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…

An improved heat transfer configuration for a solid-core nuclear thermal rocket engine

International Nuclear Information System (INIS)

Clark, J.S.; Walton, J.T.; Mcguire, M.L.

1992-07-01

Interrupted flow, impingement cooling, and axial power distribution are employed to enhance the heat-transfer configuration of a solid-core nuclear thermal rocket engine. Impingement cooling is introduced to increase the local heat-transfer coefficients between the reactor material and the coolants. Increased fuel loading is used at the inlet end of the reactor to enhance heat-transfer capability where the temperature differences are the greatest. A thermal-hydraulics computer program for an unfueled NERVA reactor core is employed to analyze the proposed configuration with attention given to uniform fuel loading, number of channels through the impingement wafers, fuel-element length, mass-flow rate, and wafer gap. The impingement wafer concept (IWC) is shown to have heat-transfer characteristics that are better than those of the NERVA-derived reactor at 2500 K. The IWC concept is argued to be an effective heat-transfer configuration for solid-core nuclear thermal rocket engines. 11 refs

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...

Validation of Supersonic Film Cooling Modeling for Liquid Rocket Engine Applications

Science.gov (United States)

Morris, Christopher I.; Ruf, Joseph H.

2010-01-01

Topics include: upper stage engine key requirements and design drivers; Calspan "stage 1" results, He slot injection into hypersonic flow (air); test articles for shock generator diagram, slot injector details, and instrumentation positions; test conditions; modeling approach; 2-d grid used for film cooling simulations of test article; heat flux profiles from 2-d flat plate simulations (run #4); heat flux profiles from 2-d backward facing step simulations (run #43); isometric sketch of single coolant nozzle, and x-z grid of half-nozzle domain; comparison of 2-d and 3-d simulations of coolant nozzles (run #45); flowfield properties along coolant nozzle centerline (run #45); comparison of 3-d CFD nozzle flow calculations with experimental data; nozzle exit plane reduced to linear profile for use in 2-d film-cooling simulations (run #45); synthetic Schlieren image of coolant injection region (run #45); axial velocity profiles from 2-d film-cooling simulation (run #45); coolant mass fraction profiles from 2-d film-cooling simulation (run #45); heat flux profiles from 2-d film cooling simulations (run #45); heat flux profiles from 2-d film cooling simulations (runs #47, #45, and #47); 3-d grid used for film cooling simulations of test article; heat flux contours from 3-d film-cooling simulation (run #45); and heat flux profiles from 3-d and 2-d film cooling simulations (runs #44, #46, and #47).

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

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.

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

Engineering thermal engine rocket adventurer for space nuclear application

International Nuclear Information System (INIS)

Nam, Seung H.; Suh, Kune Y.; Kang, Seong G.

2008-01-01

The conceptual design for the first-of-a-kind engineering of Thermal Engine Rocket Adventure (TERA) is described. TERA comprising the Battery Omnibus Reactor Integral System (BORIS) as the heat resource and the Space Propulsion Reactor Integral System (SPRIS) as the propulsion system, is one of the advanced Nuclear Thermal Rocket (NTR) engine utilizing hydrogen (H 2 ) propellant being developed at present time. BORIS in this application is an open cycle high temperature gas cooled reactor that has eighteen fuel elements for propulsion and one fuel element for electricity generation and propellant pumping. Each fuel element for propulsion has its own small nozzle. The nineteen fuel elements are arranged into hexagonal prism shape in the core and surrounded by outer Be reflector. The TERA maximum power is 1,000 MW th , specific impulse 1,000 s, thrust 250,000 N, and the total mass is 550 kg including the reactor, turbo pump and auxiliaries. Each fuel element comprises the fuel assembly, moderators, pressure tube and small nozzle. The TERA fuel assembly is fabricated of 93% enriched 1.5 mm (U, Zr, Nb)C wafers in 25.3% voided Square Lattice Honeycomb (SLHC). The H 2 propellant passes through these flow channels. This study is concerned with thermohydrodynamic analysis of the fuel element for propulsion with hypothetical axial power distribution because nuclear analysis of TERA has not been performed yet. As a result, when the power distribution of INSPI's M-SLHC is applied to the fuel assembly, the local heat concentration of fuel is more serious and the pressure of the initial inlet H 2 is higher than those of constant average power distribution applied. This means the fuel assembly geometry of 1.5 mm fuel wafers and 25.3% voided SLHC needs to be changed in order to reduce thermal and mechanical shocks. (author)

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

Performance of a RBCC Engine in Rocket-Operation

Science.gov (United States)

Tomioka, Sadatake; Kubo, Takahiro; Noboru Sakuranaka; Tani, Koichiro

Combination of a scramjet (supersonic combustion ramjet) flow-pass with embedded rocket engines (the combined system termed as Rocket-based Combined Cycle engine) are expected to be the most effective propulsion system for space launch vehicles. Either SSTO (Single Stage To Orbit) system or TSTO (Two Stage To Orbit) system with separation at high altitude needs final stage acceleration in space, so that the RBCC (Rocket Based Combined Cycle) engine should be operated as rocket engines. Performance of the scramjet combustor as the extension to the rocket nozzle, was experimentally evaluated by injecting inert gas at various pressure through the embedded rocket chamber while the whole sub-scaled model was placed in a low pressure chamber connected to an air-driven ejector system. The results showed that the thrust coefficient was about 1.2, the low value being found to mainly due to the friction force on the scramjet combustor wall, while blocking the scramjet flow pass’s opening to increase nozzle extension thrust surface, was found to have little effects on the thrust performance. The combustor was shortened to reduce the friction loss, however, degree of reduction was limited as friction decreased rapidly with distance from the onset of the scramjet combustor.

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.

Current status of rocket developments in universities -development of a small hybrid rocket with a swirling oxidizer flow type engine

OpenAIRE

Yuasa, Saburo; Kitagawa, Koki

2005-01-01

To develop an experimental small hybrid rocket with a swirling gaseous oxygen flow type engine, we made a flight model engine. Burning tests of the engine showed that a maximum thrust of 692 N and a specific impulse of 263 s (at sea level) were achieved. We designed a small hybrid rocket with this engine. The rocket measured 1.8 m in length and 15.4 kg in mass. To confirm the flight stability of the rocket, wind tunnel tests using a 112-scale model of the rocket and simulations of the flight ...

Rocket propulsion elements - An introduction to the engineering of rockets (6th revised and enlarged edition)

Science.gov (United States)

Sutton, George P.

The subject of rocket propulsion is treated with emphasis on the basic technology, performance, and design rationale. Attention is given to definitions and fundamentals, nozzle theory and thermodynamic relations, heat transfer, flight performance, chemical rocket propellant performance analysis, and liquid propellant rocket engine fundamentals. The discussion also covers solid propellant rocket fundamentals, hybrid propellant rockets, thrust vector control, selection of rocket propulsion systems, electric propulsion, and rocket testing.

Comparison of Engine Cycle Codes for Rocket-Based Combined Cycle Engines

Science.gov (United States)

Waltrup, Paul J.; Auslender, Aaron H.; Bradford, John E.; Carreiro, Louis R.; Gettinger, Christopher; Komar, D. R.; McDonald, J.; Snyder, Christopher A.

2002-01-01

This paper summarizes the results from a one day workshop on Rocket-Based Combined Cycle (RBCC) Engine Cycle Codes held in Monterey CA in November of 2000 at the 2000 JANNAF JPM with the authors as primary participants. The objectives of the workshop were to discuss and compare the merits of existing Rocket-Based Combined Cycle (RBCC) engine cycle codes being used by government and industry to predict RBCC engine performance and interpret experimental results. These merits included physical and chemical modeling, accuracy and user friendliness. The ultimate purpose of the workshop was to identify the best codes for analyzing RBCC engines and to document any potential shortcomings, not to demonstrate the merits or deficiencies of any particular engine design. Five cases representative of the operating regimes of typical RBCC engines were used as the basis of these comparisons. These included Mach 0 sea level static and Mach 1.0 and Mach 2.5 Air-Augmented-Rocket (AAR), Mach 4 subsonic combustion ramjet or dual-mode scramjet, and Mach 8 scramjet operating modes. Specification of a generic RBCC engine geometry and concomitant component operating efficiencies, bypass ratios, fuel/oxidizer/air equivalence ratios and flight dynamic pressures were provided. The engine included an air inlet, isolator duct, axial rocket motor/injector, axial wall fuel injectors, diverging combustor, and exit nozzle. Gaseous hydrogen was used as the fuel with the rocket portion of the system using a gaseous H2/O2 propellant system to avoid cryogenic issues. The results of the workshop, even after post-workshop adjudication of differences, were surprising. They showed that the codes predicted essentially the same performance at the Mach 0 and I conditions, but progressively diverged from a common value (for example, for fuel specific impulse, Isp) as the flight Mach number increased, with the largest differences at Mach 8. The example cases and results are compared and discussed in this paper.

Design considerations for a pressure-driven multi-stage rocket

Science.gov (United States)

Sauerwein, Steven Craig

2002-01-01

The purpose of this study was to examine the feasibility of using propellant tank pressurization to eliminate the use of high-pressure turbopumps in multi-stage liquid-fueled satellite launchers. Several new technologies were examined to reduce the mass of such a rocket. Composite materials have a greater strength-to-weight ratio than metals and can be used to reduce the weight of rocket propellant tanks and structure. Catalytically combined hydrogen and oxygen can be used to heat pressurization gas, greatly reducing the amount of gas required. Ablatively cooled rocket engines can reduce the complexity and cost of the rocket. Methods were derived to estimate the mass of the various rocket components. These included a method to calculate the amount of gas needed to pressurize a propellant tank by modeling the behavior of the pressurization gas as the liquid propellant flows out of the tank. A way to estimate the mass and size of a ablatively cooled composite cased rocket engine. And a method to model the flight of such a rocket through the atmosphere in conjunction with optimization of the rockets trajectory. The results show that while a liquid propellant rocket using tank pressurization are larger than solid propellant rockets and turbopump driven liquid propellant rockets, they are not impractically large.

Liquid Rocket Engine Testing

Science.gov (United States)

Rahman, Shamim

2005-01-01

Comprehensive Liquid Rocket Engine testing is essential to risk reduction for Space Flight. Test capability represents significant national investments in expertise and infrastructure. Historical experience underpins current test capabilities. Test facilities continually seek proactive alignment with national space development goals and objectives including government and commercial sectors.

Reusable Rocket Engine Advanced Health Management System. Architecture and Technology Evaluation: Summary

Science.gov (United States)

Pettit, C. D.; Barkhoudarian, S.; Daumann, A. G., Jr.; Provan, G. M.; ElFattah, Y. M.; Glover, D. E.

1999-01-01

In this study, we proposed an Advanced Health Management System (AHMS) functional architecture and conducted a technology assessment for liquid propellant rocket engine lifecycle health management. The purpose of the AHMS is to improve reusable rocket engine safety and to reduce between-flight maintenance. During the study, past and current reusable rocket engine health management-related projects were reviewed, data structures and health management processes of current rocket engine programs were assessed, and in-depth interviews with rocket engine lifecycle and system experts were conducted. A generic AHMS functional architecture, with primary focus on real-time health monitoring, was developed. Fourteen categories of technology tasks and development needs for implementation of the AHMS were identified, based on the functional architecture and our assessment of current rocket engine programs. Five key technology areas were recommended for immediate development, which (1) would provide immediate benefits to current engine programs, and (2) could be implemented with minimal impact on the current Space Shuttle Main Engine (SSME) and Reusable Launch Vehicle (RLV) engine controllers.

Development of nuclear rocket engine technology

International Nuclear Information System (INIS)

Gunn, S.V.

1989-01-01

Research sponsored by the Atomic Energy Commission, the USAF, and NASA (later on) in the area of nuclear rocket propulsion is discussed. It was found that a graphite reactor, loaded with highly concentrated Uranium 235, can be used to heat high pressure liquid hydrogen to temperatures of about 4500 R, and to expand the hydrogen through a high expansion ratio rocket nozzle assembly. The results of 20 reactor tests conducted at the Nevada Test Site between July 1959 and June 1969 are analyzed. On the basis of these results, the feasibility of solid graphite reactor/nuclear rocket engines is revealed. It is maintained that this technology will support future space propulsion requirements, using liquid hydrogen as the propellant, for thrust requirements ranging from 25,000 lbs to 250,000 lbs, with vacuum specific impulses of at least 850 sec and with full engine throttle capability. 12 refs

Manufacturing Process Developments for Regeneratively-Cooled Channel Wall Rocket Nozzles

Science.gov (United States)

Gradl, Paul; Brandsmeier, Will

2016-01-01

Regeneratively cooled channel wall nozzles incorporate a series of integral coolant channels to contain the coolant to maintain adequate wall temperatures and expand hot gas providing engine thrust and specific impulse. NASA has been evaluating manufacturing techniques targeting large scale channel wall nozzles to support affordability of current and future liquid rocket engine nozzles and thrust chamber assemblies. The development of these large scale manufacturing techniques focus on the liner formation, channel slotting with advanced abrasive water-jet milling techniques and closeout of the coolant channels to replace or augment other cost reduction techniques being evaluated for nozzles. NASA is developing a series of channel closeout techniques including large scale additive manufacturing laser deposition and explosively bonded closeouts. A series of subscale nozzles were completed evaluating these processes. Fabrication of mechanical test and metallography samples, in addition to subscale hardware has focused on Inconel 625, 300 series stainless, aluminum alloys as well as other candidate materials. Evaluations of these techniques are demonstrating potential for significant cost reductions for large scale nozzles and chambers. Hot fire testing is planned using these techniques in the future.

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.

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

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

NTR engines have continued as a main stream based on the mature technology. The typical core design of the NERVA derived engines uses hexagonal shaped fuel elements with circular cooling channels and structural tie-tube elements for supporting the fuel elements, housing moderator and regeneratively cooling the moderator. The state-of-the-art NTR designs mostly use a fast or epithermal neutron spectrum core utilizing a HEU fuel to make a high power reactor with small and simple core geometry. Nuclear propulsion is the most promising and viable option to achieve challenging deep space missions. Particularly, the attractions of a NTR include excellent thrust and propellant efficiency, bimodal capability, proven technology, and safe and reliable performance. The KANUTER-HEU and -LEU are the innovative and futuristic NTR engines to reduce the reactor size and to implement a LEU fuel in the reactor by using thermal neutron spectrum. The KANUTERs have some features in the reactor design such as the integrated fuel element and the regeneratively cooling channels to increase room for moderator and heat transfer in the core, and ensuing rocket performance. To study feasible design points in terms of thermo-hydraulics and to estimate rocket performance of the KANUTERs, the NSES is under development. The model of the NSES currently focuses on thermo-hydraulic analysis of the peculiar and complex EHTGR design during the propulsion mode in steady-state. The results indicate comparable performance for future applications, even though it uses the heavier LEU fuel. In future, the NSES will be modified to obtain temperature distribution of the entire reactor components and then more extensive design analysis of neutronics, thermohydraulics and their coupling will be conducted to validate design feasibility and to optimize the reactor design enhancing the rocket performance.

Estimates of the radiation environment for a nuclear rocket engine

International Nuclear Information System (INIS)

Courtney, J.C.; Manohara, H.M.; Williams, M.L.

1992-01-01

Ambitious missions in deep space, such as manned expeditions to Mars, require nuclear propulsion if they are to be accomplished in a reasonable length of time. Current technology is adequate to support the use of nuclear fission as a source of energy for propulsion; however, problems associated with neutrons and gammas leaking from the rocket engine must be addressed. Before manned or unmanned space flights are attempted, an extensive ground test program on the rocket engine must be completed. This paper compares estimated radiation levels and nuclear heating rates in and around the rocket engine for both a ground test and space environments

Dual Expander Cycle Rocket Engine with an Intermediate, Closed-cycle Heat Exchanger

Science.gov (United States)

Greene, William D. (Inventor)

2008-01-01

A dual expander cycle (DEC) rocket engine with an intermediate closed-cycle heat exchanger is provided. A conventional DEC rocket engine has a closed-cycle heat exchanger thermally coupled thereto. The heat exchanger utilizes heat extracted from the engine's fuel circuit to drive the engine's oxidizer turbomachinery.

Rocket Based Combined Cycle (RBCC) engine inlet

Science.gov (United States)

2004-01-01

Pictured is a component of the Rocket Based Combined Cycle (RBCC) engine. This engine was designed to ultimately serve as the near term basis for Two Stage to Orbit (TSTO) air breathing propulsion systems and ultimately a Single Stage to Orbit (SSTO) air breathing propulsion system.

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.

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.

46 CFR 119.420 - Engine cooling.

Science.gov (United States)

2010-10-01

... 46 Shipping 4 2010-10-01 2010-10-01 false Engine cooling. 119.420 Section 119.420 Shipping COAST... Machinery Requirements § 119.420 Engine cooling. (a) Except as otherwise provided in paragraph (b) of this section, all engines must be water cooled and meet the requirements of this paragraph. (1) The engine head...

Multivariable optimization of liquid rocket engines using particle swarm algorithms

Science.gov (United States)

Jones, Daniel Ray

Liquid rocket engines are highly reliable, controllable, and efficient compared to other conventional forms of rocket propulsion. As such, they have seen wide use in the space industry and have become the standard propulsion system for launch vehicles, orbit insertion, and orbital maneuvering. Though these systems are well understood, historical optimization techniques are often inadequate due to the highly non-linear nature of the engine performance problem. In this thesis, a Particle Swarm Optimization (PSO) variant was applied to maximize the specific impulse of a finite-area combustion chamber (FAC) equilibrium flow rocket performance model by controlling the engine's oxidizer-to-fuel ratio and de Laval nozzle expansion and contraction ratios. In addition to the PSO-controlled parameters, engine performance was calculated based on propellant chemistry, combustion chamber pressure, and ambient pressure, which are provided as inputs to the program. The performance code was validated by comparison with NASA's Chemical Equilibrium with Applications (CEA) and the commercially available Rocket Propulsion Analysis (RPA) tool. Similarly, the PSO algorithm was validated by comparison with brute-force optimization, which calculates all possible solutions and subsequently determines which is the optimum. Particle Swarm Optimization was shown to be an effective optimizer capable of quick and reliable convergence for complex functions of multiple non-linear variables.

Hybrid rocket engine, theoretical model and experiment

Science.gov (United States)

Chelaru, Teodor-Viorel; Mingireanu, Florin

2011-06-01

The purpose of this paper is to build a theoretical model for the hybrid rocket engine/motor and to validate it using experimental results. The work approaches the main problems of the hybrid motor: the scalability, the stability/controllability of the operating parameters and the increasing of the solid fuel regression rate. At first, we focus on theoretical models for hybrid rocket motor and compare the results with already available experimental data from various research groups. A primary computation model is presented together with results from a numerical algorithm based on a computational model. We present theoretical predictions for several commercial hybrid rocket motors, having different scales and compare them with experimental measurements of those hybrid rocket motors. Next the paper focuses on tribrid rocket motor concept, which by supplementary liquid fuel injection can improve the thrust controllability. A complementary computation model is also presented to estimate regression rate increase of solid fuel doped with oxidizer. Finally, the stability of the hybrid rocket motor is investigated using Liapunov theory. Stability coefficients obtained are dependent on burning parameters while the stability and command matrixes are identified. The paper presents thoroughly the input data of the model, which ensures the reproducibility of the numerical results by independent researchers.

46 CFR 182.420 - Engine cooling.

Science.gov (United States)

2010-10-01

... 46 Shipping 7 2010-10-01 2010-10-01 false Engine cooling. 182.420 Section 182.420 Shipping COAST...) MACHINERY INSTALLATION Specific Machinery Requirements § 182.420 Engine cooling. (a) Except as otherwise provided in paragraphs (b), (c), (d), and (e) of this section, all engines must be water cooled and meet...

14 CFR 33.21 - Engine cooling.

Science.gov (United States)

2010-01-01

... 14 Aeronautics and Space 1 2010-01-01 2010-01-01 false Engine cooling. 33.21 Section 33.21 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: AIRCRAFT ENGINES Design and Construction; General § 33.21 Engine cooling. Engine design and...

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.

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...

Transient Three-Dimensional Analysis of Nozzle Side Load in Regeneratively Cooled Engines

Science.gov (United States)

Wang, Ten-See

2005-01-01

Three-dimensional numerical investigations on the start-up side load physics for a regeneratively cooled, high-aspect-ratio nozzle were performed. The objectives of this study are to identify the three-dimensional side load physics and to compute the associated aerodynamic side load using an anchored computational methodology. The computational methodology is based on an unstructured-grid, pressure-based computational fluid dynamics formulation, and a transient inlet condition based on an engine system simulation. Computations were performed for both the adiabatic and cooled walls in order to understand the effect of boundary conditions. Finite-rate chemistry was used throughout the study so that combustion effect is always included. The results show that three types of shock evolution are responsible for side loads: generation of combustion wave; transitions among free-shock separation, restricted-shock separation, and simultaneous free-shock and restricted shock separations; along with oscillation of shocks across the lip. Wall boundary conditions drastically affect the computed side load physics: the adiabatic nozzle prefers free-shock separation while the cooled nozzle favors restricted-shock separation, resulting in higher peak side load for the cooled nozzle than that of the adiabatic nozzle. By comparing the computed physics with those of test observations, it is concluded that cooled wall is a more realistic boundary condition, and the oscillation of the restricted-shock separation flow pattern across the lip along with its associated tangential shock motion are the dominant side load physics for a regeneratively cooled, high aspect-ratio rocket engine.

Nuclear thermal rocket engine operation and control

International Nuclear Information System (INIS)

Gunn, S.V.; Savoie, M.T.; Hundal, R.

1993-06-01

The operation of a typical Rover/Nerva-derived nuclear thermal rocket (NTR) engine is characterized and the control requirements of the NTR are defined. A rationale for the selection of a candidate diverse redundant NTR engine control system is presented and the projected component operating requirements are related to the state of the art of candidate components and subsystems. The projected operational capabilities of the candidate system are delineated for the startup, full-thrust, shutdown, and decay heat removal phases of the engine operation. 9 refs

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

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...

Nuclear rocket engine reactor

Energy Technology Data Exchange (ETDEWEB)

Lanin, Anatoly

2013-07-01

Covers a new technology of nuclear reactors and the related materials aspects. Integrates physics, materials science and engineering Serves as a basic book for nuclear engineers and nuclear physicists. 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.

Improving of Hybrid Rocket Engine on the Basis of Optimizing Design Fuel Grain

Science.gov (United States)

Oriekov, K. M.; Ushkin, M. P.

2015-09-01

This article examines the processes intrachamber in hybrid rocket engine (HRE) and the comparative assessment of the use of solid rocket motors (SRM) and HRE for meteorological rockets with a mass of payload of the 364 kg. Results of the research showed the possibility of a significant increase in the ballistic effectiveness of meteorological rocket.

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.

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

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

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...

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...

Computational Fluid Dynamics Analysis Method Developed for Rocket-Based Combined Cycle Engine Inlet

Science.gov (United States)

1997-01-01

Renewed interest in hypersonic propulsion systems has led to research programs investigating combined cycle engines that are designed to operate efficiently across the flight regime. The Rocket-Based Combined Cycle Engine is a propulsion system under development at the NASA Lewis Research Center. This engine integrates a high specific impulse, low thrust-to-weight, airbreathing engine with a low-impulse, high thrust-to-weight rocket. From takeoff to Mach 2.5, the engine operates as an air-augmented rocket. At Mach 2.5, the engine becomes a dual-mode ramjet; and beyond Mach 8, the rocket is turned back on. One Rocket-Based Combined Cycle Engine variation known as the "Strut-Jet" concept is being investigated jointly by NASA Lewis, the U.S. Air Force, Gencorp Aerojet, General Applied Science Labs (GASL), and Lockheed Martin Corporation. Work thus far has included wind tunnel experiments and computational fluid dynamics (CFD) investigations with the NPARC code. The CFD method was initiated by modeling the geometry of the Strut-Jet with the GRIDGEN structured grid generator. Grids representing a subscale inlet model and the full-scale demonstrator geometry were constructed. These grids modeled one-half of the symmetric inlet flow path, including the precompression plate, diverter, center duct, side duct, and combustor. After the grid generation, full Navier-Stokes flow simulations were conducted with the NPARC Navier-Stokes code. The Chien low-Reynolds-number k-e turbulence model was employed to simulate the high-speed turbulent flow. Finally, the CFD solutions were postprocessed with a Fortran code. This code provided wall static pressure distributions, pitot pressure distributions, mass flow rates, and internal drag. These results were compared with experimental data from a subscale inlet test for code validation; then they were used to help evaluate the demonstrator engine net thrust.

