous films that can be charged and discharged with light and heat for many cycles. Thermal stability of the films, a critical metric for thermally triggerable STFs, was greatly increased by the bulky functionalization (up to 180 °C), and we were able to achieve record high energy density of 135 J/g for solid-state STFs, over a 30% improvement compared to previous solid-state reports. Furthermore, the chargeability in the solid state was improved, up to 80% charged from 40% charged in previous solid-state reports. Our results point toward molecular engineering as an effective method to increase energy storage in STFs, improve chargeability, and improve the thermal stability of the thin film.

  10. A new closed-form analytical thermal model for simulating Stirling engines based on polytropic-finite speed thermodynamics

    International Nuclear Information System (INIS)

    Hosseinzade, Hadi; Sayyaadi, Hoseyn; Babaelahi, Mojtaba

    2015-01-01

    Highlights: • A closed-form thermal model was presented for Stirling engines. • The new model was used to simulate the GPU-3 Stirling engine. • Results were compared with experimental data as well as other models. • The new model was more accurate and simple in calculation than other models. • Effects of the engines’ parameters on operation of engine were evaluated. - Abstract: Thermal models for the simulation of Stirling engines need to have greater accuracy along with simple and low-cost calculation. In this regard, a new closed-form thermal model was presented for the thermal simulation of Stirling engines. The new model called PFST (polytropic-finite speed thermodynamics) was developed based on the combination of polytropic analysis of expansion/compression processes and the concept of finite speed thermodynamics (FST). Therefore, compression/expansion works of compression/expansion processes and transferred heat into the heater of Stirling engines were determined based on polytropic analysis, instead of isothermal processes of the ideal Stirling cycle. The calculated work of polytropic processes was corrected to include the effects of internal irreversibilities including pressure throttling in heat exchangers, mechanical friction, and finite motion of the pistons. Output power and thermal efficiency of Stirling engines were calculated as functions of various engine parameters. The developed PFST model was implemented on a prototype Stirling engine, called GPU-3 engine, and the obtained results were compared with those of other closed-form and numerical models as well as experimental data. It was found that the new closed-form model, in addition to its simple and low-cost calculation, had the same order of accuracy as recently developed numerical models

  11. The 25 kWe solar thermal Stirling hydraulic engine system: Conceptual design

    Science.gov (United States)

    White, Maurice; Emigh, Grant; Noble, Jack; Riggle, Peter; Sorenson, Torvald

    1988-01-01

    The conceptual design and analysis of a solar thermal free-piston Stirling hydraulic engine system designed to deliver 25 kWe when coupled to a 11 meter test bed concentrator is documented. A manufacturing cost assessment for 10,000 units per year was made. The design meets all program objectives including a 60,000 hr design life, dynamic balancing, fully automated control, more than 33.3 percent overall system efficiency, properly conditioned power, maximum utilization of annualized insolation, and projected production costs. The system incorporates a simple, rugged, reliable pool boiler reflux heat pipe to transfer heat from the solar receiver to the Stirling engine. The free-piston engine produces high pressure hydraulic flow which powers a commercial hydraulic motor that, in turn, drives a commercial rotary induction generator. The Stirling hydraulic engine uses hermetic bellows seals to separate helium working gas from hydraulic fluid which provides hydrodynamic lubrication to all moving parts. Maximum utilization of highly refined, field proven commercial components for electric power generation minimizes development cost and risk.

  12. Thermal efficiency and environmental performances of a biogas-diesel stationary engine.

    Science.gov (United States)

    Bilcan, A; Le Corre, O; Delebarre, A

    2003-09-01

    Municipal and agricultural waste, and sludge from wastewater treatment represent a large source of pollution. Gaseous fuels can be produced from waste decomposition and then used to run internal combustion engines for power and heat generation. The present paper focuses on thermal efficiency and environmental performances of dual-fuel engines fuelled with biogas. Experiments have been carried out on a Lister-Petter single cylinder diesel engine, modified for dual-fuel operation. Natural gas was first used as the primary fuel. An empirical correlation was determined to predict the engine load for a given mass flow rate for the pilot fuel (diesel) and for the primary fuel (natural gas). That correlation has then been tested for three synthesized biogas compositions. Computations were performed and the error was estimated to be less than 10%. Additionally, NOx and CO2 contents were measured from exhaust gases. Based on exhausts gas temperature, the activation energy and the pre-exponential factor of an Arrhenius law were then proposed, resulting in a simpler mean to predict NOx.

  13. ELIMINATION OF ROCKET IGNITION SIDE LOADS, Phase I

    Data.gov (United States)

    National Aeronautics and Space Administration — This proposal is responsive to Topic H10: Ground Processing and in particular to Subtopic H10.02. When a rocket motor/engine is ignited at low altitude its...

  14. Gaseous Helium Reclamation at Rocket Test Systems, Phase II

    Data.gov (United States)

    National Aeronautics and Space Administration — GHe reclamation is critical in reducing operating costs at rocket engine test facilities. Increases in cost and shortages of helium will dramatically impact testing...

  15. Nitrous Ethane-Ethylene Rocket with Hypergolic Ignition, Phase I

    Data.gov (United States)

    National Aeronautics and Space Administration — The Nitrous Ethane-Ethylene Rocket with Hypergolic Ignition (NEERHI) engine is a proposed technology designed to provide small spacecraft with non-toxic,...

  16. Nitrous Ethane-Ethylene Rocket with Hypergolic Ignition, Phase II

    Data.gov (United States)

    National Aeronautics and Space Administration — The Nitrous Ethane-Ethylene Rocket with Hypergolic Ignition (NEERHI) engine is a proposed technology designed to provide small spacecraft with non-toxic,...