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,...

Designing Liquid Rocket Engine Injectors for Performance, Stability, and Cost

Science.gov (United States)

Westra, Douglas G.; West, Jeffrey S.

2014-01-01

NASA is developing the Space Launch System (SLS) for crewed exploration missions beyond low Earth orbit. Marshall Space Flight Center (MSFC) is designing rocket engines for the SLS Advanced Booster (AB) concepts being developed to replace the Shuttle-derived solid rocket boosters. One AB concept uses large, Rocket-Propellant (RP)-fueled engines that pose significant design challenges. The injectors for these engines require high performance and stable operation while still meeting aggressive cost reduction goals for access to space. Historically, combustion stability problems have been a critical issue for such injector designs. Traditional, empirical injector design tools and methodologies, however, lack the ability to reliably predict complex injector dynamics that often lead to combustion stability. Reliance on these tools alone would likely result in an unaffordable test-fail-fix cycle for injector development. Recently at MSFC, a massively parallel computational fluid dynamics (CFD) program was successfully applied in the SLS AB injector design process. High-fidelity reacting flow simulations were conducted for both single-element and seven-element representations of the full-scale injector. Data from the CFD simulations was then used to significantly augment and improve the empirical design tools, resulting in a high-performance, stable injector design.

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.

Computational Fluid Dynamic (CFD) analysis of axisymmetric plume and base flow of film/dump cooled rocket nozzle

Science.gov (United States)

Tucker, P. K.; Warsi, S. A.

1993-01-01

Film/dump cooling a rocket nozzle with fuel rich gas, as in the National Launch System (NLS) Space Transportation Main Engine (STME), adds potential complexities for integrating the engine with the vehicle. The chief concern is that once the film coolant is exhausted from the nozzle, conditions may exist during flight for the fuel-rich film gases to be recirculated to the vehicle base region. The result could be significantly higher base temperatures than would be expected from a regeneratively cooled nozzle. CFD analyses were conduced to augment classical scaling techniques for vehicle base environments. The FDNS code with finite rate chemistry was used to simulate a single, axisymmetric STME plume and the NLS base area. Parallel calculations were made of the Saturn V S-1 C/F1 plume base area flows. The objective was to characterize the plume/freestream shear layer for both vehicles as inputs for scaling the S-C/F1 flight data to NLS/STME conditions. The code was validated on high speed flows with relevant physics. This paper contains the calculations for the NLS/STME plume for the baseline nozzle and a modified nozzle. The modified nozzle was intended to reduce the fuel available for recirculation to the vehicle base region. Plumes for both nozzles were calculated at 10kFT and 50kFT.

NASA's Hydrogen Outpost: The Rocket Systems Area at Plum Brook Station

Science.gov (United States)

Arrighi, Robert S.

2016-01-01

"There was pretty much a general knowledge about hydrogen and its capabilities," recalled former researcher Robert Graham. "The question was, could you use it in a rocket engine? Do we have the technology to handle it? How will it cool? Will it produce so much heat release that we can't cool the engine? These were the questions that we had to address." The National Aeronautics and Space Administration's (NASA) Glenn Research Center, referred to historically as the Lewis Research Center, made a concerted effort to answer these and related questions in the 1950s and 1960s. The center played a critical role transforming hydrogen's theoretical potential into a flight-ready propellant. Since then NASA has utilized liquid hydrogen to send humans and robots to the Moon, propel dozens of spacecraft across the universe, orbit scores of satellite systems, and power 135 space shuttle flights. Rocket pioneers had recognized hydrogen's potential early on, but its extremely low boiling temperature and low density made it impracticable as a fuel. The Lewis laboratory first demonstrated that liquid hydrogen could be safely utilized in rocket and aircraft propulsion systems, then perfected techniques to store, pump, and cleanly burn the fuel, as well as use it to cool the engine. The Rocket Systems Area at Lewis's remote testing area, Plum Brook Station, played a little known, but important role in the center's hydrogen research efforts. This publication focuses on the activities at the Rocket Systems Area, but it also discusses hydrogen's role in NASA's space program and Lewis's overall hydrogen work. The Rocket Systems Area included nine physically modest test sites and three test stands dedicated to liquid-hydrogen-related research. In 1962 Cleveland Plain Dealer reporter Karl Abram claimed, "The rocket facility looks more like a petroleum refinery. Its test rigs sprout pipes, valves and tanks. During the night test runs, excess hydrogen is burned from special stacks in the best

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.

Working-cycle processes in solid-propellant rocket engines (Handbook). Rabochie protsessy v raketnykh dvigateliakh tverdogo topliva /Spravochnik/

Energy Technology Data Exchange (ETDEWEB)

Shishkov, A.A.; Panin, S.D.; Rumiantsev, B.V.

1989-01-01

Physical and mathematical models of processes taking place in solid-propellant rocket engines and gas generators are presented in a systematic manner. The discussion covers the main types of solid propellants, the general design and principal components of solid-propellant rocket engines, the combustion of a solid-propellant charge, thermodynamic calculation of combustion and outflow processes, and analysis of gasdynamic processes in solid-propellant rocket engines. 40 refs.

Rocket Based Combined Cycle (RBCC) Engine

Science.gov (United States)

2004-01-01

Pictured is an artist's concept of the Rocket Based Combined Cycle (RBCC) launch. The RBCC's overall objective is to provide a technology test bed to investigate critical technologies associated with opperational usage of these engines. The program will focus on near term technologies that can be leveraged to ultimately serve as the near term basis for Two Stage to Orbit (TSTO) air breathing propulsions systems and ultimately a Single Stage To Orbit (SSTO) air breathing propulsion system.

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.

Review on pressure swirl injector in liquid rocket engine

Science.gov (United States)

Kang, Zhongtao; Wang, Zhen-guo; Li, Qinglian; Cheng, Peng

2018-04-01

The pressure swirl injector with tangential inlet ports is widely used in liquid rocket engine. Commonly, this type of pressure swirl injector consists of tangential inlet ports, a swirl chamber, a converging spin chamber, and a discharge orifice. The atomization of the liquid propellants includes the formation of liquid film, primary breakup and secondary atomization. And the back pressure and temperature in the combustion chamber could have great influence on the atomization of the injector. What's more, when the combustion instability occurs, the pressure oscillation could further affects the atomization process. This paper reviewed the primary atomization and the performance of the pressure swirl injector, which include the formation of the conical liquid film, the breakup and atomization characteristics of the conical liquid film, the effects of the rocket engine environment, and the response of the injector and atomization on the pressure oscillation.

Optimization of the rocket mode trajectory in a rocket based combined cycle (RBCC) engine powered SSTO vehicle

Science.gov (United States)

Foster, Richard W.

1989-07-01

The application of rocket-based combined cycle (RBCC) engines to booster-stage propulsion, in combination with all-rocket second stages in orbital-ascent missions, has been studied since the mid-1960s; attention is presently given to the case of the 'ejector scramjet' RBCC configuration's application to SSTO vehicles. While total mass delivered to initial orbit is optimized at Mach 20, payload delivery capability to initial orbit optimizes at Mach 17, primarily due to the reduction of hydrogen fuel tankage structure, insulation, and thermal protection system weights.

Hybrid rocket engine research program at Ryerson University

Energy Technology Data Exchange (ETDEWEB)

Karpynczyk, J.; Greatrix, D.R. [Ryerson Polytechnic Univ., Toronto, ON (Canada). Dept. of Aerospace Engineering

2007-07-01

Hybrid rocket engines (HREs) are a combination of solid and liquid propellant rocket engine designs. A solid fuel grain is located in the main combustion chamber and nozzle aft, while a stored liquid or gaseous oxidizer source supplies the required oxygen content through a throttle valve, for combustion downstream in the main chamber. HREs have drawn significant interest in certain flight applications, as they can be advantageous in terms of cost, ease and safety in storage, controllability in flight, and availability of propellant constituents. Key factors that will lead to further practical usage of HREs for flight applications are their predictability and reproducibility of operational performance. This paper presented information on studies being conducted at Ryerson University aimed at analyzing and testing the performance of HREs. It discussed and illustrated the conventional HRE and analyzed engine performance considerations such as the fuel regression rate, mass flux about the fuel surface, burning rate, and zero transformation parameter. Other factors relating to HRE performance that were presented included induced forward and aft oxidizer flow swirl effects as a means for augmenting the fuel regression rate, stoichiometric grain length issues, and feed system stability. Last, the paper presented a simplified schematic diagram of a proposed thrust/test stand for HRE test firings. 2 refs., 3 figs.

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

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.

Reusable Rocket Engine Turbopump Health Management System

Science.gov (United States)

Surko, Pamela

1994-01-01

A health monitoring expert system software architecture has been developed to support condition-based health monitoring of rocket engines. Its first application is in the diagnosis decisions relating to the health of the high pressure oxidizer turbopump (HPOTP) of Space Shuttle Main Engine (SSME). The post test diagnostic system runs off-line, using as input the data recorded from hundreds of sensors, each running typically at rates of 25, 50, or .1 Hz. The system is invoked after a test has been completed, and produces an analysis and an organized graphical presentation of the data with important effects highlighted. The overall expert system architecture has been developed and documented so that expert modules analyzing other line replaceable units may easily be added. The architecture emphasizes modularity, reusability, and open system interfaces so that it may be used to analyze other engines as well.

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.

Combustion dynamics in cryogenic rocket engines: Research programme at DLR Lampoldshausen

Science.gov (United States)

Hardi, Justin S.; Traudt, Tobias; Bombardieri, Cristiano; Börner, Michael; Beinke, Scott K.; Armbruster, Wolfgang; Nicolas Blanco, P.; Tonti, Federica; Suslov, Dmitry; Dally, Bassam; Oschwald, Michael

2018-06-01

The Combustion Dynamics group in the Rocket Propulsion Department at the German Aerospace Center (DLR), Lampoldshausen, strives to advance the understanding of dynamic processes in cryogenic rocket engines. Leveraging the test facilities and experimental expertise at DLR Lampoldshausen, the group has taken a primarily experimental approach to investigating transient flows, ignition, and combustion instabilities for over one and a half decades. This article provides a summary of recent achievements, and an overview of current and planned research activities.

Analytical concepts for health management systems of liquid rocket engines

Science.gov (United States)

Williams, Richard; Tulpule, Sharayu; Hawman, Michael

1990-01-01

Substantial improvement in health management systems performance can be realized by implementing advanced analytical methods of processing existing liquid rocket engine sensor data. In this paper, such techniques ranging from time series analysis to multisensor pattern recognition to expert systems to fault isolation models are examined and contrasted. The performance of several of these methods is evaluated using data from test firings of the Space Shuttle main engines.

Axisymmetric Numerical Modeling of Pulse Detonation Rocket Engines

Science.gov (United States)

Morris, Christopher I.

2005-01-01

Pulse detonation rocket engines (PDREs) have generated research interest in recent years as a chemical propulsion system potentially offering improved performance and reduced complexity compared to conventional rocket engines. The detonative mode of combustion employed by these devices offers a thermodynamic advantage over the constant-pressure deflagrative combustion mode used in conventional rocket engines and gas turbines. However, while this theoretical advantage has spurred considerable interest in building PDRE devices, the unsteady blowdown process intrinsic to the PDRE has made realistic estimates of the actual propulsive performance problematic. The recent review article by Kailasanath highlights some of the progress that has been made in comparing the available experimental measurements with analytical and numerical models. In recent work by the author, a quasi-one-dimensional, finite rate chemistry CFD model was utilized to study the gasdynamics and performance characteristics of PDREs over a range of blowdown pressure ratios from 1-1000. Models of this type are computationally inexpensive, and enable first-order parametric studies of the effect of several nozzle and extension geometries on PDRE performance over a wide range of conditions. However, the quasi-one-dimensional approach is limited in that it cannot properly capture the multidimensional blast wave and flow expansion downstream of the PDRE, nor can it resolve nozzle flow separation if present. Moreover, the previous work was limited to single-pulse calculations. In this paper, an axisymmetric finite rate chemistry model is described and utilized to study these issues in greater detail. Example Mach number contour plots showing the multidimensional blast wave and nozzle exhaust plume are shown. The performance results are compared with the quasi-one-dimensional results from the previous paper. Both Euler and Navier-Stokes solutions are calculated in order to determine the effect of viscous

Space shuttle with common fuel tank for liquid rocket booster and main engines (supertanker space shuttle)

Science.gov (United States)

Thorpe, Douglas G.

1991-01-01

An operation and schedule enhancement is shown that replaces the four-body cluster (Space Shuttle Orbiter (SSO), external tank, and two solid rocket boosters) with a simpler two-body cluster (SSO and liquid rocket booster/external tank). At staging velocity, the booster unit (liquid-fueled booster engines and vehicle support structure) is jettisoned while the remaining SSO and supertank continues on to orbit. The simpler two-bodied cluster reduces the processing and stack time until SSO mate from 57 days (for the solid rocket booster) to 20 days (for the liquid rocket booster). The areas in which liquid booster systems are superior to solid rocket boosters are discussed. Alternative and future generation vehicles are reviewed to reveal greater performance and operations enhancements with more modifications to the current methods of propulsion design philosophy, e.g., combined cycle engines, and concentric propellant tanks.

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.

Rotary engine cooling system

Science.gov (United States)

Jones, Charles (Inventor); Gigon, Richard M. (Inventor); Blum, Edward J. (Inventor)

1985-01-01

A rotary engine has a substantially trochoidal-shaped housing cavity in which a rotor planetates. A cooling system for the engine directs coolant along a single series path consisting of series connected groups of passages. Coolant enters near the intake port, passes downwardly and axially through the cooler regions of the engine, then passes upwardly and axially through the hotter regions. By first flowing through the coolest regions, coolant pressure is reduced, thus reducing the saturation temperature of the coolant and thereby enhancing the nucleate boiling heat transfer mechanism which predominates in the high heat flux region of the engine during high power level operation.

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.

Using Innovative Technologies for Manufacturing Rocket Engine Hardware

Science.gov (United States)

Betts, E. M.; Eddleman, D. E.; Reynolds, D. C.; Hardin, N. A.

2011-01-01

Many of the manufacturing techniques that are currently used for rocket engine component production are traditional methods that have been proven through years of experience and historical precedence. As the United States enters into the next space age where new launch vehicles are being designed and propulsion systems are being improved upon, it is sometimes necessary to adopt innovative techniques for manufacturing hardware. With a heavy emphasis on cost reduction and improvements in manufacturing time, rapid manufacturing techniques such as Direct Metal Laser Sintering (DMLS) are being adopted and evaluated for their use on NASA s Space Launch System (SLS) upper stage engine, J-2X, with hopes of employing this technology on a wide variety of future projects. DMLS has the potential to significantly reduce the processing time and cost of engine hardware, while achieving desirable material properties by using a layered powder metal manufacturing process in order to produce complex part geometries. Marshall Space Flight Center (MSFC) has recently hot-fire tested a J-2X gas generator (GG) discharge duct that was manufactured using DMLS. The duct was inspected and proof tested prior to the hot-fire test. Using a workhorse gas generator (WHGG) test fixture at MSFC's East Test Area, the duct was subjected to extreme J-2X hot gas environments during 7 tests for a total of 537 seconds of hot-fire time. The duct underwent extensive post-test evaluation and showed no signs of degradation. DMLS manufacturing has proven to be a viable option for manufacturing rocket engine hardware, and further development and use of this manufacturing method is recommended.

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.

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

International Nuclear Information System (INIS)

Labib, Satira; King, Jeffrey

2015-01-01

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

High-speed schlieren imaging of rocket exhaust plumes

Science.gov (United States)

Coultas-McKenney, Caralyn; Winter, Kyle; Hargather, Michael

2016-11-01

Experiments are conducted to examine the exhaust of a variety of rocket engines. The rocket engines are mounted in a schlieren system to allow high-speed imaging of the engine exhaust during startup, steady state, and shutdown. A variety of rocket engines are explored including a research-scale liquid rocket engine, consumer/amateur solid rocket motors, and water bottle rockets. Comparisons of the exhaust characteristics, thrust and cost for this range of rockets is presented. The variety of nozzle designs, target functions, and propellant type provides unique variations in the schlieren imaging.

Status on Technology Development of Optic Fiber-Coupled Laser Ignition System for Rocket Engine Applications

Science.gov (United States)

Trinh, Huu P.; Early, Jim; Osborne, Robin; Thomas, Matthew; Bossard, John

2003-01-01

To pursue technology developments for future launch vehicles, NASA/Marshall Space Flight Center (MSFC) is examining vortex chamber concepts for liquid rocket engine applications. Past studies indicated that the vortex chamber schemes potentially have a number of advantages over conventional chamber methods. Due to the nature of the vortex flow, relatively cooler propellant streams tend to flow along the chamber wall. Hence, the thruster chamber can be operated without the need of any cooling techniques. This vortex flow also creates strong turbulence, which promotes the propellant mixing process. Consequently, the subject chamber concept: not only offer system simplicity, but also enhance the combustion performance. Test results have shown that chamber performance is markedly high even at a low chamber length-to-diameter ratio. This incentive can be translated to a convenience in the thrust chamber packaging.

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.

Two-dimensional motions of rockets

International Nuclear Information System (INIS)

Kang, Yoonhwan; Bae, Saebyok

2007-01-01

We analyse the two-dimensional motions of the rockets for various types of rocket thrusts, the air friction and the gravitation by using a suitable representation of the rocket equation and the numerical calculation. The slope shapes of the rocket trajectories are discussed for the three types of rocket engines. Unlike the projectile motions, the descending parts of the trajectories tend to be gentler and straighter slopes than the ascending parts for relatively large launching angles due to the non-vanishing thrusts. We discuss the ranges, the maximum altitudes and the engine performances of the rockets. It seems that the exponential fuel exhaustion can be the most potent engine for the longest and highest flights

Net-Shape HIP Powder Metallurgy Components for Rocket Engines

Science.gov (United States)

Bampton, Cliff; Goodin, Wes; VanDaam, Tom; Creeger, Gordon; James, Steve

2005-01-01

True net shape consolidation of powder metal (PM) by hot isostatic pressing (HIP) provides opportunities for many cost, performance and life benefits over conventional fabrication processes for large rocket engine structures. Various forms of selectively net-shape PM have been around for thirty years or so. However, it is only recently that major applications have been pursued for rocket engine hardware fabricated in the United States. The method employs sacrificial metallic tooling (HIP capsule and shaped inserts), which is removed from the part after HIP consolidation of the powder, by selective acid dissolution. Full exploitation of net-shape PM requires innovative approaches in both component design and materials and processing details. The benefits include: uniform and homogeneous microstructure with no porosity, irrespective of component shape and size; elimination of welds and the associated quality and life limitations; removal of traditional producibility constraints on design freedom, such as forgeability and machinability, and scale-up to very large, monolithic parts, limited only by the size of existing HIP furnaces. Net-shape PM HIP also enables fabrication of complex configurations providing additional, unique functionalities. The progress made in these areas will be described. Then critical aspects of the technology that still require significant further development and maturation will be discussed from the perspective of an engine systems builder and end-user of the technology.

Using Innovative Techniques for Manufacturing Rocket Engine Hardware

Science.gov (United States)

Betts, Erin M.; Reynolds, David C.; Eddleman, David E.; Hardin, Andy

2011-01-01

Many of the manufacturing techniques that are currently used for rocket engine component production are traditional methods that have been proven through years of experience and historical precedence. As we enter into a new space age where new launch vehicles are being designed and propulsion systems are being improved upon, it is sometimes necessary to adopt new and innovative techniques for manufacturing hardware. With a heavy emphasis on cost reduction and improvements in manufacturing time, manufacturing techniques such as Direct Metal Laser Sintering (DMLS) are being adopted and evaluated for their use on J-2X, with hopes of employing this technology on a wide variety of future projects. DMLS has the potential to significantly reduce the processing time and cost of engine hardware, while achieving desirable material properties by using a layered powder metal manufacturing process in order to produce complex part geometries. Marshall Space Flight Center (MSFC) has recently hot-fire tested a J-2X gas generator discharge duct that was manufactured using DMLS. The duct was inspected and proof tested prior to the hot-fire test. Using the Workhorse Gas Generator (WHGG) test setup at MSFC?s East Test Area test stand 116, the duct was subject to extreme J-2X gas generator environments and endured a total of 538 seconds of hot-fire time. The duct survived the testing and was inspected after the test. DMLS manufacturing has proven to be a viable option for manufacturing rocket engine hardware, and further development and use of this manufacturing method is recommended.

Numerical Simulation of Reactive Flows in Overexpanded Supersonic Nozzle with Film Cooling

Directory of Open Access Journals (Sweden)

Mohamed Sellam

2015-01-01

Full Text Available Reignition phenomena occurring in a supersonic nozzle flow may present a crucial safety issue for rocket propulsion systems. These phenomena concern mainly rocket engines which use H2 gas (GH2 in the film cooling device, particularly when the nozzle operates under over expanded flow conditions at sea level or at low altitudes. Consequently, the induced wall thermal loads can lead to the nozzle geometry alteration, which in turn, leads to the appearance of strong side loads that may be detrimental to the rocket engine structural integrity. It is therefore necessary to understand both aerodynamic and chemical mechanisms that are at the origin of these processes. This paper is a numerical contribution which reports results from CFD analysis carried out for supersonic reactive flows in a planar nozzle cooled with GH2 film. Like the experimental observations, CFD simulations showed their ability to highlight these phenomena for the same nozzle flow conditions. Induced thermal load are also analyzed in terms of cooling efficiency and the results already give an idea on their magnitude. It was also shown that slightly increasing the film injection pressure can avoid the reignition phenomena by moving the separation shock towards the nozzle exit section.

Numerical and experimental analysis of heat transfer in injector plate of hydrogen peroxide hybrid rocket motor

Science.gov (United States)

Cai, Guobiao; Li, Chengen; Tian, Hui

2016-11-01

This paper is aimed to analyze heat transfer in injector plate of hydrogen peroxide hybrid rocket motor by two-dimensional axisymmetric numerical simulations and full-scale firing tests. Long-time working, which is an advantage of hybrid rocket motor over conventional solid rocket motor, puts forward new challenges for thermal protection. Thermal environments of full-scale hybrid rocket motors designed for long-time firing tests are studied through steady-state coupled numerical simulations of flow field and heat transfer in chamber head. The motor adopts 98% hydrogen peroxide (98HP) oxidizer and hydroxyl-terminated poly-butadiene (HTPB) based fuel as the propellants. Simulation results reveal that flowing liquid 98HP in head oxidizer chamber could cool the injector plate of the motor. The cooling of 98HP is similar to the regenerative cooling in liquid rocket engines. However, the temperature of the 98HP in periphery portion of the head oxidizer chamber is higher than its boiling point. In order to prevent the liquid 98HP from unexpected decomposition, a thermal protection method for chamber head utilizing silica-phenolics annular insulating board is proposed. The simulation results show that the annular insulating board could effectively decrease the temperature of the 98HP in head oxidizer chamber. Besides, the thermal protection method for long-time working hydrogen peroxide hybrid rocket motor is verified through full-scale firing tests. The ablation of the insulating board in oxygen-rich environment is also analyzed.