  17. Advanced small rocket chambers. Option 3: 110 1bf Ir-Re rocket, volume 2

    Science.gov (United States)

    Jassowski, Donald M.; Schoenman, Leonard

    1995-02-01

    This is the second part of a two-part report that describes the AJ10-221, a high performance iridium-coated rhenium (Ir-Re) 110 lbf (490N) welded rocket chamber with 286:1 area ratio nozzle. This engine was designed, built, and hot fired for over 6 hours on this program. It demonstrated an I(s) of 321.8 sec, which is 10 sec higher than conventional 110 lbf silicide coated Cb chambers now in use. The approach used in this portion of the program was to demonstrate the performance improvement that can be made by the elimination of fuel film cooling and the use of a high temperature (4000 F) (2200 C) iridium-coated rhenium (Ir-Re) rocket chamber. Detailed thermal, performance, mechanical, and dynamic design analyses of the full engine were conducted by Aerojet. Two Ir-Re chambers were built to the Aerojet design by Ultramet, using the chemical vapor deposition (CVD) process. Incorporation of a secondary mixing device or Boundary Layer Trip (BLT) within the combustion chamber (Aerojet Patents 4882904 and 4936091) results in improvement in flow uniformity, and a significant life and performance increase. The 110 lbf engine design was verified in bolt-up hardware tests at sea level and altitude. The effects of injector design on performance were studied. Two duplicate injectors were fabricated matching the preferred design and were demonstrated to be interchangeable in operation. One of these units were welded into a flight type thruster which was tested for an accumulated duration of 22,590 sec in 93 firings, one of which was a continuous burn of two hours. A design deficiency in the C-103 nozzle near the Re-Cb transition joint was discovered, studied and corrected design has been prepared. Workhardening studies have been conducted to investigate methods for increasing the low yield strength of the Re in the annealed conditions. An advanced 490N high performance engine has been demonstrated which, when proven to be capable of withstanding launch vibration, is ready for

  18. Advanced small rocket chambers. Option 3: 110 1bf Ir-Re rocket, volume 1

    Science.gov (United States)

    Jassowski, Donald M.; Schoenman, Leonard

    1995-02-01

    This report describes the AJ10-221, a high performance Iridium-coated Rhenium (Ir-Re) 110 lbf (490N) welded rocket chamber with 286:1 area ratio nozzle. This engine was designed, built, and hot fired for over 6 hours on this program. It demonstrated an I(s) of 321.8 sec, which is 10 sec higher than conventional 110 lbf silicide coated Cb chambers now in use. The approach used in this portion of the program was to demonstrate the performance improvement that can be made by the elimination of fuel film cooling and the use of a high temperature (4000F) (2200C) iridium-coated rhenium (Ir-Re) rocket chamber. Detailed thermal, performance, mechanical, and dynamic design analyses of the full engine were conducted by Aerojet. Two Ir-Re chambers were built to the Aerojet design by Ultramet, using the chemical vapor deposition (CVD) process. Incorporation of a secondary mixing device or Boundary Layer Trip (BLT) within the combustion chamber (Aerojet Patents 4882904 and 4936091) results in improvement in flow uniformity, and a significant life and performance increase. The 110 lbf engine design was verified in bolt-up hardware tests at sea level and altitude. The effects of injector design on performance were studied. Two duplicate injectors were fabricated matching the preferred design and were demonstrated to be interchangeable in operation. One of these units was fabricated matching the preferred design and was demonstrated to be interchangeable in operation. One of these units was welded into a flight type thruster which was tested for an accumulated duration of 22,590 sec in 93 firings, one of which was a continuous burn of two hours. A design deficiency in the C-103 nozzle near the Re-Cb transition joint was discovered, studied and corrected design has been prepared. Workhardening studies have been conducted to investigate methods for increasing the low yield strength of the Re in the annealed conditions. An advanced 490N high performance engine has been demonstrated

  19. Engineering structure and thermal-technical analysis of fusion experimental breeder FEB divertor

    International Nuclear Information System (INIS)

    Feng Kaiming; Huang Jinhua; Zhu Yukun; Deng Peizhi; Zhou Xiaobing; Wang Min; Huo Tiejun

    1999-10-01

    On the basis of the physical study of FEB divertor, the engineering structure and thermal-technical analysis of FEB divertor are presented. In order to improve the impurity control and to increase ion-neutral interactions in the divertor, the configuration of the divertor is optimized to be the close type in the engineering design activity compared with the open type in the early conceptual activity. The operation mode of the divertor is designed to be partial detached plasma mode under conditions of combination gas-puffing with impurity injection. The position of gas-puffing is optimized at the torus mid-plane with NEWT1D code from the view point of impurity retention and radiation in the scrape-off layer/divertor region. The divertor structure is consisted of 48 rounded cassette modules. The thermal-technical calculations are carried out with COSMOS/M-HSTAR code for target plates. The result showed that the He-cooled target with 4 MPa coolant pressure and radial flowing is feasible

  20. Investigation of the effect of FeCl3 on combustion and emission of diesel engine with thermal barrier coating

    Directory of Open Access Journals (Sweden)

    Shakti P. Jena

    2018-03-01

    Full Text Available In the present investigation, the engine performance and emission characteristics of a single cylinder diesel engine with yttria stabilized zirconia (YSZ coating on piston crown and valves were studied. The 0.2 g L−1 of ferric chloride (FeCl3 as catalyst was added into the diesel fuel in both coated and uncoated engines. The results indicated that FeCl3 with diesel in a YSZ coated engine increased the brake thermal efficiency by 2.7%, and reduced brake specific fuel consumption by 8.3% as compared to standard diesel mode in uncoated engine. The selected thermal barrier coating improved the combustion in afterburning stage leading to effective use of intake air. Emissions such as carbon monoxide, hydrocarbons and smoke opacity were reduced with an increase in emissions of nitrogen oxide and carbon dioxide.

  1. Design study of RL10 derivatives. Volume 3, part 2: Operational and flight support plan. [analysis of transportation requirements for rocket engine in support of space tug program

    Science.gov (United States)

    Shubert, W. C.

    1973-01-01

    Transportation requirements are considered during the engine design layout reviews and maintenance engineering analyses. Where designs cannot be influenced to avoid transportation problems, the transportation representative is advised of the problems permitting remedies early in the program. The transportation representative will monitor and be involved in the shipment of development engine and GSE hardware between FRDC and vehicle manufacturing plant and thereby will be provided an early evaluation of the transportation plans, methods and procedures to be used in the space tug support program. Unanticipated problems discovered in the shipment of development hardware will be known early enough to permit changes in packaging designs and transportation plans before the start of production hardware and engine shipments. All conventional transport media can be used for the movement of space tug engines. However, truck transport is recommended for ready availability, variety of routes, short transit time, and low cost.

  2. Effect of thermal barrier coating with various blends of pumpkin seed oil methyl ester in DI diesel engine

    Science.gov (United States)

    Karthickeyan, V.; Balamurugan, P.

    2017-10-01

    The rise in oil prices, dependency on fossil fuels, degradation of non-renewable energy resources and global warming strives to find a low-carbon content alternative fuel to the conventional fuel. In the present work, Partially Stabilized Zirconia (PSZ) was used as a thermal barrier coating in piston head, cylinder head and intake and exhaust valves using plasma spray technique, which provided a rise in combustion chamber temperature. With the present study, the effects of thermal barrier coating on the blends of Pumpkin Seed Oil Methyl Ester (PSOME) were observed in both the coated and uncoated engine. Performance and emission characteristics of the PSOME in coated and uncoated engines were observed and compared. Increased thermal efficiency and reduced fuel consumption were observed for B25 and diesel in coated and uncoated engine. On comparing with the other biodiesel samples, B25 exhibited lower HC, NOx and smoke emissions in thermally coated engine than uncoated engine. After 100 h of operation, no anamolies were found in the thermally coated components except minor cracks were identified in the edges of the piston head.