Coolant Design System for Liquid Propellant Aerospike Engines

Science.gov (United States)

McConnell, Miranda; Branam, Richard

2015-11-01

Liquid propellant rocket engines burn at incredibly high temperatures making it difficult to design an effective coolant system. These particular engines prove to be extremely useful by powering the rocket with a variable thrust that is ideal for space travel. When combined with aerospike engine nozzles, which provide maximum thrust efficiency, this class of rockets offers a promising future for rocketry. In order to troubleshoot the problems that high combustion chamber temperatures pose, this research took a computational approach to heat analysis. Chambers milled into the combustion chamber walls, lined by a copper cover, were tested for their efficiency in cooling the hot copper wall. Various aspect ratios and coolants were explored for the maximum wall temperature by developing our own MATLAB code. The code uses a nodal temperature analysis with conduction and convection equations and assumes no internal heat generation. This heat transfer research will show oxygen is a better coolant than water, and higher aspect ratios are less efficient at cooling. This project funded by NSF REU Grant 1358991.

Computational Fluid Dynamic Modeling of Rocket Based Combined Cycle Engine Flowfields

Science.gov (United States)

Daines, Russell L.; Merkle, Charles L.

1994-01-01

Computational Fluid Dynamic techniques are used to study the flowfield of a fixed geometry Rocket Based Combined Cycle engine operating in rocket ejector mode. Heat addition resulting from the combustion of injected fuel causes the subsonic engine flow to choke and go supersonic in the slightly divergent combustor-mixer section. Reacting flow computations are undertaken to predict the characteristics of solutions where the heat addition is determined by the flowfield. Here, adaptive gridding is used to improve resolution in the shear layers. Results show that the sonic speed is reached in the unheated portions of the flow first, while the heated portions become supersonic later. Comparison with results from another code show reasonable agreement. The coupled solutions show that the character of the combustion-based thermal choking phenomenon can be controlled reasonably well such that there is opportunity to optimize the length and expansion ratio of the combustor-mixer.

Using Innovative Technologies for Manufacturing and Evaluating Rocket Engine Hardware

Science.gov (United States)

Betts, Erin M.; Hardin, Andy

2011-01-01

Many of the manufacturing and evaluation techniques that are currently used for rocket engine component production are traditional methods that have been proven through years of experience and historical precedence. As we enter into a new space age where new launch vehicles are being designed and propulsion systems are being improved upon, it is sometimes necessary to adopt new and innovative techniques for manufacturing and evaluating hardware. With a heavy emphasis on cost reduction and improvements in manufacturing time, manufacturing techniques such as Direct Metal Laser Sintering (DMLS) and white light scanning are being adopted and evaluated for their use on J-2X, with hopes of employing both technologies on a wide variety of future projects. DMLS has the potential to significantly reduce the processing time and cost of engine hardware, while achieving desirable material properties by using a layered powdered metal manufacturing process in order to produce complex part geometries. The white light technique is a non-invasive method that can be used to inspect for geometric feature alignment. Both the DMLS manufacturing method and the white light scanning technique have proven to be viable options for manufacturing and evaluating rocket engine hardware, and further development and use of these techniques is recommended.

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

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...

CFD Analysis of Square Flow Channel in Thermal Engine Rocket Adventurer for Space Nuclear Application

Energy Technology Data Exchange (ETDEWEB)

Nam, S. H.; Suh, K. Y. [Seoul National University, Seoul (Korea, Republic of); Kang, S. G. [PHILOSOPHIA, Inc., Seoul (Korea, Republic of)

2008-10-15

Solar system exploration relying on chemical rockets suffers from long trip time and high cost. In this regard nuclear propulsion is an attractive option for space exploration. The performance of Nuclear Thermal Rocket (NTR) is more than twice that of the best chemical rocket. Resorting to the pure hydrogen (H{sub 2}) propellant the NTRs can possibly achieve as high as 1,000 s of specific impulse (I{sub sp}) representing the ratio of the thrust over the fuel consumption rate, as compared to only 425 s of H{sub 2}/O{sub 2} rockets. If we reflect on the mission to Mars, NTRs would reduce the round trip time to less than 300 days, instead of over 600 days with chemical rockets. This work presents CFD analysis of one Fuel Element (FE) of Thermal Engine Rocket Adventurer (TERA). In particular, one Square Flow Channel (SFC) is analyzed in Square Lattice Honeycomb (SLHC) fuel to examine the effects of mass flow rate on rocket performance.

CFD Analysis of Square Flow Channel in Thermal Engine Rocket Adventurer for Space Nuclear Application

International Nuclear Information System (INIS)

Nam, S. H.; Suh, K. Y.; Kang, S. G.

2008-01-01

Solar system exploration relying on chemical rockets suffers from long trip time and high cost. In this regard nuclear propulsion is an attractive option for space exploration. The performance of Nuclear Thermal Rocket (NTR) is more than twice that of the best chemical rocket. Resorting to the pure hydrogen (H 2 ) propellant the NTRs can possibly achieve as high as 1,000 s of specific impulse (I sp ) representing the ratio of the thrust over the fuel consumption rate, as compared to only 425 s of H 2 /O 2 rockets. If we reflect on the mission to Mars, NTRs would reduce the round trip time to less than 300 days, instead of over 600 days with chemical rockets. This work presents CFD analysis of one Fuel Element (FE) of Thermal Engine Rocket Adventurer (TERA). In particular, one Square Flow Channel (SFC) is analyzed in Square Lattice Honeycomb (SLHC) fuel to examine the effects of mass flow rate on rocket performance

Construction and design of solid-propellant rocket engines. Konstruktsiia i proektirovanie raketnykh dvigatelei tverdogo topliva

Energy Technology Data Exchange (ETDEWEB)

Fakhrutdinov, I.K.; Kotel' nikov, A.V.

1987-01-01

Methods for assessing the durability of different components of solid-propellant rocket engines are presented. The following aspects of engine development are discussed: task formulation, parameter calculation, construction scheme selection, materials, and durability assessment. 45 references.

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.

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...

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...

Development of CFD model for augmented core tripropellant rocket engine

Science.gov (United States)

Jones, Kenneth M.

1994-10-01

The Space Shuttle era has made major advances in technology and vehicle design to the point that the concept of a single-stage-to-orbit (SSTO) vehicle appears more feasible. NASA presently is conducting studies into the feasibility of certain advanced concept rocket engines that could be utilized in a SSTO vehicle. One such concept is a tripropellant system which burns kerosene and hydrogen initially and at altitude switches to hydrogen. This system will attain a larger mass fraction because LOX-kerosene engines have a greater average propellant density and greater thrust-to-weight ratio. This report describes the investigation to model the tripropellant augmented core engine. The physical aspects of the engine, the CFD code employed, and results of the numerical model for a single modular thruster are discussed.

Structural strengthening of rocket nozzle extension by means of laser metal deposition

Science.gov (United States)

Honoré, M.; Brox, L.; Hallberg, M.

2012-03-01

Commercial space operations strive to maximize the payload per launch in order to minimize the costs of each kg launched into orbit; this yields demand for ever larger launchers with larger, more powerful rocket engines. Volvo Aero Corporation in collaboration with Snecma and Astrium has designed and tested a new, upgraded Nozzle extension for the Vulcain 2 engine configuration, denoted Vulcain 2+ NE Demonstrator The manufacturing process for the welding of the sandwich wall and the stiffening structure is developed in close cooperation with FORCE Technology. The upgrade is intended to be available for future development programs for the European Space Agency's (ESA) highly successful commercial launch vehicle, the ARIANE 5. The Vulcain 2+ Nozzle Extension Demonstrator [1] features a novel, thin-sheet laser-welded configuration, with laser metal deposition built-up 3D-features for the mounting of stiffening structure, flanges and for structural strengthening, in order to cope with the extreme load- and thermal conditions, to which the rocket nozzle extension is exposed during launch of the 750 ton ARIANE 5 launcher. Several millimeters of material thickness has been deposited by laser metal deposition without disturbing the intricate flow geometry of the nozzle cooling channels. The laser metal deposition process has been applied on a full-scale rocket nozzle demonstrator, and in excess of 15 kilometers of filler wire has been successfully applied to the rocket nozzle. The laser metal deposition has proven successful in two full-throttle, full-scale tests, firing the rocket engine and nozzle in the ESA test facility P5 by DLR in Lampoldshausen, Germany.

A Film-Cooling CFD Bibliography: 1971–1996

Directory of Open Access Journals (Sweden)

D. M. Kercher

1998-01-01

Full Text Available After more than 25 years of three-dimensional film cooling experimental investigations, analytical correlations and modeling, film cooling utilizing computational fluid dynamics has emerged from a similar development-applications growth process into a near-attainable heat transfer engineering tool. Analytical applications include high temperature subsonic to hypersonic flow with complex wall-geometry coolant injection film performance analysis techniques spanning usage from gas turbines to rocket engines to scramjets. In recent years there has been significant development in increased computer power and modeling capacity, increasingly more complex and successful Navier-Stokes turbulence modeling techniques, innovative labor-saving meshing techniques, and more successful validation of experimental results. These combined innovations have continued to transition computational film cooling technology from the academic, government and commercial research and development environment to the industrial design-analysis environment. This bibliography is an openliterature reference resource whose papers collectively describe the continual emerging of numerical film cooling as a viable design tool for high temperature components.

Innovative nuclear thermal rocket concept utilizing LEU fuel for space application

International Nuclear Information System (INIS)

Nam, Seung Hyun; Venneri, Paolo; Choi, Jae Young; Jeong, Yong Hoon; Chang, Soon Heung

2015-01-01

Space is one of the best places for humanity to turn to keep learning and exploiting. A Nuclear Thermal Rocket (NTR) is a viable and more efficient option for human space exploration than the existing Chemical Rockets (CRs) which are highly inefficient for long-term manned missions such as to Mars and its satellites. NERVA derived NTR engines have been studied for the human missions as a mainstream in the United States of America (USA). Actually, the NERVA technology has already been developed and successfully tested since 1950s. The state-of-the-art technology is based on a Hydrogen gas (H_2) cooled high temperature reactor with solid core utilizing High-Enriched Uranium (HEU) fuel to reduce heavy metal mass and to use fast or epithermal neutron spectrums enabling simple core designs. However, even though the NTR designs utilizing HEU is the best option in terms of rocket performance, they inevitably provoke nuclear proliferation obstacles on all Research and Development (R and D) activities by civilians and non-nuclear weapon states, and its eventual commercialization. To surmount the security issue to use HEU fuel for a NTR, a concept of the innovative NTR engine, Korea Advanced NUclear Thermal Engine Rocket utilizing Low-Enriched Uranium fuel (KANUTER-LEU) is presented in this paper. The design goal of KANUTER-LEU is to make use of a LEU fuel for its compact reactor, but does not sacrifice the rocket performance relative to the traditional NTRs utilizing HEU. KANUTER-LEU mainly consists of a fission reactor utilizing H_2 propellant, a propulsion system and an optional Electricity Generation System as a bimodal engine. To implement LEU fuel for the reactor, the innovative engine adopts W-UO_2 CERMET fuel to drastically increase uranium density and thermal neutron spectrum to improve neutron economy in the core. The moderator and structural material selections also consider neutronic and thermo-physical characteristics to reduce non-fission neutron loss and

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

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.

Research Technology (ASTP) Rocket Based Combined Cycle (RBCC) Engine

Science.gov (United States)

2004-01-01

Pictured is an artist's concept of the Rocket Based Combined Cycle (RBCC) launch. The RBCC's overall objective is to provide a technology test bed to investigate critical technologies associated with opperational usage of these engines. The program will focus on near term technologies that can be leveraged to ultimately serve as the near term basis for Two Stage to Orbit (TSTO) air breathing propulsions systems and ultimately a Single Stage To Orbit (SSTO) air breathing propulsion system.

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

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.

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.

System Engineering and Technical Challenges Overcome in the J-2X Rocket Engine Development Project

Science.gov (United States)

Ballard, Richard O.

2012-01-01

Beginning in 2006, NASA initiated the J-2X engine development effort to develop an upper stage propulsion system to enable the achievement of the primary objectives of the Constellation program (CxP): provide continued access to the International Space Station following the retirement of the Space Station and return humans to the moon. The J-2X system requirements identified to accomplish this were very challenging and the time expended over the five years following the beginning of the J- 2X effort have been noteworthy in the development of innovations in both the fields for liquid rocket propulsion and system engineering.

Two-Dimensional Motions of Rockets

Science.gov (United States)

Kang, Yoonhwan; Bae, Saebyok

2007-01-01

We analyse the two-dimensional motions of the rockets for various types of rocket thrusts, the air friction and the gravitation by using a suitable representation of the rocket equation and the numerical calculation. The slope shapes of the rocket trajectories are discussed for the three types of rocket engines. Unlike the projectile motions, the…

Focused RBCC Experiments: Two-Rocket Configuration Experiments and Hydrocarbon/Oxygen Rocket Ejector Experiments

Science.gov (United States)

Santoro, Robert J.; Pal, Sibtosh

2003-01-01

This addendum report documents the results of two additional efforts for the Rocket Based Combined Cycle (RBCC) rocket-ejector mode research work carried out at the Penn State Propulsion Engineering Research Center in support of NASA s technology development efforts for enabling 3 d generation Reusable Launch Vehicles (RLV). The tasks reported here build on an earlier NASA MSFC funded research program on rocket ejector investigations. The first task investigated the improvements of a gaseous hydrogen/oxygen twin thruster RBCC rocket ejector system over a single rocket system. The second task investigated the performance of a hydrocarbon (liquid JP-7)/gaseous oxygen single thruster rocket-ejector system. To gain a systematic understanding of the rocket-ejector s internal fluid mechanic/combustion phenomena, experiments were conducted with both direct-connect and sea-level static diffusion and afterburning (DAB) configurations for a range of rocket operating conditions. For all experimental conditions, overall system performance was obtained through global measurements of wall static pressure profiles, heat flux profiles and engine thrust. Detailed mixing and combustion information was obtained through Raman spectroscopy measurements of major species (gaseous oxygen, hydrogen, nitrogen and water vapor) for the gaseous hydrogen/oxygen rocket ejector experiments.

Radiological effluents released from nuclear rocket and ramjet engine tests at the Nevada Test Site 1959 through 1969: Fact Book

Energy Technology Data Exchange (ETDEWEB)

Friesen, H.N.

1995-06-01

Nuclear rocket and ramjet engine tests were conducted on the Nevada Test Site (NTS) in Area 25 and Area 26, about 80 miles northwest of Las Vegas, Nevada, from July 1959 through September 1969. This document presents a brief history of the nuclear rocket engine tests, information on the off-site radiological monitoring, and descriptions of the tests.

Studies of an extensively axisymmetric rocket based combined cycle (RBCC) engine powered single-stage-to-orbit (SSTO) vehicle

Science.gov (United States)

Foster, Richard W.; Escher, William J. D.; Robinson, John W.

1989-01-01

The present comparative performance study has established that rocket-based combined cycle (RBCC) propulsion systems, when incorporated by essentially axisymmetric SSTO launch vehicle configurations whose conical forebody maximizes both capture-area ratio and total capture area, are capable of furnishing payload-delivery capabilities superior to those of most multistage, all-rocket launchers. Airbreathing thrust augmentation in the rocket-ejector mode of an RBCC powerplant is noted to make a major contribution to final payload capability, by comparison to nonair-augmented rocket engine propulsion systems.

Modeling of Rocket Fuel Heating and Cooling Processes in the Interior Receptacle Space of Ground-Based Systems

Directory of Open Access Journals (Sweden)

K. I. Denisova

2016-01-01

Full Text Available The propellant to fill the fuel tanks of the spacecraft, upper stages, and space rockets on technical and ground-based launch sites before fueling should be prepared to ensure many of its parameters, including temperature, in appropriate condition. Preparation of fuel temperature is arranged through heating and cooling the rocket propellants (RP in the tanks of fueling equipment. Processes of RP temperature preparation are the most energy-intensive and timeconsuming ones, which require that a choice of sustainable technologies and modes of cooling (heating RP provided by the ground-based equipment has been made through modeling of the RP [1] temperature preparation processes at the stage of design and operation of the groundbased fueling equipment.The RP temperature preparation in the tanks of the ground-based systems can be provided through the heat-exchangers built-in the internal space and being external with respect to the tank in which antifreeze, air or liquid nitrogen may be used as the heat transfer media. The papers [1-12], which note a promising use of the liquid nitrogen to cool PR, present schematic diagrams and modeling systems for the RP temperature preparation in the fueling equipment of the ground-based systems.We consider the RP temperature preparation using heat exchangers to be placed directly in RP tanks. Feeding the liquid nitrogen into heat exchanger with the antifreeze provides the cooling mode of PR while a heated air fed there does that of heating. The paper gives the systems of equations and results of modeling the processes of RP temperature preparation, and its estimated efficiency.The systems of equations of cooling and heating RP are derived on the assumption that the heat exchange between the fuel and the antifreeze, as well as between the storage tank and the environment is quasi-stationary.The paper presents calculation results of the fuel temperature in the tank, and coolant temperature in the heat exchanger, as

Cooled Ceramic Matrix Composite Propulsion Structures Demonstrated

Science.gov (United States)

Jaskowiak, Martha H.; Dickens, Kevin W.

2005-01-01

NASA's Next Generation Launch Technology (NGLT) Program has successfully demonstrated cooled ceramic matrix composite (CMC) technology in a scramjet engine test. This demonstration represented the world s largest cooled nonmetallic matrix composite panel fabricated for a scramjet engine and the first cooled nonmetallic composite to be tested in a scramjet facility. Lightweight, high-temperature, actively cooled structures have been identified as a key technology for enabling reliable and low-cost space access. Tradeoff studies have shown this to be the case for a variety of launch platforms, including rockets and hypersonic cruise vehicles. Actively cooled carbon and CMC structures may meet high-performance goals at significantly lower weight, while improving safety by operating with a higher margin between the design temperature and material upper-use temperature. Studies have shown that using actively cooled CMCs can reduce the weight of the cooled flow-path component from 4.5 to 1.6 lb/sq ft and the weight of the propulsion system s cooled surface area by more than 50 percent. This weight savings enables advanced concepts, increased payload, and increased range. The ability of the cooled CMC flow-path components to operate over 1000 F hotter than the state-of-the-art metallic concept adds system design flexibility to space-access vehicle concepts. Other potential system-level benefits include smaller fuel pumps, lower part count, lower cost, and increased operating margin.

Analytical Modelling of the Effects of Different Gas Turbine Cooling Techniques on Engine Performance =

Science.gov (United States)

Uysal, Selcuk Can

In this research, MATLAB SimulinkRTM was used to develop a cooled engine model for industrial gas turbines and aero-engines. The model consists of uncooled on-design, mean-line turbomachinery design and a cooled off-design analysis in order to evaluate the engine performance parameters by using operating conditions, polytropic efficiencies, material information and cooling system details. The cooling analysis algorithm involves a 2nd law analysis to calculate losses from the cooling technique applied. The model is used in a sensitivity analysis that evaluates the impacts of variations in metal Biot number, thermal barrier coating Biot number, film cooling effectiveness, internal cooling effectiveness and maximum allowable blade temperature on main engine performance parameters of aero and industrial gas turbine engines. The model is subsequently used to analyze the relative performance impact of employing Anti-Vortex Film Cooling holes (AVH) by means of data obtained for these holes by Detached Eddy Simulation-CFD Techniques that are valid for engine-like turbulence intensity conditions. Cooled blade configurations with AVH and other different external cooling techniques were used in a performance comparison study. (Abstract shortened by ProQuest.).

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.

Optical measurements in rocket engine liquid sprays

Science.gov (United States)

Feikema, Douglas A.

1994-01-01

The performance of liquid propellant rocket engines is dependent upon many elements of the entire system. One of the most fundamental and most critical is the performance of the injector elements. Their characterization is an important part of the development of combustion devices. Optical measurements within these environments have proven to be invaluable tools in quantifying the physical environment of two phase flows. The effort reported herein involves the measurement of drop velocity, drop size, and most importantly mass flux using Phase-Doppler Particle Anemometry within a spray generated by a single swirl injector element operating in atmospheric pressure conditions. The mass flux has been determined and validated by mechanical patternation methods and by profile integration of the mass flux.

Evaluation of undeveloped rocket engine cycle applications to advanced transportation

Science.gov (United States)

1990-01-01

Undeveloped pump-fed, liquid propellant rocket engine cycles were assessed and evaluated for application to Next Manned Transportation System (NMTS) vehicles, which would include the evolving Space Transportation System (STS Evolution), the Personnel Launch System (PLS), and the Advanced Manned Launch System (AMLS). Undeveloped engine cycles selected for further analysis had potential for increased reliability, more maintainability, reduced cost, and improved (or possibly level) performance when compared to the existing SSME and proposed STME engines. The split expander (SX) cycle, the full flow staged combustion (FFSC) cycle, and a hybrid version of the FFSC, which has a LOX expander drive for the LOX pump, were selected for definition and analysis. Technology requirements and issues were identified and analyses of vehicle systems weight deltas using the SX and FFSC cycles in AMLS vehicles were performed. A strawman schedule and cost estimate for FFSC subsystem technology developments and integrated engine system demonstration was also provided.

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

Design of a 2000 lbf LOX/LCH4 Throttleable Rocket Engine for a Vertical Lander

Science.gov (United States)

Lopez, Israel

Liquid oxygen (LOX) and liquid methane (LCH4) has been recognized as an attractive rocket propellant combination because of its in-situ resource utilization (ISRU) capabilities, namely in Mars. ISRU would allow launch vehicles to carry greater payloads and promote missions to Mars. This has led to an increasing interest to develop spacecraft technologies that employ this propellant combination. The UTEP Center for Space Exploration and Technology Research (cSETR) has focused part of its research efforts to developing LOX/LCH4 systems. One of those projects includes the development of a vertical takeoff and landing vehicle called JANUS. This vehicle will employ a LOX/LCH 4 propulsion system. The main propulsion engine is called CROME-X and is currently being developed as part of this project. This rocket engine will employ LOX/LCH4 propellants and is intended to operate from 2000-500 lbf thrust range. This thesis describes the design and development of CROME-X. Specifically, it describes the design process for the main engine components, the design criteria for each, and plans for future engine development.

Variable cooling circuit for thermoelectric generator and engine and method of control

Science.gov (United States)

Prior, Gregory P

2012-10-30

An apparatus is provided that includes an engine, an exhaust system, and a thermoelectric generator (TEG) operatively connected to the exhaust system and configured to allow exhaust gas flow therethrough. A first radiator is operatively connected to the engine. An openable and closable engine valve is configured to open to permit coolant to circulate through the engine and the first radiator when coolant temperature is greater than a predetermined minimum coolant temperature. A first and a second valve are controllable to route cooling fluid from the TEG to the engine through coolant passages under a first set of operating conditions to establish a first cooling circuit, and from the TEG to a second radiator through at least some other coolant passages under a second set of operating conditions to establish a second cooling circuit. A method of controlling a cooling circuit is also provided.

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

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.

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

Institute of Scientific and Technical Information of China (English)

Qiang WEI; Guozhu LIANG

2017-01-01

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 couplingof 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 offcenter 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.

Identification of Noise Sources During Rocket Engine Test Firings and a Rocket Launch Using a Microphone Phased-Array

Science.gov (United States)

Panda, Jayanta; Mosher, Robert N.; Porter, Barry J.

2013-01-01

A 70 microphone, 10-foot by 10-foot, microphone phased array was built for use in the harsh environment of rocket launches. The array was setup at NASA Wallops launch pad 0A during a static test firing of Orbital Sciences' Antares engines, and again during the first launch of the Antares vehicle. It was placed 400 feet away from the pad, and was hoisted on a scissor lift 40 feet above ground. The data sets provided unprecedented insight into rocket noise sources. The duct exit was found to be the primary source during the static test firing; the large amount of water injected beneath the nozzle exit and inside the plume duct quenched all other sources. The maps of the noise sources during launch were found to be time-dependent. As the engines came to full power and became louder, the primary source switched from the duct inlet to the duct exit. Further elevation of the vehicle caused spilling of the hot plume, resulting in a distributed noise map covering most of the pad. As the entire plume emerged from the duct, and the ondeck water system came to full power, the plume itself became the loudest noise source. These maps of the noise sources provide vital insight for optimization of sound suppression systems for future Antares launches.