  3. Thermal design of a natural gas - diesel dual fuel turbocharged V18 engine for ship propulsion and power plant applications

    Science.gov (United States)

    Douvartzides, S.; Karmalis, I.

    2016-11-01

    A detailed method is presented on the thermal design of a natural gas - diesel dual fuel internal combustion engine. An 18 cylinder four stroke turbocharged engine is considered to operate at a maximum speed of 500 rpm for marine and power plant applications. Thermodynamic, heat transfer and fluid flow phenomena are mathematically analyzed to provide a real cycle analysis together with a complete set of calculated operation conditions, power characteristics and engine efficiencies. The method is found to provide results in close agreement to published data for the actual performance of similar engines such as V18 MAN 51/60DF.

  4. Microelectronic Spare and Repair Part Status Analysis for the Multiple Launch Rocket System (MLRS)

    National Research Council Canada - National Science Library

    Maddux, Gary

    1999-01-01

    .... IOD required management and engineering support In performing microelectronic technology and availability assessments for the impact of nonavailability on the Multiple Launch Rocket System (MLRS...

  5. BNFL's application of computer aided engineering to 'THORP' thermal oxide reprocessing plant

    International Nuclear Information System (INIS)

    Hall-Wilton, M.J.

    1990-01-01

    BNFL are currently constructing facilities at Sellafield, Cumbria to reprocess thermal oxide fuel for U.K., European and Japanese utilities. Faced with a 3.5bn pound capital program to provide new facilities at Sellafield, BNFL took the opportunity to embrace the new computer aided engineering technologies then emerging in 1981. To give some idea of the commitment made by BNFL to the above many millions of pounds has been invested in hardware, and more on software and people. The 'THORP' (Head End and Chemical Separation Plant) project represents 1.5bn pounds. The introduction of computer aided engineering has provided a clash free design with full assurance that all materials and components used are compatible. Planning in the design offices in conjunction with an experienced construction management team enables the sequence of piping installation to be determined long before the construction teams enter the work area. This planning aspect has been significantly improved by using EVS (Enhanced Visualisation System). The use of supercomputing graphics facilities is stimulating demands from areas of engineering who have previously not sought to use magnetic data from a variety of sources. The result is mainly due to the data being easily accessible to users who have very little computing experience. (N.K.)

  6. Characterization of Thermal Stability of Synthetic and Semi-Synthetic Engine Oils

    Directory of Open Access Journals (Sweden)

    Anand Kumar Tripathi

    2015-03-01

    Full Text Available Engine oils undergo oxidative degradation and wears out during service. Hence it is important to characterize ageing of engine oils at different simulated conditions to evaluate the performance of existing oils and also design new formulations. This work focuses on characterizing the thermo-oxidative degradation of synthetic and semi-synthetic engine oils aged at 120, 149 and 200 °C. Apparent activation energy of decomposition of aged oils evaluated using the isoconversional Kissinger-Akahira-Sunose technique was used as a thermal stability marker. The temporal variation of stability at different ageing temperatures was corroborated with kinematic viscosity, oxidation, sulfation and nitration indices, total base number, antiwear additive content and molecular structure of the organic species present in the oils. At the lowest temperature employed, synthetic oil underwent higher rate of oxidation, while semi-synthetic oil was stable for longer time periods. At higher temperatures, the initial rate of change of average apparent activation energy of synthetic oil correlated well with a similar variation in oxidation number. A mixture of long chain linear, branched, and cyclic hydrocarbons were observed when semi-synthetic oil was degraded at higher temperatures.

  7. Combining mechanical foaming and thermally induced phase separation to generate chitosan scaffolds for soft tissue engineering.

    Science.gov (United States)

    Biswas, D P; Tran, P A; Tallon, C; O'Connor, A J

    2017-02-01

    In this paper, a novel foaming methodology consisting of turbulent mixing and thermally induced phase separation (TIPS) was used to generate scaffolds for tissue engineering. Air bubbles were mechanically introduced into a chitosan solution which forms the continuous polymer/liquid phase in the foam created. The air bubbles entrained in the foam act as a template for the macroporous architecture of the final scaffolds. Wet foams were crosslinked via glutaraldehyde and frozen at -20 °C to induce TIPS in order to limit film drainage, bubble coalescence and Ostwald ripening. The effects of production parameters, including mixing speed, surfactant concentration and chitosan concentration, on foaming are explored. Using this method, hydrogel scaffolds were successfully produced with up to 80% porosity, average pore sizes of 120 μm and readily tuneable compressive modulus in the range of 2.6 to 25 kPa relevant to soft tissue engineering applications. These scaffolds supported 3T3 fibroblast cell proliferation and penetration and therefore show significant potential for application in soft tissue engineering.

  8. Pulse Phase Dynamic Thermal Tomography Investigation on the Defects of the Solid-Propellant Missile Engine Cladding Layer

    Science.gov (United States)

    Peng, Wei; Wang, Fei; Liu, Jun-yan; Xiao, Peng; Wang, Yang; Dai, Jing-min

    2018-04-01

    Pulse phase dynamic thermal tomography (PP-DTT) was introduced as a nondestructive inspection technique to detect the defects of the solid-propellant missile engine cladding layer. One-dimensional thermal wave mathematical model stimulated by pulse signal was developed and employed to investigate the thermal wave transmission characteristics. The pulse phase algorithm was used to extract the thermal wave characteristic of thermal radiation. Depth calibration curve was obtained by fuzzy c-means algorithm. Moreover, PP-DTT, a depth-resolved photothermal imaging modality, was employed to enable three-dimensional (3D) visualization of cladding layer defects. The comparison experiment between PP-DTT and classical dynamic thermal tomography was investigated. The results showed that PP-DTT can reconstruct the 3D topography of defects in a high quality.