Nuclear rockets

Energy Technology Data Exchange (ETDEWEB)

Sarram, M [Teheran Univ. (Iran). Inst. of Nuclear Science and Technology

1972-02-01

Nuclear energy has found many applications in space projects. This article deals with these applications. The first application is the use of nuclear energy for the production of electricity in space and the second main application is the use of nuclear energy for propulsion purposes in space flight. The main objective is to develop a 75000 pound thrust flight engine called NERVA by heating liquid hydrogen in a nuclear reactor. The paper describes in detail the salient features of the NERVA rocket as well as its comparison with the conventional chemical rockets. It is shown that a nuclear rocket using liquid hydrogen as medium is at least 85% more efficient as compared with the chemical rockets such as those used for the APOLLO moon flight.

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.

Effects of rocket engines on laser during lunar landing

International Nuclear Information System (INIS)

Wan, Xiong; Shu, Rong; Huang, Genghua

2013-01-01

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

Engine Cycle Analysis of Air Breathing Microwave Rocket with Reed Valves

International Nuclear Information System (INIS)

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

2011-01-01

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.

Investigation of Water-spray Cooling of Turbine Blades in a Turbojet Engine

Science.gov (United States)

Freche, John C; Stelpflug, William J

1953-01-01

An analytical and experimental investigation was made with a J33-A-9 engine to determine the effectiveness of spray cooling as a means of increasing thrust by permitting engine operation at inlet-gas temperatures and speeds above rated. With the assumption of adequate spray cooling at a coolant-to-gas flow ratio of 3 percent, calculations for the sea-level static condition indicated a thrust may be achieved by engine operation at an inlet-gas temperature of 2000 degrees F and an overspeed of 10 percent. Of the water-injection configurations investigated experimentally, those located in the inner ring of the stator diaphragm provided the best cooling at rated engine speed.

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.

The XQC microcalorimeter sounding rocket: a stable LTD platform 30 seconds after rocket motor burnout

Energy Technology Data Exchange (ETDEWEB)

Porter, F.S. E-mail: frederick.s.porter@gsfc.nasa.gov; Almy, R.; Apodaca, E.; Figueroa-Feliciano, E.; Galeazzi, M.; Kelley, R.; McCammon, D.; Stahle, C.K.; Szymkowiak, A.E.; Sanders, W.T

2000-04-07

The XQC microcalorimeter sounding rocket experiment is designed to provide a stable thermal environment for an LTD detector system within 30 s of the burnout of its second stage rocket motor. The detector system used for this instrument is a 36-pixel microcalorimeter array operated at 60 mK with a single-stage adiabatic demagnetization refrigerator (ADR). The ADR is mounted on a space-pumped liquid helium tank with vapor cooled shields which is vibration isolated from the rocket structure. We present here some of the design and performance details of this mature LTD instrument, which has just completed its third suborbital flight.

The XQC microcalorimeter sounding rocket: a stable LTD platform 30 seconds after rocket motor burnout

International Nuclear Information System (INIS)

Porter, F.S.; Almy, R.; Apodaca, E.; Figueroa-Feliciano, E.; Galeazzi, M.; Kelley, R.; McCammon, D.; Stahle, C.K.; Szymkowiak, A.E.; Sanders, W.T.

2000-01-01

The XQC microcalorimeter sounding rocket experiment is designed to provide a stable thermal environment for an LTD detector system within 30 s of the burnout of its second stage rocket motor. The detector system used for this instrument is a 36-pixel microcalorimeter array operated at 60 mK with a single-stage adiabatic demagnetization refrigerator (ADR). The ADR is mounted on a space-pumped liquid helium tank with vapor cooled shields which is vibration isolated from the rocket structure. We present here some of the design and performance details of this mature LTD instrument, which has just completed its third suborbital flight

Towards Rocket Engine Components with Increased Strength and Robust Operating Characteristics

Science.gov (United States)

Marcu, Bogdan; Hadid, Ali; Lin, Pei; Balcazar, Daniel; Rai, Man Mohan; Dorney, Daniel J.

2005-01-01

High-energy rotating machines, powering liquid propellant rocket engines, are subject to various sources of high and low cycle fatigue generated by unsteady flow phenomena. Given the tremendous need for reliability in a sustainable space exploration program, a fundamental change in the design methodology for engine components is required for both launch and space based systems. A design optimization system based on neural-networks has been applied and demonstrated in the redesign of the Space Shuttle Main Engine (SSME) Low Pressure Oxidizer Turbo Pump (LPOTP) turbine nozzle. One objective of the redesign effort was to increase airfoil thickness and thus increase its strength while at the same time detuning the vane natural frequency modes from the vortex shedding frequency. The second objective was to reduce the vortex shedding amplitude. The third objective was to maintain this low shedding amplitude even in the presence of large manufacturing tolerances. All of these objectives were achieved without generating any detrimental effects on the downstream flow through the turbine, and without introducing any penalty in performance. The airfoil redesign and preliminary assessment was performed in the Exploration Technology Directorate at NASA ARC. Boeing/Rocketdyne and NASA MSFC independently performed final CFD assessments of the design. Four different CFD codes were used in this process. They include WIL DCA T/CORSAIR (NASA), FLUENT (commercial), TIDAL (Boeing Rocketdyne) and, a new family (AardvarWPhantom) of CFD analysis codes developed at NASA MSFC employing LOX fluid properties and a Generalized Equation Set formulation. Extensive aerodynamic performance analysis and stress analysis carried out at Boeing Rocketdyne and NASA MSFC indicate that the redesign objectives have been fully met. The paper presents the results of the assessment analysis and discusses the future potential of robust optimal design for rocket engine components.

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.

Nuclear rockets

International Nuclear Information System (INIS)

Sarram, M.

1972-01-01

Nuclear energy has found many applications in space projects. This article deals with these applications. The first application is the use of nuclear energy for the production of electricity in space and the second main application is the use of nuclear energy for propulsion purposes in space flight. The main objective is to develop a 75000 pound thrust flight engine call NERVA by heating liquid hydrogen, in a nuclear reactor, from 420F to 4000 0 F. The paper describes in detail the salient features of the NERVA rocket as well as its comparison with the conventional chemical rockets. It is shown that a nuclear rocket using liquid hydrogen as medium is at least 85% more efficient as compared with the chemical rockets such as those used for the APOLLO moon flight

Design and development of LH2 cooled rolling element radial bearings for the NERVA engine turbopump. Volume 3: Phase 2: Tests on build-ups 16, 17, and 18 at NRDS, Jackass Flats, Nevada, December 1971 - March 1972

Science.gov (United States)

Accinelli, J. B.; Koch, D. A.; Reuter, F.

1972-01-01

The use of liquid hydrogen to cool the rolling element radial bearings in the nuclear engine for rocket vehicles is discussed. The fifteen hour service life goal was obtained during the tests. The increase in bearing life was also considered to be produced by: (1) improvements in bearing material, (2) bearing retainer configuration and manufacturing changes, and (3) better control of operating parameters.

Analysis and simulation of mobile air conditioning system coupled with engine cooling system

International Nuclear Information System (INIS)

Qi, Zhao-gang; Chen, Jiang-ping; Chen, Zhi-jiu

2007-01-01

Many components of the mobile air conditioning system and engine cooling system are closely interrelated and make up the vehicle climate control system. In the present paper, a vehicle climate control system model including air conditioning system and engine cooling system has been proposed under different operational conditions. All the components have been modeled on the basis of experimental data. Based on the commercial software, a computer simulation procedure of the vehicle climate control system has been developed. The performance of the vehicle climate control system is simulated, and the calculational data have good agreement with experimental data. Furthermore, the vehicle climate control simulation results have been compared with an individual air conditioning system and engine cooling system. The influences between the mobile air conditioning system and the engine cooling system are discussed

Cooled airfoil in a turbine engine

Science.gov (United States)

Vitt, Paul H; Kemp, David A; Lee, Ching-Pang; Marra, John J

2015-04-21

An airfoil in a gas turbine engine includes an outer wall and an inner wall. The outer wall includes a leading edge, a trailing edge opposed from the leading edge in a chordal direction, a pressure side, and a suction side. The inner wall is coupled to the outer wall at a single chordal location and includes portions spaced from the pressure and suction sides of the outer wall so as to form first and second gaps between the inner wall and the respective pressure and suction sides. The inner wall defines a chamber therein and includes openings that provide fluid communication between the respective gaps and the chamber. The gaps receive cooling fluid that provides cooling to the outer wall as it flows through the gaps. The cooling fluid, after traversing at least substantial portions of the gaps, passes into the chamber through the openings in the inner wall.

A reliability as an independent variable (RAIV) methodology for optimizing test planning for liquid rocket engines

Science.gov (United States)

Strunz, Richard; Herrmann, Jeffrey W.

2011-12-01

The hot fire test strategy for liquid rocket engines has always been a concern of space industry and agency alike because no recognized standard exists. Previous hot fire test plans focused on the verification of performance requirements but did not explicitly include reliability as a dimensioning variable. The stakeholders are, however, concerned about a hot fire test strategy that balances reliability, schedule, and affordability. A multiple criteria test planning model is presented that provides a framework to optimize the hot fire test strategy with respect to stakeholder concerns. The Staged Combustion Rocket Engine Demonstrator, a program of the European Space Agency, is used as example to provide the quantitative answer to the claim that a reduced thrust scale demonstrator is cost beneficial for a subsequent flight engine development. Scalability aspects of major subsystems are considered in the prior information definition inside the Bayesian framework. The model is also applied to assess the impact of an increase of the demonstrated reliability level on schedule and affordability.

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.

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).

Study of solid rocket motor for space shuttle booster, volume 2, book 1

Science.gov (United States)

1972-01-01

The technical requirements for the solid propellant rocket engine to be used with the space shuttle orbiter are presented. The subjects discussed are: (1) propulsion system definition, (2) solid rocket engine stage design, (3) solid rocket engine stage recovery, (4) environmental effects, (5) manrating of the solid rocket engine stage, (6) system safety analysis, and (7) ground support equipment.

Analysis of film cooling in rocket nozzles

Science.gov (United States)

Woodbury, Keith A.

1992-01-01

Computational Fluid Dynamics (CFD) programs are customarily used to compute details of a flow field, such as velocity fields or species concentrations. Generally they are not used to determine the resulting conditions at a solid boundary such as wall shear stress or heat flux. However, determination of this information should be within the capability of a CFD code, as the code supposedly contains appropriate models for these wall conditions. Before such predictions from CFD analyses can be accepted, the credibility of the CFD codes upon which they are based must be established. This report details the progress made in constructing a CFD model to predict the heat transfer to the wall in a film cooled rocket nozzle. Specifically, the objective of this work is to use the NASA code FDNS to predict the heat transfer which will occur during the upcoming hot-firing of the Pratt & Whitney 40K subscale nozzle (1Q93). Toward this end, an M = 3 wall jet is considered, and the resulting heat transfer to the wall is computed. The values are compared against experimental data available in Reference 1. Also, FDNS's ability to compute heat flux in a reacting flow will be determined by comparing the code's predictions against calorimeter data from the hot firing of a 40K combustor. The process of modeling the flow of combusting gases through the Pratt & Whitney 40K subscale combustor and nozzle is outlined. What follows in this report is a brief description of the FDNS code, with special emphasis on how it handles solid wall boundary conditions. The test cases and some FDNS solution are presented next, along with comparison to experimental data. The process of modeling the flow through a chamber and a nozzle using the FDNS code will also be outlined.

Ozone Depletion Caused by Rocket Engine Emissions: A Fundamental Limit on the Scale and Viability of Space-Based Geoengineering Schemes

Science.gov (United States)

Ross, M. N.; Toohey, D.

2008-12-01

Emissions from solid and liquid propellant rocket engines reduce global stratospheric ozone levels. Currently ~ one kiloton of payloads are launched into earth orbit annually by the global space industry. Stratospheric ozone depletion from present day launches is a small fraction of the ~ 4% globally averaged ozone loss caused by halogen gases. Thus rocket engine emissions are currently considered a minor, if poorly understood, contributor to ozone depletion. Proposed space-based geoengineering projects designed to mitigate climate change would require order of magnitude increases in the amount of material launched into earth orbit. The increased launches would result in comparable increases in the global ozone depletion caused by rocket emissions. We estimate global ozone loss caused by three space-based geoengineering proposals to mitigate climate change: (1) mirrors, (2) sunshade, and (3) space-based solar power (SSP). The SSP concept does not directly engineer climate, but is touted as a mitigation strategy in that SSP would reduce CO2 emissions. We show that launching the mirrors or sunshade would cause global ozone loss between 2% and 20%. Ozone loss associated with an economically viable SSP system would be at least 0.4% and possibly as large as 3%. It is not clear which, if any, of these levels of ozone loss would be acceptable under the Montreal Protocol. The large uncertainties are mainly caused by a lack of data or validated models regarding liquid propellant rocket engine emissions. Our results offer four main conclusions. (1) The viability of space-based geoengineering schemes could well be undermined by the relatively large ozone depletion that would be caused by the required rocket launches. (2) Analysis of space- based geoengineering schemes should include the difficult tradeoff between the gain of long-term (~ decades) climate control and the loss of short-term (~ years) deep ozone loss. (3) The trade can be properly evaluated only if our

Low pressure cooling seal system for a gas turbine engine

Science.gov (United States)

Marra, John J

2014-04-01

A low pressure cooling system for a turbine engine for directing cooling fluids at low pressure, such as at ambient pressure, through at least one cooling fluid supply channel and into a cooling fluid mixing chamber positioned immediately downstream from a row of turbine blades extending radially outward from a rotor assembly to prevent ingestion of hot gases into internal aspects of the rotor assembly. The low pressure cooling system may also include at least one bleed channel that may extend through the rotor assembly and exhaust cooling fluids into the cooling fluid mixing chamber to seal a gap between rotational turbine blades and a downstream, stationary turbine component. Use of ambient pressure cooling fluids by the low pressure cooling system results in tremendous efficiencies by eliminating the need for pressurized cooling fluids for sealing this gap.

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.

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

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

Rocket-Based Combined Cycle Engine Technology Development: Inlet CFD Validation and Application

Science.gov (United States)

DeBonis, J. R.; Yungster, S.

1996-01-01

A CFD methodology has been developed for inlet analyses of Rocket-Based Combined Cycle (RBCC) Engines. A full Navier-Stokes analysis code, NPARC, was used in conjunction with pre- and post-processing tools to obtain a complete description of the flow field and integrated inlet performance. This methodology was developed and validated using results from a subscale test of the inlet to a RBCC 'Strut-Jet' engine performed in the NASA Lewis 1 x 1 ft. supersonic wind tunnel. Results obtained from this study include analyses at flight Mach numbers of 5 and 6 for super-critical operating conditions. These results showed excellent agreement with experimental data. The analysis tools were also used to obtain pre-test performance and operability predictions for the RBCC demonstrator engine planned for testing in the NASA Lewis Hypersonic Test Facility. This analysis calculated the baseline fuel-off internal force of the engine which is needed to determine the net thrust with fuel on.

Effect of buoyancy on fuel containment in an open-cycle gas-core nuclear rocket engine.

Science.gov (United States)

Putre, H. A.

1971-01-01

Analysis aimed at determining the scaling laws for the buoyancy effect on fuel containment in an open-cycle gas-core nuclear rocket engine, so conducted that experimental conditions can be related to engine conditions. The fuel volume fraction in a short coaxial flow cavity is calculated with a programmed numerical solution of the steady Navier-Stokes equations for isothermal, variable density fluid mixing. A dimensionless parameter B, called the Buoyancy number, was found to correlate the fuel volume fraction for large accelerations and various density ratios. This parameter has the value B = 0 for zero acceleration, and B = 350 for typical engine conditions.

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.

Investigation of Dual-Vortical-Flow Hybrid Rocket Engine without Flame Holding Mechanism

Directory of Open Access Journals (Sweden)

A. Lai

2018-01-01

Full Text Available A 250 kgf thrust hybrid rocket engine was designed, tested, and verified in this work. Due to the injection and flow pattern of this engine, this engine was named dual-vortical-flow engine. This propulsion system uses N2O as oxidizer and HDPE as fuel. This engine was numerically investigated using a CFD tool that can handle reacting flow with finite-rate chemistry and coupled with the real-fluid model. The engine was further verified via a hot-fire test for 12 s. The ground Isp of the engine was 232 s and 221 s for numerical and hot-fire tests, respectively. An oscillation frequency with an order of 100 Hz was observed in both numerical and hot-fire tests with less than 5% of pressure oscillation. Swirling pattern on the fuel surface was also observed in both numerical and hot-fire test, which proves that this swirling dual-vortical-flow engine works exactly as designed. The averaged regression rate of the fuel surface was found to be 0.6~0.8 mm/s at the surface of disk walls and 1.5~1.7 mm/s at the surface of central core of the fuel grain.

On the hydrodynamics of rocket propellant engine inducers and turbopumps

International Nuclear Information System (INIS)

D'Agostino, L

2013-01-01

The lecture presents an overview of some recent results of the work carried out at Alta on the hydrodynamic design and rotordynamic fluid forces of cavitating turbopumps for liquid propellant feed systems of modern rocket engines. The reduced order models recently developed for preliminary geometric definition and noncavitating performance prediction of tapered-hub axial inducers and centrifugal turbopumps are illustrated. The experimental characterization of the rotordynamic forces acting on a whirling four-bladed, tapered-hub, variable-pitch high-head inducer, under different load and cavitation conditions is presented. Future perspectives of the work to be carried out at Alta in this area of research are briefly illustrated

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.

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.

46 CFR 119.425 - Engine exhaust cooling.

Science.gov (United States)

2010-10-01

..., all engine exhaust pipes must be water cooled. (1) Vertical dry exhaust pipes are permissible if installed in compliance with §§ 116.405(c) and 116.970 of this chapter. (2) Horizontal dry exhaust pipes are...) They are installed in compliance with §§ 116.405(c) and 116.970 of this chapter. (b) The exhaust pipe...

AUTOMOTIVE DIESEL MAINTENANCE 1. UNIT IV, MAINTAINING THE COOLING SYSTEM--DETROIT DIESEL ENGINES.

Science.gov (United States)

Human Engineering Inst., Cleveland, OH.

THIS MODULE OF A 30-MODULE COURSE IS DESIGNED TO DEVELOP AN UNDERSTANDING OF THE OPERATION AND MAINTENANCE OF THE DIESEL ENGINE COOLING SYSTEM. TOPICS ARE PURPOSE OF THE COOLING SYSTEM, CARE MAINTENANCE OF THE COOLING SYSTEM, COOLING SYSTEM COMPONENTS, AND TROUBLESHOOTING TIPS. THE MODULE CONSISTS OF A SELF-INSTRUCTIONAL BRANCH PROGRAMED TRAINING…

Modeling and Simulation of Truck Engine Cooling System for Onboard Diagnosis

Institute of Scientific and Technical Information of China (English)

朱正礼; 张建武; 包继华

2004-01-01

A cooling system model of a selected internal combustion engine has been built for onboard diagnosis. The model uses driving cycle data available within the production Engine Control Module (ECM): vehicle speed, engine speed, and fuel flow rate for the given ambient temperature and pressure, etc. Based on the conservation laws for heat transfer and mass flow process, the mathematical descriptions for the components involved in the cooling circuit are obtained and all the components are integrated into a model on Matlab/Simulink platform. The model can simulate the characteristics of thermostat (e.g. time-lag, hysteresis effect).The changes of coolant temperature, heat transfer flow rate, and pressure at individual component site are also shown.

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.

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

MHD thrust vectoring of a rocket engine

Science.gov (United States)

Labaune, Julien; Packan, Denis; Tholin, Fabien; Chemartin, Laurent; Stillace, Thierry; Masson, Frederic

2016-09-01

In this work, the possibility to use MagnetoHydroDynamics (MHD) to vectorize the thrust of a solid propellant rocket engine exhaust is investigated. Using a magnetic field for vectoring offers a mass gain and a reusability advantage compared to standard gimbaled, elastomer-joint systems. Analytical and numerical models were used to evaluate the flow deviation with a 1 Tesla magnetic field inside the nozzle. The fluid flow in the resistive MHD approximation is calculated using the KRONOS code from ONERA, coupling the hypersonic CFD platform CEDRE and the electrical code SATURNE from EDF. A critical parameter of these simulations is the electrical conductivity, which was evaluated using a set of equilibrium calculations with 25 species. Two models were used: local thermodynamic equilibrium and frozen flow. In both cases, chlorine captures a large fraction of free electrons, limiting the electrical conductivity to a value inadequate for thrust vectoring applications. However, when using chlorine-free propergols with 1% in mass of alkali, an MHD thrust vectoring of several degrees was obtained.

Fuel/propellant mixing in an open-cycle gas core nuclear rocket engine

International Nuclear Information System (INIS)

Guo, X.; Wehrmeyer, J.A.

1997-01-01

A numerical investigation of the mixing of gaseous uranium and hydrogen inside an open-cycle gas core nuclear rocket engine (spherical geometry) is presented. The gaseous uranium fuel is injected near the centerline of the spherical engine cavity at a constant mass flow rate, and the hydrogen propellant is injected around the periphery of the engine at a five degree angle to the wall, at a constant mass flow rate. The main objective is to seek ways to minimize the mixing of uranium and hydrogen by choosing a suitable injector geometry for the mixing of light and heavy gas streams. Three different uranium inlet areas are presented, and also three different turbulent models (k-var-epsilon model, RNG k-var-epsilon model, and RSM model) are investigated. The commercial CFD code, FLUENT, is used to model the flow field. Uranium mole fraction, axial mass flux, and radial mass flux contours are obtained. copyright 1997 American Institute of Physics

Ricardo Dyrgalla (1910-1970), pioneer of rocket development in Argentina

Science.gov (United States)

de León, Pablo

2009-12-01

One of the most important developers of liquid propellant rocket engines in Argentina was Polish-born Ricardo Dyrgalla. Dyrgalla immigrated to Argentina from the United Kingdom in 1946, where he had been studying German weapons development at the end of the Second World War. A trained pilot and aeronautical engineer, he understood the intricacies of rocket propulsion and was eager to find practical applications to his recently gained knowledge. Dyrgalla arrived in Argentina during Juan Perón's first presidency, a time when technicians from all over Europe were being recruited to work in various projects for the recently created Argentine Air Force. Shortly after immigrating, Dyrgalla proposed to develop an advanced air-launched weapon, the Tábano, based on a rocket engine of his design, the AN-1. After a successful development program, the Tábano was tested between 1949 and 1951; however, the project was canceled by the government shortly after. Today, the AN-1 rocket engine is recognized as the first liquid propellant rocket to be developed in South America. Besides the AN-1, Dyrgalla also developed several other rockets systems in Argentina, including the PROSON, a solid-propellant rocket launcher developed by the Argentine Institute of Science and Technology for the Armed Forces (CITEFA). In the late 1960s, Dyrgalla and his family relocated to Brazil due mostly to the lack of continuation of rocket development in Argentina. There, he worked for the Institute of Aerospace Technology (ITA) until his untimely death in 1970. Ricardo Dyrgalla deserves to be recognized among the world's rocket pioneers and his contribution to the science and engineering of rocketry deserves a special place in the history of South America's rocketry and space flight advocacy programs.

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.

Airbreathing engine selection criteria for SSTO propulsion system

Science.gov (United States)

Ohkami, Yoshiaki; Maita, Masataka

1995-02-01

This paper presents airbreathing engine selection criteria to be applied to the propulsion system of a Single Stage To Orbit (SSTO). To establish the criteria, a relation among three major parameters, i.e., delta-V capability, weight penalty, and effective specific impulse of the engine subsystem, is derived as compared to these parameters of the LH2/LOX rocket engine. The effective specific impulse is a function of the engine I(sub sp) and vehicle thrust-to-drag ratio which is approximated by a function of the vehicle velocity. The weight penalty includes the engine dry weight, cooling subsystem weight. The delta-V capability is defined by the velocity region starting from the minimum operating velocity up to the maximum velocity. The vehicle feasibility is investigated in terms of the structural and propellant weights, which requires an iteration process adjusting the system parameters. The system parameters are computed by iteration based on the Newton-Raphson method. It has been concluded that performance in the higher velocity region is extremely important so that the airbreathing engines are required to operate beyond the velocity equivalent to the rocket engine exhaust velocity (approximately 4500 m/s).