  9. Another Look at Rocket Thrust

    Science.gov (United States)

    Hester, Brooke; Burris, Jennifer

    2012-01-01

    Rocket propulsion is often introduced as an example of Newton's third law. The rocket exerts a force on the exhaust gas being ejected; the gas exerts an equal and opposite force--the thrust--on the rocket. Equivalently, in the absence of a net external force, the total momentum of the system, rocket plus ejected gas, remains constant. The law of…

  10. Non-linear thermal engineering, chaotic advection and mixing; Thermique non-lineaire, melange et advection chaotique

    Energy Technology Data Exchange (ETDEWEB)

    NONE

    1998-12-31

    This conference day was jointly organized by the `university group of thermal engineering (GUT)` and the French association of thermal engineers. This book of proceedings contains 7 papers entitled: `energy spectra of a passive scalar undergoing advection by a chaotic flow`; `analysis of chaotic behaviours: from topological characterization to modeling`; `temperature homogeneity by Lagrangian chaos in a direct current flow heat exchanger: numerical approach`; ` thermal instabilities in a mixed convection phenomenon: nonlinear dynamics`; `experimental characterization study of the 3-D Lagrangian chaos by thermal analogy`; `influence of coherent structures on the mixing of a passive scalar`; `evaluation of the performance index of a chaotic advection effect heat exchanger for a wide range of Reynolds numbers`. (J.S.)

  11. Biomimetic thermal barrier coating in jet engine to resist volcanic ash deposition

    Science.gov (United States)

    Song, Wenjia; Major, Zsuzsanna; Schulz, Uwe; Muth, Tobias; Lavallée, Yan; Hess, Kai-Uwe; Dingwell, Donald B.

    2017-04-01

    The threat of volcanic ash to aviation safety is attracting extensive attention when several commercial jet aircraft were damaged after flying through volcanic ash clouds from the May 1980 eruptions of Mount St. Helen in Washington, U.S. and especially after the air traffic disruption in 2010 Eyjafjallajökull eruption. A major hazard presented by volcanic ash to aircraft is linked to the wetting and spreading of molten ash droplets on engine component surfaces. Due to the fact ash has a lower melting point, around 1100 °C, than the gas temperature in the hot section (between 1400 to 2000 °C), this cause the ash to melt and potentially stick to the internal components (e.g., combustor and turbine blades), this cause the ash to melt and potentially stick to the internal components of the engine creating, substantial damage or even engine failure after ingestion. Here, inspiring form the natural surface of lotus leaf (exhibiting extreme water repellency, known as 'lotus effect'), we firstly create the multifunctional surface thermal barrier coatings (TBCs) by producing a hierarchical structure with femtosecond laser pulses. In detail, we investigate the effect of one of primary femtosecond laser irradiation process parameter (scanning speed) on the hydrophobicity of water droplets onto the two kinds of TBCs fabricated by electron-beam physical vapor deposition (EB-PVD) and air plasma spray (APS), respectively as well as their corresponding to morphology. It is found that, comparison with the original surface (without femtosecond laser ablation), all of the irradiated samples demonstrate more significant hydrophobic properties due to nanostructuring. On the basis of these preliminary room-temperature results, the wettability of volcanic ash droplets will be analysed at the high temperature to constrain the potential impact of volcanic ash on the jet engines.

  12. Evaluation by Rocket Combustor of C/C Composite Cooled Structure Using Metallic Cooling Tubes

    Science.gov (United States)

    Takegoshi, Masao; Ono, Fumiei; Ueda, Shuichi; Saito, Toshihito; Hayasaka, Osamu

    In this study, the cooling performance of a C/C composite material structure with metallic cooling tubes fixed by elastic force without chemical bonding was evaluated experimentally using combustion gas in a rocket combustor. The C/C composite chamber was covered by a stainless steel outer shell to maintain its airtightness. Gaseous hydrogen as a fuel and gaseous oxygen as an oxidizer were used for the heating test. The surface of these C/C composites was maintained below 1500 K when the combustion gas temperature was about 2800 K and the heat flux to the combustion chamber wall was about 9 MW/m2. No thermal damage was observed on the stainless steel tubes that were in contact with the C/C composite materials. The results of the heating test showed that such a metallic tube-cooled C/C composite structure is able to control the surface temperature as a cooling structure (also as a heat exchanger) as well as indicated the possibility of reducing the amount of coolant even if the thermal load to the engine is high. Thus, application of this metallic tube-cooled C/C composite structure to reusable engines such as a rocket-ramjet combined-cycle engine is expected.

  13. US Rocket Propulsion Industrial Base Health Metrics

    Science.gov (United States)

    Doreswamy, Rajiv

    2013-01-01

    The number of active liquid rocket engine and solid rocket motor development programs has severely declined since the "space race" of the 1950s and 1960s center dot This downward trend has been exacerbated by the retirement of the Space Shuttle, transition from the Constellation Program to the Space launch System (SLS) and similar activity in DoD programs center dot In addition with consolidation in the industry, the rocket propulsion industrial base is under stress. To Improve the "health" of the RPIB, we need to understand - The current condition of the RPIB - How this compares to past history - The trend of RPIB health center dot This drives the need for a concise set of "metrics" - Analogous to the basic data a physician uses to determine the state of health of his patients - Easy to measure and collect - The trend is often more useful than the actual data point - Can be used to focus on problem areas and develop preventative measures The nation's capability to conceive, design, develop, manufacture, test, and support missions using liquid rocket engines and solid rocket motors that are critical to its national security, economic health and growth, and future scientific needs. center dot The RPIB encompasses US government, academic, and commercial (including industry primes and their supplier base) research, development, test, evaluation, and manufacturing capabilities and facilities. center dot The RPIB includes the skilled workforce, related intellectual property, engineering and support services, and supply chain operations and management. This definition touches the five main segments of the U.S. RPIB as categorized by the USG: defense, intelligence community, civil government, academia, and commercial sector. The nation's capability to conceive, design, develop, manufacture, test, and support missions using liquid rocket engines and solid rocket motors that are critical to its national security, economic health and growth, and future scientific needs

  14. An historical collection of papers on nuclear thermal propulsion

    Science.gov (United States)

    The present volume of historical papers on nuclear thermal propulsion (NTP) encompasses NTP technology development regarding solid-core NTP technology, advanced concepts from the early years of NTP research, and recent activities in the field. Specific issues addressed include NERVA rocket-engine technology, the development of nuclear rocket propulsion at Los Alamos, fuel-element development, reactor testing for the Rover program, and an overview of NTP concepts and research emphasizing two decades of NASA research. Also addressed are the development of the 'nuclear light bulb' closed-cycle gas core and a demonstration of a fissioning UF6 gas in an argon vortex. The recent developments reviewed include the application of NTP to NASA's Lunar Space Transportation System, the use of NTP for the Space Exploration Initiative, and the development of nuclear rocket engines in the former Soviet Union.

  15. Thermal performance analysis of Brayton cycle with waste heat recovery boiler for diesel engines of offshore oil production facilities

    International Nuclear Information System (INIS)

    Liu, Xianglong; Gong, Guangcai; Wu, Yi; Li, Hangxin

    2016-01-01

    Highlights: • Comparison of Brayton cycle with WHRB adopted in diesel engines with and without fans by thermal performance. • Waste heat recovery technology for FPSO. • The thermoeconomic analysis for the heat recovery for FPSO. - Abstract: This paper presents the theoretical analysis and on-site testing on the thermal performance of the waste heat recovery system for offshore oil production facilities, including the components of diesel engines, thermal boilers and waste heat boilers. We use the ideal air standard Brayton cycle to analyse the thermal performance. In comparison with the traditional design, the fans at the engine outlet of the waste heat recovery boiler is removed due to the limited space of the offshore platform. The cases with fan and without fan are compared in terms of thermal dynamics performance, energy efficiency and thermo-economic index of the system. The results show that the application of the WHRB increases the energy efficiency of the whole system, but increases the flow resistance in the duct. It is proved that as the waste heat recovery boiler takes the place of the thermal boiler, the energy efficiency of whole system without fan is slightly reduced but heat recovery efficiency is improved. This research provides an important guidance to improve the waste heat recovery for offshore oil production facilities.

  16. ATTIRE (analytical tools for thermal infrared engineering): A sensor simulation and modeling package

    Science.gov (United States)

    Jaggi, S.

    1993-01-01

    The Advanced Sensor Development Laboratory (ASDL) at the Stennis Space Center develops, maintains and calibrates remote sensing instruments for the National Aeronautics & Space Administration (NASA). To perform system design trade-offs, analysis, and establish system parameters, ASDL has developed a software package for analytical simulation of sensor systems. This package called 'Analytical Tools for Thermal InfraRed Engineering' - ATTIRE, simulates the various components of a sensor system. The software allows each subsystem of the sensor to be analyzed independently for its performance. These performance parameters are then integrated to obtain system level information such as Signal-to-Noise Ratio (SNR), Noise Equivalent Radiance (NER), Noise Equivalent Temperature Difference (NETD) etc. This paper describes the uses of the package and the physics that were used to derive the performance parameters.

  17. Interring Gas Dynamic Analysis of Piston in a Diesel Engine considering the Thermal Effect

    Directory of Open Access Journals (Sweden)

    Wanyou Li

    2015-01-01

    Full Text Available Understanding the interaction between ring dynamics and gas transport in ring pack systems is crucial and needs to be imperatively studied. The present work features detailed interring gas dynamics of piston ring pack behavior in internal combustion engines. The model is developed for a ring pack with four rings. The dynamics of ring pack are simulated. Due to the fact that small changes in geometry of the grooves and lands would have a significant impact on the interring gas dynamics, the thermal deformation of piston has been considered during the ring pack motion analysis in this study. In order to get the temperature distribution of piston head more quickly and accurately, an efficient method utilizing the concept of inverse heat conduction is presented. Moreover, a sensitive analysis based on the analysis of partial regression coefficients is presented to investigate the effect of groove parameters on blowby.

  18. Numerical method for assessing the potential of smart engine thermal management: Application to a medium-upper segment passenger car

    International Nuclear Information System (INIS)

    Caresana, F.; Bilancia, M.; Bartolini, C.M.

    2011-01-01

    Significant reductions in vehicle fuel consumption can be obtained through a greater control of the thermal status of the engine, especially under partial load conditions. Different systems have been proposed to implement this concept, referred to as improved engine thermal management. The amount of fuel saved depends on the components and layout of the engine cooling plant and on the performance of its control system. In this work, a method was developed to calculate the theoretical minimum fuel consumption of a passenger car and used as a reference in comparing different engine cooling system concepts. A high-medium class car was taken as an example and simulated on standard cycles. Models for power train and cooling system components were developed and linked to simulate the vehicle. A preliminary analysis of the engine was performed using AVL's Boost program. The fuel consumption of the complete vehicle, equipped with a conventional cooling plant, was determined on standard cycles and compared with that of a vehicle equipped with a 'perfect' cooling system, to calculate the theoretical reduction in fuel consumption. - Highlights: → We propose a method for assessing the potential of smart engine thermal management. → A conventional cooling system is compared to a 'perfect' one to estimate fuel economy. → We tested the method in an upper-medium segment passenger car.

  19. Thermo-acoustic instabilities of high-frequency combustion in rocket engines; Instabilites thermo-acoustiques de combustion haute-frequence dans les moteurs fusees

    Energy Technology Data Exchange (ETDEWEB)

    Cheuret, F.

    2005-10-15

    Rocket motors are confined environments where combustion occurs in extreme conditions. Combustion instabilities can occur at high frequencies; they are tied to the acoustic modes of the combustion chamber. A common research chamber, CRC, allows us to study the response of a turbulent two-phase flame to acoustic oscillations of low or high amplitudes. The chamber is characterised under cold conditions to obtain, in particular, the relative damping coefficient of acoustic oscillations. The structure and frequency of the modes are determined in the case where the chamber is coupled to a lateral cavity. We have used a powder gun to study the response to a forced acoustic excitation at high amplitude. The results guide us towards shorter flames. The injectors were then modified to study the combustion noise level as a function of injection conditions. The speed of the gas determines whether the flames are attached or lifted. The noise level of lifted flames is higher. That of attached flames is proportional to the Weber number. The shorter flames whose length is less than the radius of the CRC, necessary condition to obtain an effective coupling, are the most sensitive to acoustic perturbations. The use of a toothed wheel at different positions in the chamber allowed us to obtain informations on the origin of the thermo-acoustic coupling, main objective of this thesis. The flame is sensitive to pressure acoustic oscillations, with a quasi-zero response time. These observations suggest that under the conditions of the CRC, we observe essentially the response of chemical kinetics to pressure oscillations. (author)

  20. Investigation of Thermal Comfort Conditions in Higher Education Facilities: A Case Study for Engineering Faculty in Edirne

    Directory of Open Access Journals (Sweden)

    E. Mıhlayanlar

    2017-02-01

    Full Text Available In this study, a higher education institution in Edirne (Trakya University Engineering Faculty is investigated for indoor thermal comfort conditions of the classrooms (indoor temperature, relative humidity, average radiant temperature, “Satisfaction from thermal environment” (PMV and “Dissatisfaction from thermal environment” (PPD. The classrooms in the institution are heated by a central heating system and utilise natural ventilation system. Measurements were taken with the proper devices at the same time of the weekdays during lecture times in winter (heating season in December. The results obtained from measurements are given in graphics and compared with the values given in ASHRAE 55 and ISO 7730 standards.