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.

Ambient air cooling arrangement having a pre-swirler for gas turbine engine blade cooling

Science.gov (United States)

Lee, Ching-Pang; Tham, Kok-Mun; Schroeder, Eric; Meeroff, Jamie; Miller, Jr., Samuel R; Marra, John J

2015-01-06

A gas turbine engine including: an ambient-air cooling circuit (10) having a cooling channel (26) disposed in a turbine blade (22) and in fluid communication with a source (12) of ambient air: and an pre-swirler (18), the pre-swirler having: an inner shroud (38); an outer shroud (56); and a plurality of guide vanes (42), each spanning from the inner shroud to the outer shroud. Circumferentially adjacent guide vanes (46, 48) define respective nozzles (44) there between. Forces created by a rotation of the turbine blade motivate ambient air through the cooling circuit. The pre-swirler is configured to impart swirl to ambient air drawn through the nozzles and to direct the swirled ambient air toward a base of the turbine blade. The end walls (50, 54) of the pre-swirler may be contoured.

Simulation research on the effect of cooled EGR, supercharging and compression ratio on downsized SI engine knock

Science.gov (United States)

Shu, Gequn; Pan, Jiaying; Wei, Haiqiao; Shi, Ning

2013-03-01

Knock in spark-ignition(SI) engines severely limits engine performance and thermal efficiency. The researches on knock of downsized SI engine have mainly focused on structural design, performance optimization and advanced combustion modes, however there is little for simulation study on the effect of cooled exhaust gas recirculation(EGR) combined with downsizing technologies on SI engine performance. On the basis of mean pressure and oscillating pressure during combustion process, the effect of different levels of cooled EGR ratio, supercharging and compression ratio on engine dynamic and knock characteristic is researched with three-dimensional KIVA-3V program coupled with pressure wave equation. The cylinder pressure, combustion temperature, ignition delay timing, combustion duration, maximum mean pressure, and maximum oscillating pressure at different initial conditions are discussed and analyzed to investigate potential approaches to inhibiting engine knock while improving power output. The calculation results of the effect of just cooled EGR on knock characteristic show that appropriate levels of cooled EGR ratio can effectively suppress cylinder high-frequency pressure oscillations without obvious decrease in mean pressure. Analysis of the synergistic effect of cooled EGR, supercharging and compression ratio on knock characteristic indicates that under the condition of high supercharging and compression ratio, several times more cooled EGR ratio than that under the original condition is necessarily utilized to suppress knock occurrence effectively. The proposed method of synergistic effect of cooled EGR and downsizing technologies on knock characteristic, analyzed from the aspects of mean pressure and oscillating pressure, is an effective way to study downsized SI engine knock and provides knock inhibition approaches in practical engineering.

SAFE testing nuclear rockets economically

International Nuclear Information System (INIS)

Howe, Steven D.; Travis, Bryan; Zerkle, David K.

2003-01-01

Several studies over the past few decades have recognized the need for advanced propulsion to explore the solar system. As early as the 1960s, Werner Von Braun and others recognized the need for a nuclear rocket for sending humans to Mars. The great distances, the intense radiation levels, and the physiological response to zero-gravity all supported the concept of using a nuclear rocket to decrease mission time. These same needs have been recognized in later studies, especially in the Space Exploration Initiative in 1989. One of the key questions that has arisen in later studies, however, is the ability to test a nuclear rocket engine in the current societal environment. Unlike the Rover/NERVA programs in the 1960s, the rocket exhaust can no longer be vented to the open atmosphere. As a consequence, previous studies have examined the feasibility of building a large-scale version of the Nuclear Furnace Scrubber that was demonstrated in 1971. We have investigated an alternative that would deposit the rocket exhaust along with any entrained fission products directly into the ground. The Subsurface Active Filtering of Exhaust, or SAFE, concept would allow variable sized engines to be tested for long times at a modest expense. A system overview, results of preliminary calculations, and cost estimates of proof of concept demonstrations are presented. The results indicate that a nuclear rocket could be tested at the Nevada Test Site for under $20 M

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

An Object-Oriented Graphical User Interface for a Reusable Rocket Engine Intelligent Control System

Science.gov (United States)

Litt, Jonathan S.; Musgrave, Jeffrey L.; Guo, Ten-Huei; Paxson, Daniel E.; Wong, Edmond; Saus, Joseph R.; Merrill, Walter C.

1994-01-01

An intelligent control system for reusable rocket engines under development at NASA Lewis Research Center requires a graphical user interface to allow observation of the closed-loop system in operation. The simulation testbed consists of a real-time engine simulation computer, a controls computer, and several auxiliary computers for diagnostics and coordination. The system is set up so that the simulation computer could be replaced by the real engine and the change would be transparent to the control system. Because of the hard real-time requirement of the control computer, putting a graphical user interface on it was not an option. Thus, a separate computer used strictly for the graphical user interface was warranted. An object-oriented LISP-based graphical user interface has been developed on a Texas Instruments Explorer 2+ to indicate the condition of the engine to the observer through plots, animation, interactive graphics, and text.

Large Liquid Rocket Testing: Strategies and Challenges

Science.gov (United States)

Rahman, Shamim A.; Hebert, Bartt J.

2005-01-01

Rocket propulsion development is enabled by rigorous ground testing in order to mitigate the propulsion systems risks that are inherent in space flight. This is true for virtually all propulsive devices of a space vehicle including liquid and solid rocket propulsion, chemical and non-chemical propulsion, boost stage and in-space propulsion and so forth. In particular, large liquid rocket propulsion development and testing over the past five decades of human and robotic space flight has involved a combination of component-level testing and engine-level testing to first demonstrate that the propulsion devices were designed to meet the specified requirements for the Earth to Orbit launchers that they powered. This was followed by a vigorous test campaign to demonstrate the designed propulsion articles over the required operational envelope, and over robust margins, such that a sufficiently reliable propulsion system is delivered prior to first flight. It is possible that hundreds of tests, and on the order of a hundred thousand test seconds, are needed to achieve a high-reliability, flight-ready, liquid rocket engine system. This paper overviews aspects of earlier and recent experience of liquid rocket propulsion testing at NASA Stennis Space Center, where full scale flight engines and flight stages, as well as a significant amount of development testing has taken place in the past decade. The liquid rocket testing experience discussed includes testing of engine components (gas generators, preburners, thrust chambers, pumps, powerheads), as well as engine systems and complete stages. The number of tests, accumulated test seconds, and years of test stand occupancy needed to meet varying test objectives, will be selectively discussed and compared for the wide variety of ground test work that has been conducted at Stennis for subscale and full scale liquid rocket devices. Since rocket propulsion is a crucial long-lead element of any space system acquisition or

Nuclear Thermal Rocket Element Environmental Simulator (NTREES)

International Nuclear Information System (INIS)

Emrich, William J. Jr.

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

Aero-Thermo-Structural Analysis of Inlet for Rocket Based Combined Cycle Engines

Science.gov (United States)

Shivakumar, K. N.; Challa, Preeti; Sree, Dave; Reddy, Dhanireddy R. (Technical Monitor)

2000-01-01

NASA has been developing advanced space transportation concepts and technologies to make access to space less costly. One such concept is the reusable vehicles with short turn-around times. The NASA Glenn Research Center's concept vehicle is the Trailblazer powered by a rocket-based combined cycle (RBCC) engine. Inlet is one of the most important components of the RBCC engine. This paper presents fluid flow, thermal, and structural analysis of the inlet for Mach 6 free stream velocity for fully supersonic and supercritical with backpressure conditions. The results concluded that the fully supersonic condition was the most severe case and the largest stresses occur in the ceramic matrix composite layer of the inlet cowl. The maximum tensile and the compressive stresses were at least 3.8 and 3.4, respectively, times less than the associated material strength.

Preliminary Thermo-hydraulic Core Design Analysis of Korea Advanced Nuclear Thermal Engine Rocket for Space Application

Energy Technology Data Exchange (ETDEWEB)

Nam, Seung Hyun; Lee, Jeong Ik; Chang, Soon Heung [Korea Advanced Institute of Science and Technology, Daejeon (Korea, Republic of)

2014-05-15

Nclear rockets improve the propellant efficiency more than twice compared to CRs and thus significantly reduce the propellant requirement. The superior efficiency of nuclear rockets is due to the combination of the huge energy density and a single low molecular weight propellant utilization. Nuclear Thermal Rockets (NTRs) are particularly suitable for manned missions to Mars because it satisfies a relatively high thrust as well as a high propellant efficiency. NTRs use thermal energy released from a nuclear fission reactor to heat a single low molecular weight propellant, i. e., Hydrogen (H{sub 2}) and then exhausted the extremely heated propellant through a thermodynamic nozzle to produce thrust. A propellant efficiency parameter of rocket engines is specific impulse (I{sub sp}) which represents the ratio of the thrust over the rate of propellant consumption. The difference of I{sub sp} makes over three times propellant savings of NTRs for a manned Mars mission compared to CRs. NTRs can also be configured to operate bimodally by converting the surplus nuclear energy to auxiliary electric power required for the operation of a spacecraft. Moreover, the concept and technology of NTRs are very simple, already proven, and safe. Thus, NTRs can be applied to various space missions such as solar system exploration, International Space Station (ISS) transport support, Near Earth Objects (NEOs) interception, etc. Nuclear propulsion is the most promising and viable option to achieve challenging deep space missions. Particularly, the attractions of a NTR include excellent thrust and propellant efficiency, bimodal capability, proven technology, and safe and reliable performance. The ROK has also begun the research for space nuclear systems as a volunteer of the international space race and a major world nuclear energy country. KANUTER is one of the advanced NTR engines currently under development at KAIST. This bimodal engine is operated in two modes of propulsion with 100 MW

U.S./CIS eye joint nuclear rocket venture

Science.gov (United States)

Clark, John S.; Mcilwain, Melvin C.; Smetanikov, Vladimir; D'Yakov, Evgenij K.; Pavshuk, Vladimir A.

1993-01-01

An account is given of the significance for U.S. spacecraft development of a nuclear thermal rocket (NTR) reactor concept that has been developed in the (formerly Soviet) Commonwealth of Independent States (CIS). The CIS NTR reactor employs a hydrogen-cooled zirconium hydride moderator and ternary carbide fuels; the comparatively cool operating temperatures associated with this design promise overall robustness.

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.

Fundamental rocket injector/spray programs at the Phillips Laboratory

Science.gov (United States)

Talley, D. G.

1993-11-01

injectors, but can be a consideration in preburners, where the desire to keep turbine inlet temperatures as cool as possible can make it advantageous for the preburners to operate as far from stoichiometry as can be tolerated. ignitable during startup where pressures and temperatures are far from design conditions, and ignition transients must be minimized in order to avoid damage to engine components. In order to satisfy these various constraints, the injector designer must be able to perform design tradeoff studies, and it is important that this be done with minimal time and costs. In fact, it can easily be argued that reducing engine development time and costs is essential to maintaining U.S. competitiveness in space. *The Propulsion Directorate of the Phillips Laboratory has invested in a number of programs to advance liquid rocket engine technology, and several of these are directed at improving design tools for liquid rocket injectors. -The purpose of the presentation will be to describe some of these latter programs.

Subsonic Glideback Rocket Demonstrator Flight Testing

Science.gov (United States)

DeTurris, Dianne J.; Foster, Trevor J.; Barthel, Paul E.; Macy, Daniel J.; Droney, Christopher K.; Talay, Theodore A. (Technical Monitor)

2001-01-01

For the past two years, Cal Poly's rocket program has been aggressively exploring the concept of remotely controlled, fixed wing, flyable rocket boosters. This program, embodied by a group of student engineers known as Cal Poly Space Systems, has successfully demonstrated the idea of a rocket design that incorporates a vertical launch pattern followed by a horizontal return flight and landing. Though the design is meant for supersonic flight, CPSS demonstrators are deployed at a subsonic speed. Many steps have been taken by the club that allowed the evolution of the StarBooster prototype to reach its current size: a ten-foot tall, one-foot diameter, composite material rocket. Progress is currently being made that involves multiple boosters along with a second stage, third rocket.

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

Major accomplishments of America's nuclear rocket program (ROVER)

International Nuclear Information System (INIS)

Finseth, J.L.

1991-01-01

The United States embarked on a program to develop nuclear rocket engines in 1955. This program was known as project Rover. Initially nuclear rockets were considered as a potential backup for intercontinental ballistic missile propulsion but later proposed applications included both a lunar second stage as well as use in manned-Mars flights. Under the Rover program, 19 different reactors were built and tested during the period of 1959-1969. Additionally, several cold flow (non-fuelled) reactors were tested as well as a nuclear fuels test cell. The Rover program was terminated in 1973, due to budget constraints and an evolving political climate. The Rover program would have led to the development of a flight engine had the program continued through a logical continuation. The Rover program was responsible for a number of technological achievements. The successful operation of nuclear rocket engines on a system level represents the pinnacle of accomplishment. This paper will discuss the engine test program as well as several subsystems

Cooling system with compressor bleed and ambient air for gas turbine engine

Science.gov (United States)

Marsh, Jan H.; Marra, John J.

2017-11-21

A cooling system for a turbine engine for directing cooling fluids from a compressor to a turbine blade cooling fluid supply and from an ambient air source to the turbine blade cooling fluid supply to supply cooling fluids to one or more airfoils of a rotor assembly is disclosed. The cooling system may include a compressor bleed conduit extending from a compressor to the turbine blade cooling fluid supply that provides cooling fluid to at least one turbine blade. The compressor bleed conduit may include an upstream section and a downstream section whereby the upstream section exhausts compressed bleed air through an outlet into the downstream section through which ambient air passes. The outlet of the upstream section may be generally aligned with a flow of ambient air flowing in the downstream section. As such, the compressed air increases the flow of ambient air to the turbine blade cooling fluid supply.

Integrated System Health Management (ISHM) Implementation in Rocket Engine Testing

Science.gov (United States)

Figueroa, Fernando; Morris, Jon; Turowski, Mark; Franzl, Richard; Walker, Mark; Kapadia, Ravi; Venkatesh, Meera

2010-01-01

A pilot operational ISHM capability has been implemented for the E-2 Rocket Engine Test Stand (RETS) and a Chemical Steam Generator (CSG) test article at NASA Stennis Space Center. The implementation currently includes an ISHM computer and a large display in the control room. The paper will address the overall approach, tools, and requirements. It will also address the infrastructure and architecture. Specific anomaly detection algorithms will be discussed regarding leak detection and diagnostics, valve validation, and sensor validation. It will also describe development and use of a Health Assessment Database System (HADS) as a repository for measurements, health, configuration, and knowledge related to a system with ISHM capability. It will conclude with a discussion of user interfaces, and a description of the operation of the ISHM system prior, during, and after testing.

Approaches to Low Fuel Regression Rate in Hybrid Rocket Engines

Directory of Open Access Journals (Sweden)

Dario Pastrone

2012-01-01

Full Text Available Hybrid rocket engines are promising propulsion systems which present appealing features such as safety, low cost, and environmental friendliness. On the other hand, certain issues hamper the development hoped for. The present paper discusses approaches addressing improvements to one of the most important among these issues: low fuel regression rate. To highlight the consequence of such an issue and to better understand the concepts proposed, fundamentals are summarized. Two approaches are presented (multiport grain and high mixture ratio which aim at reducing negative effects without enhancing regression rate. Furthermore, fuel material changes and nonconventional geometries of grain and/or injector are presented as methods to increase fuel regression rate. Although most of these approaches are still at the laboratory or concept scale, many of them are promising.

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

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

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

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

Studies of Fission Fragment Rocket Engine Propelled Spacecraft

Science.gov (United States)

Werka, Robert O.; Clark, Rodney; Sheldon, Rob; Percy, Thomas K.

2014-01-01

The NASA Office of Chief Technologist has funded from FY11 through FY14 successive studies of the physics, design, and spacecraft integration of a Fission Fragment Rocket Engine (FFRE) that directly converts the momentum of fission fragments continuously into spacecraft momentum at a theoretical specific impulse above one million seconds. While others have promised future propulsion advances if only you have the patience, the FFRE requires no waiting, no advances in physics and no advances in manufacturing processes. Such an engine unequivocally can create a new era of space exploration that can change spacecraft operation. The NIAC (NASA Institute for Advanced Concepts) Program Phase 1 study of FY11 first investigated how the revolutionary FFRE technology could be integrated into an advanced spacecraft. The FFRE combines existent technologies of low density fissioning dust trapped electrostatically and high field strength superconducting magnets for beam management. By organizing the nuclear core material to permit sufficient mean free path for escape of the fission fragments and by collimating the beam, this study showed the FFRE could convert nuclear power to thrust directly and efficiently at a delivered specific impulse of 527,000 seconds. The FY13 study showed that, without increasing the reactor power, adding a neutral gas to the fission fragment beam significantly increased the FFRE thrust through in a manner analogous to a jet engine afterburner. This frictional interaction of gas and beam resulted in an engine that continuously produced 1000 pound force of thrust at a delivered impulse of 32,000 seconds, thereby reducing the currently studied DRM 5 round trip mission to Mars from 3 years to 260 days. By decreasing the gas addition, this same engine can be tailored for much lower thrust at much higher impulse to match missions to more distant destinations. These studies created host spacecraft concepts configured for manned round trip journeys. While the

Liquid rocket propulsion dynamic flow modeling using the ROCETS engineering modules in the EASY5x environment

Science.gov (United States)

Follett, Randolph F.; Taylor, Robert P.; Nunez, Stephen C.

1993-01-01

A report on the progress of porting the ROCETS (ROCket Engine Transient Simulator) into the EASY5x simulation environment is presented. Brief descriptions of each of the software systems, information regarding the actual port process, and examples comparing the results of the two systems are given. It is shown that EASY5x is a suitable environment for utilization of the ROCETS engineering modules, and that, for the example systems shown, EASY5x actually seems to give more accurate solutions than the straight ROCETS code.

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.

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.

Correlation of cylinder-head temperatures and coolant heat rejections of a multicylinder, liquid-cooled engine of 1710-cubic-inch displacement

Science.gov (United States)

Lundin, Bruce T; Povolny, John H; Chelko, Louis J

1949-01-01

Data obtained from an extensive investigation of the cooling characteristics of four multicylinder, liquid-cooled engines have been analyzed and a correlation of both the cylinder-head temperatures and the coolant heat rejections with the primary engine and coolant variables was obtained. The method of correlation was previously developed by the NACA from an analysis of the cooling processes involved in a liquid-cooled-engine cylinder and is based on the theory of nonboiling, forced-convection heat transfer. The data correlated included engine power outputs from 275 to 1860 brake horsepower; coolant flows from 50 to 320 gallons per minute; coolants varying in composition from 100 percent water to 97 percent ethylene glycol and 3 percent water; and ranges of engine speed, manifold pressure, carburetor-air temperature, fuel-air ratio, exhaust-gas pressure, ignition timing, and coolant temperature. The effect on engine cooling of scale formation on the coolant passages of the engine and of boiling of the coolant under various operating conditions is also discussed.

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.

Effect of cooling the recirculated exhaust gases on diesel engine emissions

International Nuclear Information System (INIS)

Abu-Hamdeh, Nidal H.

2003-01-01

Although combustion is essential in most energy generation processes, it is one of the major causes of air pollution. Spiral fin exhaust pipes were designed to study the effect of cooling the recirculated exhaust gases (EGR) of Diesel engines on the chemical composition of the exhaust gases and the reduction in the percentages of pollutant emissions. The gases examined in this study were oxides of nitrogen (NO x ), carbon dioxide (CO 2 ) and carbon monoxide (CO). In addition, O 2 concentration in the exhaust was measured. The two designs adopted in this study were exhaust pipes with solid and hollow fins around them. The first type uses air flow around the fins to cool the exhaust gases. The second type consists of hollow fins around the exhaust pipe to allow cooling water to flow in the hollow passage. Different combinations and arrangements of the solid and hollow fins exhaust pipes were used. It was found that decreasing the temperature of the EGR resulted in reductions in the oxides of nitrogen (NO x ) and carbon dioxide (CO 2 ) but increased the carbon monoxide (CO) in the exhaust gases. In addition, the oxygen (O 2 ) concentration in the exhaust was decreased. As a general trend, the percentages of reduction in the NO x gas concentrations were lower than the percentages of increase in the CO emissions as a result of cooling the EGR of a Diesel engine by a heat exchanger. Using water as a cooling medium decreased the exhaust gases temperature and the amount of pollutants more than did air as a cooling medium. In a separate series of tests, increasing the cooled EGR ratios decreased the exhaust NO x but increased the particulate matter concentrations in the exhaust gases

Effect of cooling the recirculated exhaust gases on diesel engine emissions

Energy Technology Data Exchange (ETDEWEB)

Abu-Hamdeh, Nidal H. [Jordan Univ. of Science and Technology, Irbid (Jordan)

2003-11-01

Although combustion is essential in most energy generation processes, it is one of the major causes of air pollution. Spiral fin exhaust pipes were designed to study the effect of cooling the recirculated exhaust gases (EGR) of Diesel engines on the chemical composition of the exhaust gases and the reduction in the percentages of pollutant emissions. The gases examined in this study were oxides of nitrogen (NO{sub x}), carbon dioxide (CO{sub 2}) and carbon monoxide (CO). In addition, O{sub 2} concentration in the exhaust was measured. The two designs adopted in this study were exhaust pipes with solid and hollow fins around them. The first type uses air flow around the fins to cool the exhaust gases. The second type consists of hollow fins around the exhaust pipe to allow cooling water to flow in the hollow passage. Different combinations and arrangements of the solid and hollow fins exhaust pipes were used. It was found that decreasing the temperature of the EGR resulted in reductions in the oxides of nitrogen (NO{sub x}) and carbon dioxide (CO{sub 2}) but increased the carbon monoxide (CO) in the exhaust gases. In addition, the oxygen (O{sub 2}) concentration in the exhaust was decreased. As a general trend, the percentages of reduction in the NO{sub x} gas concentrations were lower than the percentages of increase in the CO emissions as a result of cooling the EGR of a Diesel engine by a heat exchanger. Using water as a cooling medium decreased the exhaust gases temperature and the amount of pollutants more than did air as a cooling medium. In a separate series of tests, increasing the cooled EGR ratios decreased the exhaust NO{sub x} but increased the particulate matter concentrations in the exhaust gases. (Author)

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

Passive Rocket Diffuser Testing: Reacting Flow Performance of Four Second-Throat Geometries

Science.gov (United States)

Jones, Daniel R.; Allgood, Daniel C.; Saunders, Grady P.

2016-01-01

Second-throat diffusers serve to isolate rocket engines from the effects of ambient back pressure. As one of the nation's largest rocket testing facilities, the performance and design limitations of diffusers are of great interest to NASA's Stennis Space Center. This paper describes a series of tests conducted on four diffuser configurations to better understand the effects of inlet geometry and throat area on starting behavior and boundary layer separation. The diffusers were tested for a duration of five seconds with a 1455-pound thrust, LO2/GH2 thruster to ensure they each reached aerodynamic steady state. The effects of a water spray ring at the diffuser exits and a water-cooled deflector plate were also evaluated. Static pressure and temperature measurements were taken at multiple axial locations along the diffusers, and Computational Fluid Dynamics (CFD) simulations were used as a tool to aid in the interpretation of data. The hot combustion products were confirmed to enable the diffuser start condition with tighter second throats than predicted by historical cold-flow data or the theoretical normal shock method. Both aerodynamic performance and heat transfer were found to increase with smaller diffuser throats. Spray ring and deflector cooling water had negligible impacts on diffuser boundary layer separation. CFD was found to accurately capture diffuser shock structures and full-flowing diffuser wall pressures, and the qualitative behavior of heat transfer. However, the ability to predict boundary layer separated flows was not consistent.

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.

Kinetic---a system code for analyzing nuclear thermal propulsion rocket engine transients

International Nuclear Information System (INIS)

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

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

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.

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.