  1. C/C composites for rocket chamber applications. Part 2: Fabrication and evaluation tests of rocket chamber

    Science.gov (United States)

    Sato, Masahiro; Tadano, Makoto; Ueda, Shuichi; Kuroda, Yukio; Kusaka, Kazuo; Suemitsu, Takeshi; Hasegawa, Satoshi; Kude, Yukinori

    1995-05-01

    Carbon fiber-reinforced carbon matrix (C/C) composites coated with SiC are promising candidates for use in the main structural materials of the body of spaceplanes and combustion chambers of rocket engines, because of their superior properties of high specific strength, specific modulus, and fracture strength at high temperatures. However, C/C composite has poor resistance to oxidation, and protection from the oxidating environment is crucial. Conventional C/C composites for use in the high-temperature components of rocket engines are coated with SiC. However, due to the difference in the thermal expansion rates of the SiC coating layer and the base materials, cracks occur in the SiC coating layer during the coating process, and oxygen diffuses to the base material through the cracks during repeated temperature cycling in the rocket combustion environment. To protect the base materials from oxidation at high temperatures, we have employed SiC C/C-coated composites with a modified matrix and also developed SiC C/C functionally gradient materials (FGM's). In this test series, three kinds of combustion chambers were constructed for the Reaction Control System (RCS) subscale engine of H-II Orbiting Plane (HOPE): (1) Conventional C/C composites, (2) SiC C/C-coated composites with a modified matrix, and (3) SiC C/C FGM's. Firing tests were performed at sea level at a temperature around 2000 K using nitrogen tetroxide (NTO)/monomethyl hydrazine (MMH) propellant to evaluate the durability of these chambers. This test series showed that conventional C/C composite developed no microcracks and delamination in the coating layer at 1940 K. Modified matrix C/C composite also did not suffer microcracks and delamination at the boundary between the SiC and the base materials when the inner surface temperature was 1875 K. However, microcracks were observed at injector flange surface after these test cycles. In the test series of FGM's chamber, it was shown that coating with FGM

  2. Hydrocarbon Rocket Technology Impact Forecasting

    Science.gov (United States)

    Stuber, Eric; Prasadh, Nishant; Edwards, Stephen; Mavris, Dimitri N.

    2012-01-01

    Forecasting method is a normative forecasting technique that allows the designer to quantify the effects of adding new technologies on a given design. This method can be used to assess and identify the necessary technological improvements needed to close the gap that exists between the current design and one that satisfies all constraints imposed on the design. The TIF methodology allows for more design knowledge to be brought to the earlier phases of the design process, making use of tools such as Quality Function Deployments, Morphological Matrices, Response Surface Methodology, and Monte Carlo Simulations.2 This increased knowledge allows for more informed decisions to be made earlier in the design process, resulting in shortened design cycle time. This paper will investigate applying the TIF method, which has been widely used in aircraft applications, to the conceptual design of a hydrocarbon rocket engine. In order to reinstate a manned presence in space, the U.S. must develop an affordable and sustainable launch capability. Hydrocarbon-fueled rockets have drawn interest from numerous major government and commercial entities because they offer a low-cost heavy-lift option that would allow for frequent launches1. However, the development of effective new hydrocarbon rockets would likely require new technologies in order to overcome certain design constraints. The use of advanced design methods, such as the TIF method, enables the designer to identify key areas in need of improvement, allowing one to dial in a proposed technology and assess its impact on the system. Through analyses such as this one, a conceptual design for a hydrocarbon-fueled vehicle that meets all imposed requirements can be achieved.

  3. Thermal green protein, an extremely stable, nonaggregating fluorescent protein created by structure-guided surface engineering.

    Science.gov (United States)

    Close, Devin W; Paul, Craig Don; Langan, Patricia S; Wilce, Matthew C J; Traore, Daouda A K; Halfmann, Randal; Rocha, Reginaldo C; Waldo, Geoffery S; Payne, Riley J; Rucker, Joseph B; Prescott, Mark; Bradbury, Andrew R M

    2015-07-01

    In this article, we describe the engineering and X-ray crystal structure of Thermal Green Protein (TGP), an extremely stable, highly soluble, non-aggregating green fluorescent protein. TGP is a soluble variant of the fluorescent protein eCGP123, which despite being highly stable, has proven to be aggregation-prone. The X-ray crystal structure of eCGP123, also determined within the context of this paper, was used to carry out rational surface engineering to improve its solubility, leading to TGP. The approach involved simultaneously eliminating crystal lattice contacts while increasing the overall negative charge of the protein. Despite intentional disruption of lattice contacts and introduction of high entropy glutamate side chains, TGP crystallized readily in a number of different conditions and the X-ray crystal structure of TGP was determined to 1.9 Å resolution. The structural reasons for the enhanced stability of TGP and eCGP123 are discussed. We demonstrate the utility of using TGP as a fusion partner in various assays and significantly, in amyloid assays in which the standard fluorescent protein, EGFP, is undesirable because of aberrant oligomerization. © 2014 Wiley Periodicals, Inc.

  4. ROCKET PORT CLOSURE

    Science.gov (United States)

    Mattingly, J.T.

    1963-02-12

    This invention provides a simple pressure-actuated closure whereby windowless observation ports are opened to the atmosphere at preselected altitudes. The closure comprises a disk which seals a windowless observation port in rocket hull. An evacuated instrument compartment is affixed to the rocket hull adjacent the inner surface of the disk, while the outer disk surface is exposed to the atmosphere through which the rocket is traveling. The pressure differential between the evacuated instrument compartment and the relatively high pressure external atmosphere forces the disk against the edge of the observation port, thereby effecting a tight seai. The instrument compartment is evacuated to a pressure equal to the atmospheric pressure existing at the altitude at which it is desiretl that the closure should open. When the rocket reaches this preselected altitude, the inwardly directed atmospheric force on the disk is just equaled by the residual air pressure force within the instrument compartment. Consequently, the closure disk falls away and uncovers the open observation port. The separation of the disk from the rocket hull actuates a switch which energizes the mechanism of a detecting instrument disposed within the instrument compartment. (AE C)

  5. Phase of Photothermal Emission Analysis as a Diagnostic Tool for Thermal Barrier Coatings on Serviceable Engine Components

    Science.gov (United States)

    Kakuda, Tyler

    Power generation and aircraft companies are continuously improving the efficiency of gas turbines to meet economic and environmental goals. The trend towards higher efficiency has been achieved in part by raising the operating temperature of engines. At elevated temperatures, engine components are subject to many forms of degradation including oxidation, creep deformation and thermal cycle fatigue. To minimize these harmful effects, ceramic thermal barrier coatings (TBCs) are routinely used to insulate metal components from excessive heat loads. Efforts to make realistic performance assessments of current and candidate coating materials has led to a diverse battery of creative measurement techniques. While it is unrealistic to envision a single measurement that would provide all conceivable information about the TBC, it is arguable that the capability for the single most important measurement is still lacking. A quantitative and nondestructive measurement of the thermal protection offered by a coating is not currently among the measurements one can employ on a serviceable engine part (or even many experimental specimens). In this contribution, phase of photothermal emission analysis (PopTea) is presented as a viable thermal property measurement for serviceable engine components. As it will be shown, PopTea has the versatility to make measurements on gas turbine parts in situ, with the goal of monitoring TBCs over the lifetime of the engine. The main challenges toward this goal are dealing with changes that occur to the TBC during service. Several of the main degradations seen on engine equipment include: aging, surface contamination and infiltration of foreign deposits. Measuring coatings under these conditions, is the impetus of this work. Furthermore, it is demonstrated that PopTea can be used on real engine equipment with measurements made on an actual turbine blade.

  6. Durability of zirconia thermal-barrier ceramic coatings on air-cooled turbine blades in cyclic jet engine operation

    Science.gov (United States)

    Liebert, C. H.; Jacobs, R. E.; Stecura, S.; Morse, C. R.

    1976-01-01

    Thermal barrier ceramic coatings of stabilized zirconia over a bond coat of Ni Cr Al Y were tested for durability on air cooled turbine rotor blades in a research turbojet engine. Zirconia stabilized with either yttria, magnesia, or calcia was investigated. On the basis of durability and processing cost, the yttria stabilized zirconia was considered the best of the three coatings investigated.

  7. Analysis of thermal stress of the piston during non-stationary heat flow in a turbocharged Diesel engine

    Science.gov (United States)

    Gustof, P.; Hornik, A.

    2016-09-01

    In the paper, numeric calculations of thermal stresses of the piston in a turbocharged Diesel engine in the initial phase of its work were carried out based on experimental studies and the data resulting from them. The calculations were made using a geometrical model of the piston in a five-cylinder turbocharged Diesel engine with a capacity of about 2300 cm3, with a direct fuel injection to the combustion chamber and a power rating of 85 kW. In order to determine the thermal stress, application of own mathematical models of the heat flow in characteristic surfaces of the piston was required to show real processes occurring on the surface of the analysed component. The calculations were performed using a Geostar COSMOS/M program module. A three-dimensional geometric model of the piston was created in this program based on a real component, in order to enable the calculations and analysis of thermal stresses during non-stationary heat flow. Modelling of the thermal stresses of the piston for the engine speed n=4250 min-1 and engine load λ=1.69 was carried out.

  8. Rocket Flight Path

    Directory of Open Access Journals (Sweden)

    Jamie Waters

    2014-09-01

    Full Text Available This project uses Newton’s Second Law of Motion, Euler’s method, basic physics, and basic calculus to model the flight path of a rocket. From this, one can find the height and velocity at any point from launch to the maximum altitude, or apogee. This can then be compared to the actual values to see if the method of estimation is a plausible. The rocket used for this project is modeled after Bullistic-1 which was launched by the Society of Aeronautics and Rocketry at the University of South Florida.

  9. Performance and emission characteristics of the thermal barrier coated SI engine by adding argon inert gas to intake mixture

    Directory of Open Access Journals (Sweden)

    T. Karthikeya Sharma

    2015-11-01

    Full Text Available Dilution of the intake air of the SI engine with the inert gases is one of the emission control techniques like exhaust gas recirculation, water injection into combustion chamber and cyclic variability, without scarifying power output and/or thermal efficiency (TE. This paper investigates the effects of using argon (Ar gas to mitigate the spark ignition engine intake air to enhance the performance and cut down the emissions mainly nitrogen oxides. The input variables of this study include the compression ratio, stroke length, and engine speed and argon concentration. Output parameters like TE, volumetric efficiency, heat release rates, brake power, exhaust gas temperature and emissions of NOx, CO2 and CO were studied in a thermal barrier coated SI engine, under variable argon concentrations. Results of this study showed that the inclusion of Argon to the input air of the thermal barrier coated SI engine has significantly improved the emission characteristics and engine’s performance within the range studied.

  10. Highlights of 50 years of Aerojet, a pioneering American rocket company, 1942-1992

    Science.gov (United States)

    Winter, Frank H.; James, George S.

    1995-05-01

    The "pre-history" of Aerojet is recalled, followed by a survey of Aerojet's solid-fuel and liquid-fuel JATOs (Jet-Assisted Take-Off) to aircraft prime powerplants, missile sustainer motors, boosters, sounding rocket engines and, finally, nuclear powered rocket engines (NERVA).

  11. Thermal barrier coatings: Coating methods, performance, and heat engine applications. (Latest citations from the EI Compendex*plus database). Published Search

    Energy Technology Data Exchange (ETDEWEB)

    NONE

    1997-02-01

    The bibliography contains citations concerning conference proceedings on coating methods, performance evaluations, and applications of thermal barrier coatings as protective coatings for heat engine components against high temperature corrosions and chemical erosions. The developments of thermal barrier coating techniques for high performance and reliable gas turbines, diesel engines, jet engines, and internal combustion engines are presented. Topics include plasma sprayed coating methods, yttria stabilized zirconia coatings, coating life models, coating failure and durability, thermal shock and cycling, and acoustic emission analysis of coatings. (Contains 50-250 citations and includes a subject term index and title list.) (Copyright NERAC, Inc. 1995)

  12. Thermal barrier coatings: Coating methods, performance, and heat engine applications. (Latest citations from the EI Compendex*plus database). Published Search

    Energy Technology Data Exchange (ETDEWEB)

    NONE

    1995-11-01

    The bibliography contains citations concerning conference proceedings on coating methods, performance evaluations, and applications of thermal barrier coatings as protective coatings for heat engine components against high temperature corrosions and chemical erosions. The developments of thermal barrier coating techniques for high performance and reliable gas turbines, diesel engines, jet engines, and internal combustion engines are presented. Topics include plasma sprayed coating methods, yttria stabilized zirconia coatings, coating life models, coating failure and durability, thermal shock and cycling, and acoustic emission analysis of coatings. (Contains 50-250 citations and includes a subject term index and title list.) (Copyright NERAC, Inc. 1995)

  13. Rockets in World War I

    Science.gov (United States)

    2004-01-01

    World War I enlisted rockets once again for military purposes. French pilots rigged rockets to the wing struts of their airplanes and aimed them at enemy observation balloons filled with highly inflammable hydrogen.