Acoustic streaming in simplified liquid rocket engines with transverse mode oscillations

Science.gov (United States)

Fischbach, Sean R.; Flandro, Gary A.; Majdalani, Joseph

2010-06-01

This study considers a simplified model of a liquid rocket engine in which uniform injection is imposed at the faceplate. The corresponding cylindrical chamber has a small length-to-diameter ratio with respect to solid and hybrid rockets. Given their low chamber aspect ratios, liquid thrust engines are known to experience severe tangential and radial oscillation modes more often than longitudinal ones. In order to model this behavior, tangential and radial waves are superimposed onto a basic mean-flow model that consists of a steady, uniform axial velocity throughout the chamber. Using perturbation tools, both potential and viscous flow equations are then linearized in the pressure wave amplitude and solved to the second order. The effects of the headwall Mach number are leveraged as well. While the potential flow analysis does not predict any acoustic streaming effects, the viscous solution carried out to the second order gives rise to steady secondary flow patterns near the headwall. These axisymmetric, steady contributions to the tangential and radial traveling waves are induced by the convective flow motion through interactions with inertial and viscous forces. We find that suppressing either the convective terms or viscosity at the headwall leads to spurious solutions that are free from streaming. In our problem, streaming is initiated at the headwall, within the boundary layer, and then extends throughout the chamber. We find that nonlinear streaming effects of tangential and radial waves act to alter the outer solution inside a cylinder with headwall injection. As a result of streaming, the radial wave velocities are intensified in one-half of the domain and reduced in the opposite half at any instant of time. Similarly, the tangential waves are either enhanced or weakened in two opposing sectors that are at 90° angle to the radial velocity counterparts. The second-order viscous solution that we obtain clearly displays both an oscillating and a steady flow

Hot-gas-side heat transfer characteristics of subscale, plug-nozzle rocket calorimeter chamber

Science.gov (United States)

Quentmeyer, Richard J.; Roncace, Elizabeth A.

1993-01-01

An experimental investigation was conducted to determine the hot-gas-side heat transfer characteristics for a liquid-hydrogen-cooled, subscale, plug-nozzle rocket test apparatus. This apparatus has been used since 1975 to evaluate rocket engine advanced cooling concepts and fabrication techniques, to screen candidate combustion chamber liner materials, and to provide data for model development. In order to obtain the data, a water-cooled calorimeter chamber having the same geometric configuration as the plug-nozzle test apparatus was tested. It also used the same two showerhead injector types that were used on the test apparatus: one having a Rigimesh faceplate and the other having a platelet faceplate. The tests were conducted using liquid oxygen and gaseous hydrogen as the propellants over a mixture ratio range of 5.8 to 6.3 at a nominal chamber pressure of 4.14 MPa abs (600 psia). The two injectors showed similar performance characteristics with the Rigimesh faceplate having a slightly higher average characteristic-exhaust-velocity efficiency of 96 percent versus 94.4 percent for the platelet faceplate. The throat heat flux was 54 MW/m(sup 2) (33 Btu/in.(sup 2)-sec) at the nominal operating condition, which was a chamber pressure of 4.14 MPa abs (600 psia), a hot-gas-side wall temperature of 730 K (1314 R), and a mixture ratio of 6.0. The chamber throat region correlation coefficient C(sub g) for a Nusselt number correlation of the form Nu =C(sub g)Re(sup 0.8)Pr(sup 0.3) averaged 0.023 for the Rigimesh faceplate and 0.026 for the platelet faceplate.

A review of findings of a study of rocket based combined cycle engines applied to extensively axisymmetric single stage to orbit vehicles

Science.gov (United States)

Foster, Richard W.

1992-01-01

Extensively axisymmetric and non-axisymmetric Single Stage To Orbit (SSTO) vehicles are considered. The information is presented in viewgraph form and the following topics are presented: payload comparisons; payload as a percent of dry weight - a system hardware cost indicator; life cycle cost estimations; operations and support costs estimation; selected engine type; and rocket engine specific impulse calculation.

A new facility for advanced rocket propulsion research

Science.gov (United States)

Zoeckler, Joseph G.; Green, James M.; Raitano, Paul

1993-06-01

A new test facility was constructed at the NASA Lewis Research Center Rocket Laboratory for the purpose of conducting rocket propulsion research at up to 8.9 kN (2000 lbf) thrust, using liquid oxygen and gaseous hydrogen propellants. A laser room adjacent to the test cell provides access to the rocket engine for advanced laser diagnostic systems. The size and location of the test cell provide the ability to conduct large amounts of testing in short time periods, with rapid turnover between programs. These capabilities make the new test facility an important asset for basic and applied rocket propulsion research.

Nuclear rockets: High-performance propulsion for Mars

International Nuclear Information System (INIS)

Watson, C.W.

1994-05-01

A new impetus to manned Mars exploration was introduced by President Bush in his Space Exploration Initiative. This has led, in turn, to a renewed interest in high-thrust nuclear thermal rocket propulsion (NTP). The purpose of this report is to give a brief tutorial introduction to NTP and provide a basic understanding of some of the technical issues in the realization of an operational NTP engine. Fundamental physical principles are outlined from which a variety of qualitative advantages of NTP over chemical propulsion systems derive, and quantitative performance comparisons are presented for illustrative Mars missions. Key technologies are described for a representative solid-core heat-exchanger class of engine, based on the extensive development work in the Rover and NERVA nuclear rocket programs (1955 to 1973). The most driving technology, fuel development, is discussed in some detail for these systems. Essential highlights are presented for the 19 full-scale reactor and engine tests performed in these programs. On the basis of these tests, the practicality of graphite-based nuclear rocket engines was established. Finally, several higher-performance advanced concepts are discussed. These have received considerable attention, but have not, as yet, developed enough credibility to receive large-scale development

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.

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.

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

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.

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

Energy Technology Data Exchange (ETDEWEB)

Nam, Seung Hyun; Chang, Soon Heung [Korea Advanced Institute of Science and Technology, Daejeon (Korea, Republic of)

2013-05-15

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.

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

Magnetocaloric Effect and Thermoelectric Cooling - A Synergistic Cooling Technology

Science.gov (United States)

2018-01-16

Thermoelectric Cooling - A Synergistic Cooling Technology Sb. GRANT NUMBER N00173-14-1-G016 Sc. PROGRAM ELEMENT NUMBER 82-2020-17 6. AUTHOR(S) 5d...Magnetocaloric Effect and Thermoelectric Cooling - A Synergistic Cooling Technology NRL Grant N00173-14-l-G016 CODE 8200: Spacecraft Engineering Department...82-11-0 1: Space and Space Systems Technology General Engineering & Research, L.L.C. Technical & Administrative point of contact: Dr. Robin

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.

Experimental study of gas engine driven air to water heat pump in cooling mode

International Nuclear Information System (INIS)

Elgendy, E.; Schmidt, J.

2010-01-01

Nowadays a sustainable development for more efficient use of energy and protection of the environment is of increasing importance. Gas engine heat pumps represent one of the most practicable solutions which offer high energy efficiency and environmentally friendly for heating and cooling applications. In this paper, the performance characteristics of gas engine driven heat pump used in water cooling were investigated experimentally without engine heat recovery. The effects of several important factors (evaporator water inlet temperature, evaporator water volume flow rate, ambient air temperature, and engine speed) on the performance of gas engine driven heat pump were studied in a wide range of operating conditions. The results showed that primary energy ratio of the system increased by 22.5% as evaporator water inlet temperature increased from 13 o C to 24 o C. On the other hand, varying of engine speed from 1300 rpm to 1750 rpm led to decrease in system primary energy ratio by 13%. Maximum primary energy ratio has been estimated with a value of two over a wide range of operating conditions.

Experimental characterization of mass, work and heat flows in an air cooled, single cylinder engine

International Nuclear Information System (INIS)

Perez-Blanco, H.

2004-01-01

Small air cooled engines, although large in numbers, receive scant attention in the literature. Experimental data for a four stroke, air cooled, single cylinder engine are presented in this report. Air to fuel ratios, indicated and output power, exhaust composition and heat loss are determined to result in suitable thermal and mechanical efficiencies. The data obtained are discussed with the perspective obtained from other literature references. Exhaust composition figures appear reasonable, but the measurement of the transient exhaust flows is still a concern. Based on the measurements, a graph illustrating the different energy transformations in the engine is produced. Undergraduate students in the curriculum routinely use the engine and the present work allows one to conclude that the measurement approach produces reasonable results. These results could be used by engine modelers and others interested in this wide field of technology

Calculated concentrations of any radionuclide deposited on the ground by release from underground nuclear detonations, tests of nuclear rockets, and tests of nuclear ramjet engines

International Nuclear Information System (INIS)

Hicks, H.G.

1981-11-01

This report presents calculated gamma radiation exposure rates and ground deposition of related radionuclides resulting from three types of event that deposited detectable radioactivity outside the Nevada Test Site complex, namely, underground nuclear detonations, tests of nuclear rocket engines and tests of nuclear ramjet engines

Analysis of film cooling in rocket nozzles

Science.gov (United States)

Woodbury, Keith A.

1993-01-01

This report summarizes the findings on the NASA contract NAG8-212, Task No. 3. The overall project consists of three tasks, all of which have been successfully completed. In addition, some supporting supplemental work, not required by the contract, has been performed and is documented herein. Task 1 involved the modification of the wall functions in the code FDNS (Finite Difference Navier-Stokes) to use a Reynolds Analogy-based method. This task was completed in August, 1992. Task 2 involved the verification of the code against experimentally available data. The data chosen for comparison was from an experiment involving the injection of helium from a wall jet. Results obtained in completing this task also show the sensitivity of the FDNS code to unknown conditions at the injection slot. This task was completed in September, 1992. Task 3 required the computation of the flow of hot exhaust gases through the P&W 40K subscale nozzle. Computations were performed both with and without film coolant injection. This task was completed in July, 1993. The FDNS program tends to overpredict heat fluxes, but, with suitable modeling of backside cooling, may give reasonable wall temperature predictions. For film cooling in the P&W 40K calorimeter subscale nozzle, the average wall temperature is reduced from 1750R to about 1050R by the film cooling. The average wall heat flux is reduced by a factor of 3.

AUTOMOTIVE DIESEL MAINTENACE 1. UNIT XV, I--MAINTAINING THE COOLING SYSTEM, CUMMINS DIESEL ENGINE, I--UNIT INSTALLATION--TRANSMISSION.

Science.gov (United States)

Human Engineering Inst., Cleveland, OH.

THIS MODULE OF A 30-MODULE COURSE IS DESIGNED TO DEVELOP AN UNDERSTANDING OF THE FUNCTION AND MAINTENANCE OF THE DIESEL ENGINE COOLING SYSTEM AND THE PROCEDURES FOR TRANSMISSION INSTALLATION. TOPICS ARE (1) IMPORTANCE OF THE COOLING SYSTEM, (2) COOLING SYSTEM COMPONENTS, (3) EVALUATING COOLING SYSTEM FAILURES, (4) CARING FOR THE COOLING SYSTEM,…

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.

Ongoing Analysis of Rocket Based Combined Cycle Engines by the Applied Fluid Dynamics Analysis Group at Marshall Space Flight Center

Science.gov (United States)

Ruf, Joseph; Holt, James B.; Canabal, Francisco

1999-01-01

This paper presents the status of analyses on three Rocket Based Combined Cycle configurations underway in the Applied Fluid Dynamics Analysis Group (TD64). TD64 is performing computational fluid dynamics analysis on a Penn State RBCC test rig, the proposed Draco axisymmetric RBCC engine and the Trailblazer engine. The intent of the analysis on the Penn State test rig is to benchmark the Finite Difference Navier Stokes code for ejector mode fluid dynamics. The Draco engine analysis is a trade study to determine the ejector mode performance as a function of three engine design variables. The Trailblazer analysis is to evaluate the nozzle performance in scramjet mode. Results to date of each analysis are presented.

Introduction to the Special Issue on Sounding Rockets and Instrumentation

OpenAIRE

Christe, Steven; Zeiger, Ben; Pfaff, Rob; Garcia, Michael

2016-01-01

Rocket technology, originally developed for military applications, has provided a low-cost observing platform to carry critical and rapid-response scientific investigations for over 70 years. Even with the development of launch vehicles that could put satellites into orbit, high altitude sounding rockets have remained relevant. In addition to science observations, sounding rockets provide a unique technology test platform and a valuable training ground for scientists and engineers. Most impor...

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

Computational study of variable area ejector rocket flowfields

Science.gov (United States)

Etele, Jason

Access to space has always been a scientific priority for countries which can afford the prohibitive costs associated with launch. However, the large scale exploitation of space by the business community will require the cost of placing payloads into orbit be dramatically reduced for space to become a truly profitable commodity. To this end, this work focuses on a next generation propulsive technology called the Rocket Based Combined Cycle (RBCC) engine in which rocket, ejector, ramjet, and scramjet cycles operate within the same engine environment. Using an in house numerical code solving the axisymmetric version of the Favre averaged Navier Stokes equations (including the Wilcox ko turbulence model with dilatational dissipation) a systematic study of various ejector designs within an RBCC engine is undertaken. It is shown that by using a central rocket placed along the axisymmetric axis in combination with an annular rocket placed along the outer wall of the ejector, one can obtain compression ratios of approximately 2.5 for the case where both the entrained air and rocket exhaust mass flows are equal. Further, it is shown that constricting the exit area, and the manner in which this constriction is performed, has a significant positive impact on the compression ratio. For a decrease in area of 25% a purely conical ejector can increase the compression ratio by an additional 23% compared to an equal length unconstricted ejector. The use of a more sharply angled conical section followed by a cylindrical section to maintain equivalent ejector lengths can further increase the compression ratio by 5--7% for a total increase of approximately 30%.

The main indicators of the health of children and adolescents in residential zone of the facility for disposal of rocket engines

Directory of Open Access Journals (Sweden)

Tarakanova S.Y.

2014-12-01

39.5%. The main cause of morbidity in children is diseases of the nervous system and mental disorders, and congenital anomalies. Conclusion. Operation of installations for the disposal of rocket engines solid fuel according to the official reporting forms medical institutions has no effect on child health.

Mid-section of a can-annular gas turbine engine with a cooling system for the transition

Science.gov (United States)

Wiebe, David J.; Rodriguez, Jose L.

2015-12-08

A cooling system is provided for a transition (420) of a gas turbine engine (410). The cooling system includes a cowling (460) configured to receive an air flow (111) from an outlet of a compressor section of the gas turbine engine (410). The cowling (460) is positioned adjacent to a region of the transition (420) to cool the transition region upon circulation of the air flow within the cowling (460). The cooling system further includes a manifold (121) to directly couple the air flow (111) from the compressor section outlet to an inlet (462) of the cowling (460). The cowling (460) is configured to circulate the air flow (111) within an interior space (426) of the cowling (460) that extends radially outward from an inner diameter (423) of the cowling to an outer diameter (424) of the cowling at an outer surface.

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.

A Programmatic and Engineering Approach to the Development of a Nuclear Thermal Rocket for Space Exploration

Science.gov (United States)

Bordelon, Wayne J., Jr.; Ballard, Rick O.; Gerrish, Harold P., Jr.

2006-01-01

With the announcement of the Vision for Space Exploration on January 14, 2004, there has been a renewed interest in nuclear thermal propulsion. Nuclear thermal propulsion is a leading candidate for in-space propulsion for human Mars missions; however, the cost to develop a nuclear thermal rocket engine system is uncertain. Key to determining the engine development cost will be the engine requirements, the technology used in the development and the development approach. The engine requirements and technology selection have not been defined and are awaiting definition of the Mars architecture and vehicle definitions. The paper discusses an engine development approach in light of top-level strategic questions and considerations for nuclear thermal propulsion and provides a suggested approach based on work conducted at the NASA Marshall Space Flight Center to support planning and requirements for the Prometheus Power and Propulsion Office. This work is intended to help support the development of a comprehensive strategy for nuclear thermal propulsion, to help reduce the uncertainty in the development cost estimate, and to help assess the potential value of and need for nuclear thermal propulsion for a human Mars mission.

Research and development of cooled turbine for aircraft engines. Koku engine yo reikyaku turbine no kenkyu kaihatsu

Energy Technology Data Exchange (ETDEWEB)

Maya, T; Yamawaki, S [Ishikawajima-Harima Heavy Industries, Co. Ltd., Tokyo (Japan)

1994-05-01

For the turbine which is one of the principal elements of aircraft engine, progress in turbine use material development and cooling performance further heightened for the turbine are needed to grapple with the required heightening of turbine inlet temperature. In the present paper based on the turbine inlet temperature designed to be 1600[degree]C as a target, a two-dimensional model used for the turbine cooling performance test was structurally given together with the result of the above test which aimed at confirming the design calculation. As a result of cooling design for the turbine which was about 1600[degree]C in inlet temperature, the highest gas temperature was 1890 and 1470[degree]C on the stator blade and rotor blade, respectively. Both those blades were 0.66 and 0.62, respectively in cooling efficiency. To test the cooling performance, a two-dimensional cascade was tested with a doubly amplified model of cooling blade, the use of which could set its Reynolds number near that of the actual one. As compared with the actual operation, the test was made at low temperatures of 400 to 500[degree]C and low pressures of 0.02 to 0.03MPa. The test agreed with the design calculation in result. 4 refs., 8 figs.

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...

Turbulent Mixing of Primary and Secondary Flow Streams in a Rocket-Based Combined Cycle Engine

Science.gov (United States)

Cramer, J. M.; Greene, M. U.; Pal, S.; Santoro, R. J.; Turner, Jim (Technical Monitor)

2002-01-01

This viewgraph presentation gives an overview of the turbulent mixing of primary and secondary flow streams in a rocket-based combined cycle (RBCC) engine. A significant RBCC ejector mode database has been generated, detailing single and twin thruster configurations and global and local measurements. On-going analysis and correlation efforts include Marshall Space Flight Center computational fluid dynamics modeling and turbulent shear layer analysis. Potential follow-on activities include detailed measurements of air flow static pressure and velocity profiles, investigations into other thruster spacing configurations, performing a fundamental shear layer mixing study, and demonstrating single-shot Raman measurements.

Combustion response to acoustic perturbation in liquid rocket engines

Science.gov (United States)

Ghafourian, Akbar

An experimental study of the effect of acoustic perturbations on combustion behavior of a model liquid propellant rocket engine has been carried out. A pair of compression drivers were used to excite transverse and longitudinal acoustic fields at strengths of up to 156.6 dB and 159.5 dB respectively in the combustion chamber of the experimental rocket engine. Propellant simulants were injected into the combustion chamber through a single element shear coaxial injector. Water and air were used in cold flow studies and ethanol and oxygen-enriched air were used as fuel and oxidizer in reacting hot flow studies. In cold flow studies an imposed transverse acoustic field had a more pronounced effect on the spray pattern than a longitudinal acoustic fields. A transverse acoustic field widened the spray by as much as 33 percent and the plane of impingement of the spray with chamber walls moved up closer to the injection plane. The behavior was strongly influenced by the gas phase velocity but was less sensitive to changes in the liquid phase velocity. In reacting hot flow studies the effects of changes in equivalence ratio, excitation amplitude, excitation frequency, liquid and gas phase velocity and chamber pressure on the response of the injector to imposed high frequency transverse acoustic excitation were measured. Reducing the equivalence ratio from 7.4 to 3.8 increased the chamber pressure response to the imposed excitation at 3000 Hz. Increasing the excitation amplitude from 147 dB to 155.6 dB at 3000 Hz increased the chamber pressure response to the excitation. In the frequency range of 1240 Hz to 3220 Hz, an excitation frequency of 3000 Hz resulted in the largest response of the chamber pressure indicating the importance of fluid dynamic coupling. Increasing the liquid phase velocity from 9.2 m/sec to 22.7 m/sec, did not change the amplitude of the chamber pressure response to excitation. This implied the importance of local equivalence ratio and not the overall

Performance Evaluation of a Software Engineering Tool for Automated Design of Cooling Systems in Injection Moulding

DEFF Research Database (Denmark)

Jauregui-Becker, Juan M.; Tosello, Guido; van Houten, Fred J.A.M.

2013-01-01

This paper presents a software tool for automating the design of cooling systems for injection moulding and a validation of its performance. Cooling system designs were automatically generated by the proposed software tool and by applying a best practice tool engineering design approach. The two...

Heat removal performance of auxiliary cooling system for the high temperature engineering test reactor during scrams

International Nuclear Information System (INIS)

Takeda, Takeshi; Tachibana, Yukio; Iyoku, Tatsuo; Takenaka, Satsuki

2003-01-01

The auxiliary cooling system of the high temperature engineering test reactor (HTTR) is employed for heat removal as an engineered safety feature when the reactor scrams in an accident when forced circulation can cool the core. The HTTR is the first high temperature gas-cooled reactor in Japan with reactor outlet gas temperature of 950 degree sign C and thermal power of 30 MW. The auxiliary cooling system should cool the core continuously avoiding excessive cold shock to core graphite components and water boiling of itself. Simulation tests on manual trip from 9 MW operation and on loss of off-site electric power from 15 MW operation were carried out in the rise-to-power test up to 20 MW of the HTTR. Heat removal characteristics of the auxiliary cooling system were examined by the tests. Empirical correlations of overall heat transfer coefficients were acquired for a helium/water heat exchanger and air cooler for the auxiliary cooling system. Temperatures of fluids in the auxiliary cooling system were predicted on a scram event from 30 MW operation at 950 degree sign C of the reactor outlet coolant temperature. Under the predicted helium condition of the auxiliary cooling system, integrity of fuel blocks among the core graphite components was investigated by stress analysis. Evaluation results showed that overcooling to the core graphite components and boiling of water in the auxiliary cooling system should be prevented where open area condition of louvers in the air cooler is the full open

Turbopump Design and Analysis Approach for Nuclear Thermal Rockets

International Nuclear Information System (INIS)

Chen, Shucheng S.; Veres, Joseph P.; Fittje, James E.

2006-01-01

A rocket propulsion system, whether it is a chemical rocket or a nuclear thermal rocket, is fairly complex in detail but rather simple in principle. Among all the interacting parts, three components stand out: they are pumps and turbines (turbopumps), and the thrust chamber. To obtain an understanding of the overall rocket propulsion system characteristics, one starts from analyzing the interactions among these three components. It is therefore of utmost importance to be able to satisfactorily characterize the turbopump, level by level, at all phases of a vehicle design cycle. Here at the NASA Glenn Research Center, as the starting phase of a rocket engine design, specifically a Nuclear Thermal Rocket Engine design, we adopted the approach of using a high level system cycle analysis code (NESS) to obtain an initial analysis of the operational characteristics of a turbopump required in the propulsion system. A set of turbopump design codes (PumpDes and TurbDes) were then executed to obtain sizing and performance parameters of the turbopump that were consistent with the mission requirements. A set of turbopump analyses codes (PUMPA and TURBA) were applied to obtain the full performance map for each of the turbopump components; a two dimensional layout of the turbopump based on these mean line analyses was also generated. Adequacy of the turbopump conceptual design will later be determined by further analyses and evaluation. In this paper, descriptions and discussions of the aforementioned approach are provided and future outlooks are discussed

The Alfred Nobel rocket camera. An early aerial photography attempt

Science.gov (United States)

Ingemar Skoog, A.

2010-02-01

Alfred Nobel (1833-1896), mainly known for his invention of dynamite and the creation of the Nobel Prices, was an engineer and inventor active in many fields of science and engineering, e.g. chemistry, medicine, mechanics, metallurgy, optics, armoury and rocketry. Amongst his inventions in rocketry was the smokeless solid propellant ballistite (i.e. cordite) patented for the first time in 1887. As a very wealthy person he actively supported many Swedish inventors in their work. One of them was W.T. Unge, who was devoted to the development of rockets and their applications. Nobel and Unge had several rocket patents together and also jointly worked on various rocket applications. In mid-1896 Nobel applied for patents in England and France for "An Improved Mode of Obtaining Photographic Maps and Earth or Ground Measurements" using a photographic camera carried by a "…balloon, rocket or missile…". During the remaining of 1896 the mechanical design of the camera mechanism was pursued and cameras manufactured. In April 1897 (after the death of Alfred Nobel) the first aerial photos were taken by these cameras. These photos might be the first documented aerial photos taken by a rocket borne camera. Cameras and photos from 1897 have been preserved. Nobel did not only develop the rocket borne camera but also proposed methods on how to use the photographs taken for ground measurements and preparing maps.