  14. This Is Rocket Science!

    Science.gov (United States)

    Keith, Wayne; Martin, Cynthia; Veltkamp, Pamela

    2013-09-01

    Using model rockets to teach physics can be an effective way to engage students in learning. In this paper, we present a curriculum developed in response to an expressed need for helping high school students review physics equations in preparation for a state-mandated exam. This required a mode of teaching that was more advanced and analytical than that offered by Estes Industries, but more basic than the analysis of Nelson et al. In particular, drag is neglected until the very end of the exercise, which allows the concept of conservation of energy to be shown when predicting the rocket's flight. Also, the variable mass of the rocket motor is assumed to decrease linearly during the flight (while the propulsion charge and recovery delay charge are burning) and handled simplistically by using an average mass value. These changes greatly simplify the equations needed to predict the times and heights at various stages of flight, making it more useful as a review of basic physics. Details about model rocket motors, range safety, and other supplemental information may be found online at Apogee Components4 and the National Association of Rocketry.5

  15. Thiokol Solid Rocket Motors

    Science.gov (United States)

    Graves, S. R.

    2000-01-01

    This paper presents viewgraphs on thiokol solid rocket motors. The topics include: 1) Communications; 2) Military and government intelligence; 3) Positioning satellites; 4) Remote sensing; 5) Space burial; 6) Science; 7) Space manufacturing; 8) Advertising; 9) Space rescue space debris management; 10) Space tourism; 11) Space settlements; 12) Hazardous waste disposal; 13) Extraterrestrial resources; 14) Fast package delivery; and 15) Space utilities.

  16. The Relativistic Rocket

    Science.gov (United States)

    Antippa, Adel F.

    2009-01-01

    We solve the problem of the relativistic rocket by making use of the relation between Lorentzian and Galilean velocities, as well as the laws of superposition of successive collinear Lorentz boosts in the limit of infinitesimal boosts. The solution is conceptually simple, and technically straightforward, and provides an example of a powerful…

  17. This "Is" Rocket Science!

    Science.gov (United States)

    Keith, Wayne; Martin, Cynthia; Veltkamp, Pamela

    2013-01-01

    Using model rockets to teach physics can be an effective way to engage students in learning. In this paper, we present a curriculum developed in response to an expressed need for helping high school students review physics equations in preparation for a state-mandated exam. This required a mode of teaching that was more advanced and analytical…

  18. Low toxicity rocket propellants

    NARCIS (Netherlands)

    Wink, J.

    2014-01-01

    Hydrazine (N2H4) and its hypergolic mate nitrogen tetroxide (N2O4) are used on virtually all spacecraft and on a large number of launch vehicles. In recent years however, there has been an effort in identifying and developing alternatives to replace hydrazine as a rocket propellant.

  19. L'évolution des combustibles pour moteurs thermiques Evolution of Fuels for Thermal Engines

    Directory of Open Access Journals (Sweden)

    Balaceanu J. C.

    2006-11-01

    Full Text Available Depuis la crise pétrolière, l'accroissement des prix et les craintes de pénurie ont conduit à adapter les moteurs aux combustibles disponibles. Si la situation apparaît comme moins tendue, l'industrie des machines thermiques, qui pendant longtemps a eu comme partenaire une industrie du pétrole très sûre, se trouve cependant confrontée à un marché des combustibles incertain dans son ravitaillement et surtout dans ses prix. Les progrès des moteurs diesel et des turbines à gaz, dûs à une meilleure adaptation à leur usage et aussi à l'évolution de la technologie, supposent que les combustibles n'apporteront aucune contrainte majeure en quantité ou en qualité. La modification des usages dévolus désormais au pétrole entraînera une réduction de la coupe lourde et un raffinage plus profond des bruts avec en particulier un développement du craquage catalytique et de la viscoréduction. Or, ces différentes opérations de conversion peuvent conduire à une détérioration de la qualité des combustibles moins grave pour le gazole que pour le fuel lourd. Dans les différents domaines impliqués, les parades technologiques sont en cours de développement. L'industrie des machines thermiques, qui poursuit l'amélioration des engins, et l'industrie du pétrole, qui recherche une réduction des prix des combustibles, sont donc conduites à un compromis optimal auquel elles ne peuvent accéder efficacement qu'en définissant les règles du jeu c'est-à-dire des spécifications internationales rigoureuses des combustibles. Since the oil crisis, the increase in prices and fears of a shortage have led to the adapting of engines to what fuels are available. Whereas the situation now seems somewhat less tense, the thermal machinery industry, which for a long time had a very reliable petroleum industry as its partner, nonetheless finds itself confronted with an uncertain fuel market with regard to supplies and especially to prices. Progress

  20. Multi-Sensing system for outdoor thermal monitoring: Application to large scale civil engineering components

    Science.gov (United States)

    Crinière, Antoine; Dumoulin, Jean; Manceau, Jean-Luc; Perez, Laetitia; Bourquin, Frederic

    2014-05-01

    Aging of transport infrastructures combined with traffic and climatic solicitations contribute to the reduction of their performances. To address and quantify the resilience of civil engineering structure, investigations on robust, fast and efficient methods are required. Among research works carried out at IFSTTAR, methods for long term monitoring face an increasing demand. Such works take benefits of this last decade technological progresses in ICT domain. The present study follows the ISTIMES European project [1], which aimed at demonstrate the ability of different electromagnetic sensing techniques, processing methods and ICT architecture, to be used for long term monitoring of critical transport infrastructures. Thanks to this project a multi-sensing techniques system, able to date and synchronize measurements carried out by infrared thermography coupled with various measurements data (i.e. weather parameters), have been designed, developed and implemented on real site [2]. Among experiments carried out on real transport infrastructure, it has been shown, for the "Musmesci" bridge deck (Italy), that by using infrared thermal image sequence with weather measurements during sevral days it was possible to develop analysis methods able to produce qualitative and quantitative data [3]. In the present study, added functionalities were designed and added to the "IrLAW" system in order to reach full autonomy in term of power supply, very long term measurement capability (at least 1 year) and automated data base feeding. The surveyed civil engineering structures consist in two concrete beams of 16 m long and 21 T weight each. One of the two beams was damage by high energy mechanical impact at the IFSTTAR falling rocks test station facilities located in the French Alpes [4]. The system is composed of one IR uncooled microbolometric camera (FLIR SC325) with a 320X240 Focal Plane Array detector in band III, a weather station VAISALA WXT520, a GPS, a failover power supply