LOX/hydrocarbon rocket engine analytical design methodology development and validation. Volume 2: Appendices

Science.gov (United States)

Niiya, Karen E.; Walker, Richard E.; Pieper, Jerry L.; Nguyen, Thong V.

1993-05-01

This final report includes a discussion of the work accomplished during the period from Dec. 1988 through Nov. 1991. The objective of the program was to assemble existing performance and combustion stability models into a usable design methodology capable of designing and analyzing high-performance and stable LOX/hydrocarbon booster engines. The methodology was then used to design a validation engine. The capabilities and validity of the methodology were demonstrated using this engine in an extensive hot fire test program. The engine used LOX/RP-1 propellants and was tested over a range of mixture ratios, chamber pressures, and acoustic damping device configurations. This volume contains time domain and frequency domain stability plots which indicate the pressure perturbation amplitudes and frequencies from approximately 30 tests of a 50K thrust rocket engine using LOX/RP-1 propellants over a range of chamber pressures from 240 to 1750 psia with mixture ratios of from 1.2 to 7.5. The data is from test configurations which used both bitune and monotune acoustic cavities and from tests with no acoustic cavities. The engine had a length of 14 inches and a contraction ratio of 2.0 using a 7.68 inch diameter injector. The data was taken from both stable and unstable tests. All combustion instabilities were spontaneous in the first tangential mode. Although stability bombs were used and generated overpressures of approximately 20 percent, no tests were driven unstable by the bombs. The stability instrumentation included six high-frequency Kistler transducers in the combustion chamber, a high-frequency Kistler transducer in each propellant manifold, and tri-axial accelerometers. Performance data is presented, both characteristic velocity efficiencies and energy release efficiencies, for those tests of sufficient duration to record steady state values.

Ongoing Analyses of Rocket Based Combined Cycle Engines by the Applied Fluid Dynamics Analysis Group at Marshall Space Flight Center

Science.gov (United States)

Ruf, Joseph H.; Holt, James B.; Canabal, Francisco

2001-01-01

This paper presents the status of analyses on three Rocket Based Combined Cycle (RBCC) configurations underway in the Applied Fluid Dynamics Analysis Group (TD64). TD64 is performing computational fluid dynamics (CFD) analysis on a Penn State RBCC test rig, the proposed Draco axisymmetric RBCC engine and the Trailblazer engine. The intent of the analysis on the Penn State test rig is to benchmark the Finite Difference Navier Stokes (FDNS) code for ejector mode fluid dynamics. The Draco analysis was a trade study to determine the ejector mode performance as a function of three engine design variables. The Trailblazer analysis is to evaluate the nozzle performance in scramjet mode. Results to date of each analysis are presented.

High-Temperature Polymer Composites Tested for Hypersonic Rocket Combustor Backup Structure

Science.gov (United States)

Sutter, James K.; Shin, E. Eugene; Thesken, John C.; Fink, Jeffrey E.

2005-01-01

Significant component weight reductions are required to achieve the aggressive thrust-toweight goals for the Rocket Based Combined Cycle (RBCC) third-generation, reusable liquid propellant rocket engine, which is one possible engine for a future single-stage-toorbit vehicle. A collaboration between the NASA Glenn Research Center and Boeing Rocketdyne was formed under the Higher Operating Temperature Propulsion Components (HOTPC) program and, currently, the Ultra-Efficient Engine Technology (UEET) Project to develop carbon-fiber-reinforced high-temperature polymer matrix composites (HTPMCs). This program focused primarily on the combustor backup structure to replace all metallic support components with a much lighter polymer-matrixcomposite- (PMC-) titanium honeycomb sandwich structure.

Rhenium Rocket Manufacturing Technology

Science.gov (United States)

1997-01-01

The NASA Lewis Research Center's On-Board Propulsion Branch has a research and technology program to develop high-temperature (2200 C), iridium-coated rhenium rocket chamber materials for radiation-cooled rockets in satellite propulsion systems. Although successful material demonstrations have gained much industry interest, acceptance of the technology has been hindered by a lack of demonstrated joining technologies and a sparse materials property data base. To alleviate these concerns, we fabricated rhenium to C-103 alloy joints by three methods: explosive bonding, diffusion bonding, and brazing. The joints were tested by simulating their incorporation into a structure by welding and by simulating high-temperature operation. Test results show that the shear strength of the joints degrades with welding and elevated temperature operation but that it is adequate for the application. Rhenium is known to form brittle intermetallics with a number of elements, and this phenomena is suspected to cause the strength degradation. Further bonding tests with a tantalum diffusion barrier between the rhenium and C-103 is planned to prevent the formation of brittle intermetallics.

Current and Future Critical Issues in Rocket Propulsion Systems

Science.gov (United States)

Navaz, Homayun K.; Dix, Jeff C.

1998-01-01

The objective of this research was to tackle several problems that are currently of great importance to NASA. In a liquid rocket engine several complex processes take place that are not thoroughly understood. Droplet evaporation, turbulence, finite rate chemistry, instability, and injection/atomization phenomena are some of the critical issues being encountered in a liquid rocket engine environment. Pulse Detonation Engines (PDE) performance, combustion chamber instability analysis, 60K motor flowfield pattern from hydrocarbon fuel combustion, and 3D flowfield analysis for the Combined Cycle engine were of special interest to NASA. During the summer of 1997, we made an attempt to generate computational results for all of the above problems and shed some light on understanding some of the complex physical phenomena. For this purpose, the Liquid Thrust Chamber Performance (LTCP) code, mainly designed for liquid rocket engine applications, was utilized. The following test cases were considered: (1) Characterization of a detonation wave in a Pulse Detonation Tube; (2) 60K Motor wall temperature studies; (3) Propagation of a pressure pulse in a combustion chamber (under single and two-phase flow conditions); (4) Transonic region flowfield analysis affected by viscous effects; (5) Exploring the viscous differences between a smooth and a corrugated wall; and (6) 3D thrust chamber flowfield analysis of the Combined Cycle engine. It was shown that the LTCP-2D and LTCP-3D codes are capable of solving complex and stiff conservation equations for gaseous and droplet phases in a very robust and efficient manner. These codes can be run on a workstation and personal computers (PC's).

Robust Low Cost Liquid Rocket Combustion Chamber by Advanced Vacuum Plasma Process

Science.gov (United States)

Holmes, Richard; Elam, Sandra; Ellis, David L.; McKechnie, Timothy; Hickman, Robert; Rose, M. Franklin (Technical Monitor)

2001-01-01

Next-generation, regeneratively cooled rocket engines will require materials that can withstand high temperatures while retaining high thermal conductivity. Fabrication techniques must be cost efficient so that engine components can be manufactured within the constraints of shrinking budgets. Three technologies have been combined to produce an advanced liquid rocket engine combustion chamber at NASA-Marshall Space Flight Center (MSFC) using relatively low-cost, vacuum-plasma-spray (VPS) techniques. Copper alloy NARloy-Z was replaced with a new high performance Cu-8Cr-4Nb alloy developed by NASA-Glenn Research Center (GRC), which possesses excellent high-temperature strength, creep resistance, and low cycle fatigue behavior combined with exceptional thermal stability. Functional gradient technology, developed building composite cartridges for space furnaces was incorporated to add oxidation resistant and thermal barrier coatings as an integral part of the hot wall of the liner during the VPS process. NiCrAlY, utilized to produce durable protective coating for the space shuttle high pressure fuel turbopump (BPFTP) turbine blades, was used as the functional gradient material coating (FGM). The FGM not only serves as a protection from oxidation or blanching, the main cause of engine failure, but also serves as a thermal barrier because of its lower thermal conductivity, reducing the temperature of the combustion liner 200 F, from 1000 F to 800 F producing longer life. The objective of this program was to develop and demonstrate the technology to fabricate high-performance, robust, inexpensive combustion chambers for advanced propulsion systems (such as Lockheed-Martin's VentureStar and NASA's Reusable Launch Vehicle, RLV) using the low-cost VPS process. VPS formed combustion chamber test articles have been formed with the FGM hot wall built in and hot fire tested, demonstrating for the first time a coating that will remain intact through the hot firing test, and with

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.

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

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

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.

Numerical Model of a Variable-Combined-Cycle Engine for Dual Subsonic and Supersonic Cruise

Directory of Open Access Journals (Sweden)

Victor Fernandez-Villace

2013-02-01

Full Text Available Efficient high speed propulsion requires exploiting the cooling capability of the cryogenic fuel in the propulsion cycle. This paper presents the numerical model of a combined cycle engine while in air turbo-rocket configuration. Specific models of the various heat exchanger modules and the turbomachinery elements were developed to represent the physical behavior at off-design operation. The dynamic nature of the model allows the introduction of the engine control logic that limits the operation of certain subcomponents and extends the overall engine operational envelope. The specific impulse and uninstalled thrust are detailed while flying a determined trajectory between Mach 2.5 and 5 for varying throttling levels throughout the operational envelope.

Data on test results of vessel cooling system of high temperature engineering test reactor

International Nuclear Information System (INIS)

Saikusa, Akio; Nakagawa, Shigeaki; Fujimoto, Nozomu; Tachibana, Yukio; Iyoku, Tatsuo

2003-02-01

High Temperature Engineering Test Reactor (HTTR) is the first graphite-moderated helium gas cooled reactor in Japan. The rise-to-power test of the HTTR started on September 28, 1999 and thermal power of the HTTR reached its full power of 30 MW on December 7, 2001. Vessel Cooling System (VCS) of the HTTR is the first Reactor Cavity Cooling System (RCCS) applied for High Temperature Gas Cooled Reactors. The VCS cools the core indirectly through the reactor pressure vessel to keep core integrity during the loss of core flow accidents such as depressurization accident. Minimum heat removal of the VCS to satisfy its safety requirement is 0.3MW at 30 MW power operation. Through the performance test of the VCS in the rise-to-power test of the HTTR, it was confirmed that the VCS heat removal at 30 MW power operation was higher than 0.3 MW. This paper shows outline of the VCS and test results on the VCS performance. (author)

Engineered design of SSC cooling ponds

International Nuclear Information System (INIS)

Bear, J.B.

1993-05-01

The cooling requirements of the SSC are significant and adequate cooling water systems to meet these requirements are critical to the project's successful operation. The use of adequately designed cooling ponds will provide reliable cooling for operation while also meeting environmental goals of the project to maintain streamflow and flood peaks to preconstruction levels as well as other streamflow and water quality requirements of the Texas Water Commission and the Environmental Protection Agency

Parametric Data from a Wind Tunnel Test on a Rocket-Based Combined-Cycle Engine Inlet

Science.gov (United States)

Fernandez, Rene; Trefny, Charles J.; Thomas, Scott R.; Bulman, Mel J.

2001-01-01

A 40-percent scale model of the inlet to a rocket-based combined-cycle (RBCC) engine was tested in the NASA Glenn Research Center 1- by 1-Foot Supersonic Wind Tunnel (SWT). The full-scale RBCC engine is scheduled for test in the Hypersonic Tunnel Facility (HTF) at NASA Glenn's Plum Brook Station at Mach 5 and 6. This engine will incorporate the configuration of this inlet model which achieved the best performance during the present experiment. The inlet test was conducted at Mach numbers of 4.0, 5.0, 5.5, and 6.0. The fixed-geometry inlet consists of an 8 deg.. forebody compression plate, boundary layer diverter, and two compressive struts located within 2 parallel sidewalls. These struts extend through the inlet, dividing the flowpath into three channels. Test parameters investigated included strut geometry, boundary layer ingestion, and Reynolds number (Re). Inlet axial pressure distributions and cross-sectional Pitot-pressure surveys at the base of the struts were measured at varying back-pressures. Inlet performance and starting data are presented. The inlet chosen for the RBCC engine self-started at all Mach numbers from 4 to 6. Pitot-pressure contours showed large flow nonuniformity on the body-side of the inlet. The inlet provided adequate pressure recovery and flow quality for the RBCC cycle even with the flow separation.

Taming Liquid Hydrogen: The Centaur Upper Stage Rocket

Science.gov (United States)

Dawson, Virginia P.; Bowles, Mark D.

2004-01-01

The Centaur is one of the most powerful rockets in the world. As an upper-stage rocket for the Atlas and Titan boosters it has been a reliable workhorse for NASA for over forty years and has played an essential role in many of NASA's adventures into space. In this CD-ROM you will be able to explore the Centaur's history in various rooms to this virtual museum. Visit the "Movie Theater" to enjoy several video documentaries on the Centaur. Enter the "Interview Booth" to hear and read interviews with scientists and engineers closely responsible for building and operating the rocket. Go to the "Photo Gallery" to look at numerous photos of the rocket throughout its history. Wander into the "Centaur Library" to read various primary documents of the Centaur program. Finally, stop by the "Observation Deck" to watch a virtual Centaur in flight.

Cooling towers

International Nuclear Information System (INIS)

Boernke, F.

1975-01-01

The need for the use of cooling systems in power plant engineering is dealt with from the point of view of a non-polluting form of energy production. The various cooling system concepts up to the modern natural-draught cooling towers are illustrated by examples. (TK/AK) [de

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.

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.

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...

Turbine airfoil cooling system with cooling systems using high and low pressure cooling fluids

Science.gov (United States)

Marsh, Jan H.; Messmann, Stephen John; Scribner, Carmen Andrew

2017-10-25

A turbine airfoil cooling system including a low pressure cooling system and a high pressure cooling system for a turbine airfoil of a gas turbine engine is disclosed. In at least one embodiment, the low pressure cooling system may be an ambient air cooling system, and the high pressure cooling system may be a compressor bleed air cooling system. In at least one embodiment, the compressor bleed air cooling system in communication with a high pressure subsystem that may be a snubber cooling system positioned within a snubber. A delivery system including a movable air supply tube may be used to separate the low and high pressure cooling subsystems. The delivery system may enable high pressure cooling air to be passed to the snubber cooling system separate from low pressure cooling fluid supplied by the low pressure cooling system to other portions of the turbine airfoil cooling system.

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.

Adaptive individual-cylinder thermal state control using piston cooling for a GDCI engine

Science.gov (United States)

Roth, Gregory T; Husted, Harry L; Sellnau, Mark C

2015-04-07

A system for a multi-cylinder compression ignition engine includes a plurality of nozzles, at least one nozzle per cylinder, with each nozzle configured to spray oil onto the bottom side of a piston of the engine to cool that piston. Independent control of the oil spray from the nozzles is provided on a cylinder-by-cylinder basis. A combustion parameter is determined for combustion in each cylinder of the engine, and control of the oil spray onto the piston in that cylinder is based on the value of the combustion parameter for combustion in that cylinder. A method for influencing combustion in a multi-cylinder engine, including determining a combustion parameter for combustion taking place in in a cylinder of the engine and controlling an oil spray targeted onto the bottom of a piston disposed in that cylinder is also presented.

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.

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.

Engineering Encounters: Blasting off with Engineering

Science.gov (United States)

Dare, Emily A.; Childs, Gregory T.; Cannaday, E. Ashley; Roehrig, Gillian H

2014-01-01

What better way to engage young students in physical science concepts than to have them engineer flying toy rockets? The integration of engineering into science classrooms is advocated by the "Next Generation Science Standards" (NGSS) and researchers alike (Brophy et al. 2008), as engineering provides: (1) A "real-world…

Impingement jet cooling in gas turbines

CERN Document Server

Amano, R S

2014-01-01

Due to the requirement for enhanced cooling technologies on modern gas turbine engines, advanced research and development has had to take place in field of thermal engineering. Impingement jet cooling is one of the most effective in terms of cooling, manufacturability and cost. This is the first to book to focus on impingement cooling alone.

Development and analysis of a variable position thermostat for smart cooling system of a light duty diesel vehicles and engine emissions assessment during NEDC

International Nuclear Information System (INIS)

Mohamed, Eid S.

2016-01-01

Highlights: • A new concept of the variable position electromagnetic thermostat in MCS is proposed. • A series of experiments were conducted on a light duty diesel vehicle operated over the NEDC test. • A comparative study was done on emission characteristics of the MCS and the conventional cooling system. • Engine cold start and steady-state coolant flow rate and emissions are presented. • The effect of MCS on engine accumulation FC and emissions over NEDC are evaluated. - Graphical Abstract: Display Omitted - Abstract: Smart cooling control systems for IC engines can better regulate the combustion process and heat, a variable position thermostat and electric coolant pumps (EWP) for IC engines are under development by a number of researchers. However, the aim of this study is to assess the performance of a variable position electromagnetic thermostat (VPEMT) to provide more flexible control of the engine temperature and coolant mass flow rate of modification cooling system (MCS). The measurement procedure was applied to two phases under new European drive cycle (NEDC) on a chassis dynamometer, with conventional cooling system (baseline engine) and MCS of a light duty diesel engine. The experimental results revealed that MCS using a VPEMT and EWP contributed to a reduction of engine warm-up period. As a consequence, important reduces in coolant flow rate and most exhaust emission compounds (THC, CO_2, CO and smoke opacity) were obtained. In contrast, NOx emission was observed to increase in these conditions. Comparative results are given for various engine speeds during a cold start and engine fully warm-up tests when the engine was equipped by conventional cooling system and MCS operation under NEDC, revealing the effect of MCS on engine fuel consumption and exhaust emissions.

Storage of HLW in engineered structures: air-cooled and water-cooled concepts

International Nuclear Information System (INIS)

Ahner, S.; Dekais, J.J.; Puttke, B.; Staner, P.

1981-01-01

A comparative study on an air-cooled and a water-cooled intermediate storage of vitrified, highly radioactive waste (HLW) in overground installations has been performed by Nukem and Belgonucleaire respectively. In the air-cooled storage concept the decay heat from the storage area will be removed using natural convection. In the water-cooled storage concept the decay heat is carried off by a primary and secondary forced-cooling system with redundant and diverse devices. The safety study carried out by Nukem used a fault tree method. It shows that the reliability of the designed water-cooled system is very high and comparable to the inherent, safe, air-cooled system. The impact for both concepts on the environment is determined by the release route, but even during accident conditions the release is far below permissible limits. The economic analysis carried out by Belgonucleaire shows that the construction costs for both systems do not differ very much, but the operation and maintenance costs for the water-cooled facility are higher than for the air cooled facility. The result of the safety and economic analysis and the discussions with the members of the working group have shown some possible significant modifications for both systems, which are included in this report. The whole study has been carried out using certain national criteria which, in certain Member States at least, would lead to a higher standard of safety than can be justified on any social, political or economic grounds

Subscale Winged Rocket Development and Application to Future Reusable Space Transportation

Directory of Open Access Journals (Sweden)

Koichi YONEMOTO

2018-03-01

Full Text Available Kyushu Institute of Technology has been studying unmanned suborbital winged rocket called WIRES (WInged REusable Sounding rocket and its research subjects concerning aerodynamics, NGC (Navigation, Guidance and Control, cryogenic composite tanks etc., and conducting flight demonstration of small winged rocket since 2005. WIRES employs the original aerodynamic shape of HIMES (HIghly Maneuverable Experimental Sounding rocket studied by ISAS (Institute of Space and Astronautical Science of JAXA (Japan Aerospace Exploration Agency in 1980s. This paper presents the preliminary design of subscale non-winged and winged rockets called WIRES#013 and WIRES#015, respectively, that are developed in collaboration with JAXA, USC (University of Southern California, UTEP (University of Texas at El Paso and Japanese industries. WIRES#013 is a conventional pre-test rocket propelled by two IPA-LOX (Isopropyl Alcohol and Liquid Oxygen engines under development by USC. It has the total length of 4.6m, and the weight of 1000kg to reach the altitude of about 6km. The flight objective is validation of the telemetry and ground communication system, recovery parachute system, and launch operation of liquid engine. WIRES#015, which has the same length of WIRES#013 and the weight of 1000kg, is a NGC technology demonstrator propelled by a fully expander-cycle LOX-Methane engine designed and developed by JAXA to reach the altitude more than 6km. The flight tests of both WIRES#013 and WIRES#015 will be conducted at the launch facility of FAR (Friends of Amateur Rocketry, Inc., which is located at Mojave Desert of California in United States of America, in May 2018 and March 2019 respectively. After completion of WIRES#015 flight tests, the suborbital demonstrator called WIRES-X will be developed and its first flight test well be performed in 2020. Its application to future fully reusable space transportation systems, such as suborbital space tour vehicles and two

Feasibility study of palm-based fuels for hybrid rocket motor applications

Science.gov (United States)

Tarmizi Ahmad, M.; Abidin, Razali; Taha, A. Latif; Anudip, Amzaryi

2018-02-01

This paper describes the combined analysis done in pure palm-based wax that can be used as solid fuel in a hybrid rocket engine. The measurement of pure palm wax calorific value was performed using a bomb calorimeter. An experimental rocket engine and static test stand facility were established. After initial measurement and calibration, repeated procedures were performed. Instrumentation supplies carried out allow fuel regression rate measurements, oxidizer mass flow rates and stearic acid rocket motors measurements. Similar tests are also carried out with stearate acid (from palm oil by-products) dissolved with nitrocellulose and bee solution. Calculated data and experiments show that rates and regression thrust can be achieved even in pure-tested palm-based wax. Additionally, palm-based wax is mixed with beeswax characterized by higher nominal melting temperatures to increase moisturizing points to higher temperatures without affecting regression rate values. Calorie measurements and ballistic experiments were performed on this new fuel formulation. This new formulation promises driving applications in a wide range of temperatures.

Extension of a simplified computer program for analysis of solid-propellant rocket motors

Science.gov (United States)

Sforzini, R. H.

1973-01-01

A research project to develop a computer program for the preliminary design and performance analysis of solid propellant rocket engines is discussed. The following capabilities are included as computer program options: (1) treatment of wagon wheel cross sectional propellant configurations alone or in combination with circular perforated grains, (2) calculation of ignition transients with the igniter treated as a small rocket engine, (3) representation of spherical circular perforated grain ends as an alternative to the conical end surface approximation used in the original program, and (4) graphical presentation of program results using a digital plotter.

Cooling performance and energy saving of a compression-absorption refrigeration system driven by a gas engine

Energy Technology Data Exchange (ETDEWEB)

Sun, Z.G.; Guo, K.H. [Sun Yat-Sen University, Guangzhou (China). Engineering School

2006-07-01

The prototype of combined vapour compression-absorption refrigeration system was set up, where a gas engine drove directly an open screw compressor in a vapour compression refrigeration chiller and waste heat from the gas engine was used to operate absorption refrigeration cycle. The experimental procedure and results showed that the combined refrigeration system was feasible. The cooling capacity of the prototype reached about 589 kW at the Chinese rated conditions of air conditioning (the inlet and outlet temperatures of chilled water are 12 and 7{sup o}C, the inlet and outlet temperatures of cooling water are 30 and 35{sup o}C, respectively). Primary energy rate (PER) and comparative primary energy saving were used to evaluate energy utilization efficiency of the combined refrigeration system. The calculated results showed that the PER of the prototype was about 1.81 and the prototype saved more than 25% of primary energy compared to a conventional electrically driven vapour compression refrigeration unit. Error analysis showed that the total error of the combined cooling system measurement was about 4.2% in this work. (author)

Re-Engineering Control Systems using Automatic Generation Tools and Process Simulation: the LHC Water Cooling Case

CERN Document Server

Booth, W; Bradu, B; Gomez Palacin, L; Quilichini, M; Willeman, D

2014-01-01

This paper presents the approach used at CERN (European Organization for Nuclear Research) to perform the re-engineering of the control systems dedicated to the LHC (Large Hadron Collider) water cooling systems.

Computational Fluid Dynamics (CFD) applications in rocket propulsion analysis and design

Science.gov (United States)

Mcconnaughey, P. K.; Garcia, R.; Griffin, L. W.; Ruf, J. H.

1993-01-01

Computational Fluid Dynamics (CFD) has been used in recent applications to affect subcomponent designs in liquid propulsion rocket engines. This paper elucidates three such applications for turbine stage, pump stage, and combustor chamber geometries. Details of these applications include the development of a high turning airfoil for a gas generator (GG) powered, liquid oxygen (LOX) turbopump, single-stage turbine using CFD as an integral part of the design process. CFD application to pump stage design has emphasized analysis of inducers, impellers, and diffuser/volute sections. Improvements in pump stage impeller discharge flow uniformity have been seen through CFD optimization on coarse grid models. In the area of combustor design, recent CFD analysis of a film cooled ablating combustion chamber has been used to quantify the interaction between film cooling rate, chamber wall contraction angle, and geometry and their effects of these quantities on local wall temperature. The results are currently guiding combustion chamber design and coolant flow rate for an upcoming subcomponent test. Critical aspects of successful integration of CFD into the design cycle includes a close-coupling of CFD and design organizations, quick turnaround of parametric analyses once a baseline CFD benchmark has been established, and the use of CFD methodology and approaches that address pertinent design issues. In this latter area, some problem details can be simplified while retaining key physical aspects to maintain analytical integrity.

NERVA-Derived Concept for a Bimodal Nuclear Thermal Rocket

International Nuclear Information System (INIS)

Fusselman, Steven P.; Frye, Patrick E.; Gunn, Stanley V.; Morrison, Calvin Q.; Borowski, Stanley K.

2005-01-01

The Nuclear Thermal Rocket is an enabling technology for human exploration missions. The 'bimodal' NTR (BNTR) provides a novel approach to meeting both propulsion and power requirements of future manned and robotic missions. The purpose of this study was to evaluate tie-tube cooling configurations, NTR performance, Brayton cycle performance, and LOX-Augmented NTR (LANTR) feasibility to arrive at a point of departure BNTR configuration for subsequent system definition

AUTOMOTIVE DIESEL MAINTENANCE 1. UNIT XXV, I--CATERPILLAR DIESEL ENGINE COOLING SYSTEM D-8 AND 824 MODELS, II--TIRES AND TIRE HARDWARE.

Science.gov (United States)

Minnesota State Dept. of Education, St. Paul. Div. of Vocational and Technical Education.

THIS MODULE OF A 30-MODULE COURSE IS DESIGNED TO DEVELOP AN UNDERSTANDING OF THE OPERATION AND MAINTENANCE OF THE DIESEL ENGINE COOLING SYSTEM AND TO PROVIDE A DESCRIPTION OF HEAVY TIRES AND WHEELS USED ON DIESEL POWERED VEHICLES. TOPICS ARE (1) THEORY OF THE COOLING SYSTEM, (2) COOLING SYSTEM COMPONENTS, (3) MAINTENANCE TIPS (COOLING SYSTEM), (4)…

NASA rocket launches student project into space

OpenAIRE

Crumbley, Liz

2005-01-01

A project that began in 2002 will culminate at sunrise on Tuesday, March 15, when a team of Virginia Tech engineering students watch a payload section they designed lift off aboard a sounding rocket from a launch pad at NASA's Wallops Island Flight Facility and travel 59 miles into space.

Determination of the Flow Field in the Propellant Tank of a Rocket Engine on Completion of the Mission

Science.gov (United States)

Fedorov, A. V.; Bedarev, I. A.; Lavruk, S. A.; Trushlyakov, V. I.; Kudentsov, V. Yu.

2018-03-01

In the present work, a method of mathematical simulation is employed to describe processes occurring in the specimens of new equipment and using the remaining propellant in rocket-engine tanks. Within the framework of certain turbulence models, the authors perform a calculation of the flow field in the volume of the tank of the launch-vehicle stage when a hot gas jet is injected into it. A vortex flow structure is revealed; the characteristics of heat transfer for different angles of injection of the jet are determined. The obtained correlation Nu = Nu(Re) satisfactorily describes experimental data.

Engineering aspect of the microwave ionosphere nonlinear interaction experiment (MINIX) with a sounding rocket

Science.gov (United States)

Nagatomo, Makoto; Kaya, Nobuyuki; Matsumoto, Hiroshi

The Microwave Ionosphere Nonlinear Interaction Experiment (MINIX) is a sounding rocket experiment to study possible effects of strong microwave fields in case it is used for energy transmission from the Solar Power Satellite (SPS) upon the Earth's atmosphere. Its secondary objective is to develop high power microwave technology for space use. Two rocket-borne magnetrons were used to emit 2.45 GHz microwave in order to make a simulated condition of power transmission from an SPS to a ground station. Sounding of the environment radiated by microwave was conducted by the diagnostic package onboard the daughter unit which was separated slowly from the mother unit. The main design drivers of this experiment were to build such high power equipments in a standard type of sounding rocket, to keep the cost within the budget and to perform a series of experiments without complete loss of the mission. The key technology for this experiment is a rocket-borne magnetron and high voltage converter. Location of position of the daughter unit relative to the mother unit was a difficult requirement for a spin-stabilized rocket. These problems were solved by application of such a low cost commercial products as a magnetron for microwave oven and a video tape recorder and camera.

Rocket Engine Turbine Blade Surface Pressure Distributions Experiment and Computations

Science.gov (United States)

Hudson, Susan T.; Zoladz, Thomas F.; Dorney, Daniel J.; Turner, James (Technical Monitor)

2002-01-01

Understanding the unsteady aspects of turbine rotor flow fields is critical to successful future turbine designs. A technology program was conducted at NASA's Marshall Space Flight Center to increase the understanding of unsteady environments for rocket engine turbines. The experimental program involved instrumenting turbine rotor blades with miniature surface mounted high frequency response pressure transducers. The turbine model was then tested to measure the unsteady pressures on the rotor blades. The data obtained from the experimental program is unique in two respects. First, much more unsteady data was obtained (several minutes per set point) than has been possible in the past. Also, an extensive steady performance database existed for the turbine model. This allowed an evaluation of the effect of the on-blade instrumentation on the turbine's performance. A three-dimensional unsteady Navier-Stokes analysis was also used to blindly predict the unsteady flow field in the turbine at the design operating conditions and at +15 degrees relative incidence to the first-stage rotor. The predicted time-averaged and unsteady pressure distributions show good agreement with the experimental data. This unique data set, the lessons learned for acquiring this type of data, and the improvements made to the data analysis and prediction tools are contributing significantly to current Space Launch Initiative turbine airflow test and blade surface pressure prediction efforts.

'RCHX-1-STORM' first Slovenian meteorological rocket program

Science.gov (United States)

Kerstein, Aleksander; Matko, Drago; Trauner, Amalija; Britovšek, Zvone

2004-08-01

Astronautic and Rocket Society Celje (ARSC) formed a special working team for research and development of a small meteorological hail suppression rocket in the 70th. The hail suppression system was established in former Yugoslavia in the late 60th as an attempt to protect important agricultural regions from one of the summer's most vicious storm. In this time Slovenia was a part of Yugoslavia as one of the federal republic with relative high developed agricultural region production. The Rocket program 'RCHX-STORM' was a second attempt, for Slovenia indigenously developed in the production of meteorological hail suppression rocket. ARSC has designed a family of small sounding rocket that were based on highly promising hybrid propellant propulsion. Hybrid propulsion was selected for this family because it was offering low cost, save production and operation and simple logistics. Conventional sounding rockets use solid propellant motor for their propulsion. The introduction of hybrid motors has enabled a considerable decrease in overall cost. The transportation handling and storage procedures were greatly simplified due to the fact that a hybrid motor was not considered as explosive matter. A hybrid motor may also be designed to stand a severe environment without resorting to conditioning arrangements. The program started in the late 70th when the team ARSC was integrated in the Research and Development Institute in Celje (RDIC). The development program aimed to produce three types of meteorological rockets with diameters 76, 120 and 160 mm. Development of the RCHX-76 engine and rocket vehicle including flight certification has been undertaken by a joint team comprising of the ARCS, RDIC and the company Cestno podjetje Celje (CPC), Road building company Celje. Many new techniques and methods were used in this program such as computer simulation of external and internal ballistics, composite materials for rocket construction, intensive static testing of models and

Radiation effect on rocket engine performance

Science.gov (United States)

Chiu, Huei-Huang; Kross, K. W.; Krebsbach, A. N.

1990-01-01

Critical problem areas involving the effect of radiation on the combustion of bipropellants are addressed by formulating a universal scaling law in combination with a radiation-enhanced vaporization combustion model. Numerical algorithms are developed and data pertaining to the Variable Thrust Engine (VTE) and the Space Shuttle Main Engine (SSME) are used to conduct parametric sensitivity studies to predict the principal intercoupling effects of radiation. The analysis reveals that low-enthalpy engines, such as the VTE, are vulnerable to a substantial performance setback due to radiative loss, whereas the performance of high-enthalpy engines such as the SSME are hardly affected over a broad range of engine operation. Combustion enhancement by radiative heating of the propellant has a significant impact on propellants with high absorptivity.

Ramjet Application Possibilities for Increasing Fire Range of the Multiple Launch Rocket Systems Ammunition

Directory of Open Access Journals (Sweden)

V. N. Zubov

2015-01-01

Full Text Available The article considers a possibility to increase a flying range of the perspective rockets equipped with the control unit with aerodynamic controllers for the multiple launch rocket systems “Smerch”.To increase a flying range and reduce a starting mass of the rocket, the paper studies a possibility to replace the single-mode rocket engine used in the solid-fuel rocket motor for the direct-flow propulsion jet engine (DFPJE with not head sector air intakes. The DFPJE is implemented according to the classical scheme with a fuel charged in the combustion chamber. A separated solid propellant starting accelerator provides the rocket acceleration to reach a speed necessary for the DFPJE to run.When designing the DFPJE a proper choice of not head air intake parameters is one of the most difficult points. For this purpose a COSMOS Flow Simulation software package and analytical dependences were used to define the following: a boundary layer thickness where an air intake is set, maximum permissible and appropriate angles of attack and deviation angles of controllers at the section where the DFPJE works, and some other parameters as well.Calculation of DFPJE characteristics consisted in determining parameters of an air-gas path of the propulsion system, geometrical sizes of the pipeline flow area, sizes of a fuel charge, and dependence of the propulsion system impulse on the flight height and speed. Calculations were performed both in thermodynamic statement of problem and in using software package of COSMOS Flow Simulation.As a result of calculations and design engineering activities the air intake profile is created and mass-dimensional characteristics of DFPJE are defined. Besides, calculations of the starting solid fuel accelerator were carried out. Further design allowed us to create the rocket shape, estimate its mass-dimensional characteristics, and perform ballistic calculations, which proved that achieving a range of 120 km for the rocket is

Significant Climate Changes Caused by Soot Emitted From Rockets in the Stratosphere

Science.gov (United States)

Mills, M. J.; Ross, M.; Toohey, D. W.

2010-12-01

A new type of hydrocarbon rocket engine with a larger soot emission index than current kerosene rockets is expected to power a fleet of suborbital rockets for commercial and scientific purposes in coming decades. At projected launch rates, emissions from these rockets will create a persistent soot layer in the northern middle stratosphere that would disproportionally affect the Earth’s atmosphere and cryosphere. A global climate model predicts that thermal forcing in the rocket soot layer will cause significant changes in the global atmospheric circulation and distributions of ozone and temperature. Tropical ozone columns decline as much as 1%, while polar ozone columns increase by up to 6%. Polar surface temperatures rise one Kelvin regionally and polar summer sea ice fractions shrink between 5 - 15%. After 20 years of suborbital rocket fleet operation, globally averaged radiative forcing (RF) from rocket soot exceeds the RF from rocket CO_{2} by six orders of magnitude, but remains small, comparable to the global RF from aviation. The response of the climate system is surprising given the small forcing, and should be investigated further with different climate models.

Combined effects of cooled EGR and a higher geometric compression ratio on thermal efficiency improvement of a downsized boosted spark-ignition direct-injection engine

International Nuclear Information System (INIS)

Su, Jianye; Xu, Min; Li, Tie; Gao, Yi; Wang, Jiasheng

2014-01-01

Highlights: • Experiments for the effects of cooled EGR and two compression ratios (CR) on fuel efficiency were conducted. • The mechanism for the observed fuel efficiency behaviors by cooled EGR and high CR was clarified. • Cooled EGR offers more fuel efficiency improvement than elevating CR from 9.3 to 10.9. • Combining 18–25% cooled EGR with 10.9 CR lead to 2.1–3.5% brake thermal efficiency improvements. - Abstract: The downsized boosted spark-ignition direct-injection (SIDI) engine has proven to be one of the most promising concepts to improve vehicle fuel economy. However, the boosted engine is typically designed at a lower geometric compression ratio (CR) due to the increased knock tendency in comparison to naturally aspirated engines, limiting the potential of improving fuel economy. On the other hand, cooled exhaust gas recirculation (EGR) has drawn attention due to the potential to suppress knock and improve fuel economy. Combing the effects of boosting, increased CR and cooled EGR to further improve fuel economy within acceptable knock tolerance has been investigated using a 2.0 L downsized boosted SIDI engine over a wide range of engine operating conditions from 1000 rpm to 3000 rpm at low to high loads. To clarify the mechanism of this complicated effects, the first law of thermodynamics analysis was conducted with the inputs from GT-Power® engine simulation. Experiment results indicate that cooled EGR provides more brake thermal efficiency improvement than increasing geometric CR from 9.3 to 10.9. The benefit of brake thermal efficiency from the higher CR is limited to low load conditions. The attributes for improving brake thermal efficiency by cooled EGR include reduced heat transfer loss, reduced pumping work and increased ratio of specific heats for all the engine operating conditions, as well as higher degree of constant volume heat release only for the knock-limited high load conditions. The combined effects of 18–25% cooled EGR

Rocket science

International Nuclear Information System (INIS)

Upson Sandra

2011-01-01

Expanding across the Solar System will require more than a simple blast off, a range of promising new propulsion technologies are being investigated by ex- NASA shuttle astronaut Chang Diaz. He is developing an alternative to chemical rockets, called VASIMR -Variable Specific Impulse Magnetoplasm Rocket. In 2012 Ad Astra plans to test a prototype, using solar power rather than nuclear, on the International Space Station. Development of this rocket for human space travel is discussed. The nuclear reactor's heat would be converted into electricity in an electric rocket such as VASIMR, and at the peak of nuclear rocket research thrust levels of almost one million newtons were reached.

Air-Powered Rockets.

Science.gov (United States)

Rodriguez, Charley; Raynovic, Jim

This document describes methods for designing and building two types of rockets--rockets from paper and rockets from bottles. Devices used for measuring the heights that the rockets obtain are also discussed. (KHR)

Contingency power for small turboshaft engines using water injection into turbine cooling air

Science.gov (United States)

Biesiadny, Thomas J.; Berger, Brett; Klann, Gary A.; Clark, David A.

1987-01-01

Because of one engine inoperative requirements, together with hot-gas reingestion and hot day, high altitude takeoff situations, power augmentation for multiengine rotorcraft has always been of critical interest. However, power augmentation using overtemperature at the turbine inlet will shorten turbine life unless a method of limiting thermal and mechanical stresses is found. A possible solution involves allowing the turbine inlet temperature to rise to augment power while injecting water into the turbine cooling air to limit hot-section metal temperatures. An experimental water injection device was installed in an engine and successfully tested. Although concern for unprotected subcomponents in the engine hot section prevented demonstration of the technique's maximum potential, it was still possible to demonstrate increases in power while maintaining nearly constant turbine rotor blade temperature.

Thrust Vector Control of an Upper-Stage Rocket with Multiple Propellant Slosh Modes

Directory of Open Access Journals (Sweden)

Jaime Rubio Hervas

2012-01-01

Full Text Available The thrust vector control problem for an upper-stage rocket with propellant slosh dynamics is considered. The control inputs are defined by the gimbal deflection angle of a main engine and a pitching moment about the center of mass of the spacecraft. The rocket acceleration due to the main engine thrust is assumed to be large enough so that surface tension forces do not significantly affect the propellant motion during main engine burns. A multi-mass-spring model of the sloshing fuel is introduced to represent the prominent sloshing modes. A nonlinear feedback controller is designed to control the translational velocity vector and the attitude of the spacecraft, while suppressing the sloshing modes. The effectiveness of the controller is illustrated through a simulation example.

X-ray Radiography Measurements of Shear Coaxial Rocket Injectors

Science.gov (United States)

2013-05-07

Shear coaxial jets can be found in a number of combustion devices – Turbofan engine exhaust , air blast furnaces, and liquid rocket engines ...water and gaseous nitro-gen as propellant simulants at atmospheric backpressure , the effect of momentum flux ratio and mass flux ratio, are...the effect of momentum flux ratio, mass flux ratio and post thickness on the liquid mass distribution – Use quantitative centerline profiles to

Saving Humanity from Catastrophic Cooling with Geo-Engineering

Science.gov (United States)

Haapala, K.; Singer, S. F.

2016-02-01

There are two kinds of ice ages; they are fundamentally different and therefore require different methods of mitigation: (i) Major (Milankovitch-style) glaciations occur on a 100,000-year time-scale and are controlled astronomically. (ii) "Little" ice ages were discovered in ice cores; they have been occurring on an approx. 1000-1500-yr cycle and are likely controlled by the Sun [Ref: Singer & Avery 2007. Unstoppable Global Warming: Every 1500 years]. The current cycle's cooling phase may be imminent - hence there may be urgent need for action. To stop onset of a major (Milankovitch) glaciation 1. Locate a "trigger" - a growing perennial snow/ice field - using satellites 2. Spread soot, to lower the albedo, and use Sun to melt 3. How much soot? How to apply soot? Learn by experimentation To lessen (regional, intermittent) cooling of DOB (Dansgaard-Oeschger-Bond) cycles1. Use greenhouse effect of manmade cirrus (ice particles) [Ref: Singer 1988. Meteorology and Atmospheric Physics 38:228 - 239]2. Inject water droplets (mist) near tropopause 3. Trace dispersion of cirrus cloud by satellite and observe warming at surface 4. How much water; over what area? How often to inject? Learn by experimentation Many scientific questions remain. While certainly interesting and important, there is really no need to delay the crucial and urgent tests of geo-engineering, designed to validate schemes of mitigation. Such proposed tests involve only minor costs and present negligible risks to the environment.

Evaluation of heat exchange performance for the auxiliary component cooling water system cooling tower in HTTR

International Nuclear Information System (INIS)

Tochio, Daisuke; Kameyama, Yasuhiko; Shimizu, Atsushi; Inoi, Hiroyuki; Yamazaki, Kazunori; Shimizu, Yasunori; Aragaki, Etsushi; Ota, Yukimaru; Fujimoto, Nozomu

2006-09-01

The auxiliary component cooling water system (ACCWS) is one of the cooling system in High Temperature Engineering Test Reactor (HTTR). The ACCWS has main two features, many facilities cooling, and heat sink of the vessel cooling system which is one of the engineering safety features. Therefore, the ACCWS is required to satisfy the design criteria of heat removal performance. In this report, heat exchange performance data of the rise-to-power-up test and the in-service operation for the ACCWS cooling tower was evaluated. Moreover, the evaluated values were compared with the design values, and it is confirmed that ACCWS cooling tower has the required heat exchange performance in the design. (author)

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.

Verification on spray simulation of a pintle injector for liquid rocket engine

Science.gov (United States)

Son, Min; Yu, Kijeong; Radhakrishnan, Kanmaniraja; Shin, Bongchul; Koo, Jaye

2016-02-01

The pintle injector used for a liquid rocket engine is a newly re-attracted injection system famous for its wide throttle ability with high efficiency. The pintle injector has many variations with complex inner structures due to its moving parts. In order to study the rotating flow near the injector tip, which was observed from the cold flow experiment using water and air, a numerical simulation was adopted and a verification of the numerical model was later conducted. For the verification process, three types of experimental data including velocity distributions of gas flows, spray angles and liquid distribution were all compared using simulated results. The numerical simulation was performed using a commercial simulation program with the Eulerian multiphase model and axisymmetric two dimensional grids. The maximum and minimum velocities of gas were within the acceptable range of agreement, however, the spray angles experienced up to 25% error when the momentum ratios were increased. The spray density distributions were quantitatively measured and had good agreement. As a result of this study, it was concluded that the simulation method was properly constructed to study specific flow characteristics of the pintle injector despite having the limitations of two dimensional and coarse grids.

Multi-Stage Hybrid Rocket Conceptual Design for Micro-Satellites Launch using Genetic Algorithm

Science.gov (United States)

Kitagawa, Yosuke; Kitagawa, Koki; Nakamiya, Masaki; Kanazaki, Masahiro; Shimada, Toru

The multi-objective genetic algorithm (MOGA) is applied to the multi-disciplinary conceptual design problem for a three-stage launch vehicle (LV) with a hybrid rocket engine (HRE). MOGA is an optimization tool used for multi-objective problems. The parallel coordinate plot (PCP), which is a data mining method, is employed in the post-process in MOGA for design knowledge discovery. A rocket that can deliver observing micro-satellites to the sun-synchronous orbit (SSO) is designed. It consists of an oxidizer tank containing liquid oxidizer, a combustion chamber containing solid fuel, a pressurizing tank and a nozzle. The objective functions considered in this study are to minimize the total mass of the rocket and to maximize the ratio of the payload mass to the total mass. To calculate the thrust and the engine size, the regression rate is estimated based on an empirical model for a paraffin (FT-0070) propellant. Several non-dominated solutions are obtained using MOGA, and design knowledge is discovered for the present hybrid rocket design problem using a PCP analysis. As a result, substantial knowledge on the design of an LV with an HRE is obtained for use in space transportation.

Effect of cooled EGR on performance and exhaust gas emissions in EFI spark ignition engine fueled by gasoline and wet methanol blends

Science.gov (United States)

Rohadi, Heru; Syaiful, Bae, Myung-Whan

2016-06-01

Fuel needs, especially the transport sector is still dominated by fossil fuels which are non-renewable. However, oil reserves are very limited. Furthermore, the hazardous components produced by internal combustion engine forces many researchers to consider with alternative fuel which is environmental friendly and renewable sources. Therefore, this study intends to investigate the impact of cooled EGR on the performance and exhaust gas emissions in the gasoline engine fueled by gasoline and wet methanol blends. The percentage of wet methanol blended with gasoline is in the range of 5 to 15% in a volume base. The experiment was performed at the variation of engine speeds from 2500 to 4000 rpm with 500 intervals. The re-circulated exhaust gasses into combustion chamber was 5%. The experiment was performed at the constant engine speed. The results show that the use of cooled EGR with wet methanol of 10% increases the brake torque up to 21.3%. The brake thermal efficiency increases approximately 39.6% using cooled EGR in the case of the engine fueled by 15% wet methanol. Brake specific fuel consumption for the engine using EGR fueled by 10% wet methanol decreases up to 23% at the engine speed of 2500 rpm. The reduction of CO, O2 and HC emissions was found, while CO2 increases.

The nuclear thermal electric rocket: a proposed innovative propulsion concept for manned interplanetary missions

Science.gov (United States)

Dujarric, C.; Santovincenzo, A.; Summerer, L.

2013-03-01

Conventional propulsion technology (chemical and electric) currently limits the possibilities for human space exploration to the neighborhood of the Earth. If farther destinations (such as Mars) are to be reached with humans on board, a more capable interplanetary transfer engine featuring high thrust, high specific impulse is required. The source of energy which could in principle best meet these engine requirements is nuclear thermal. However, the nuclear thermal rocket technology is not yet ready for flight application. The development of new materials which is necessary for the nuclear core will require further testing on ground of full-scale nuclear rocket engines. Such testing is a powerful inhibitor to the nuclear rocket development, as the risks of nuclear contamination of the environment cannot be entirely avoided with current concepts. Alongside already further matured activities in the field of space nuclear power sources for generating on-board power, a low level investigation on nuclear propulsion has been running since long within ESA, and innovative concepts have already been proposed at an IAF conference in 1999 [1, 2]. Following a slow maturation process, a new concept was defined which was submitted to a concurrent design exercise in ESTEC in 2007. Great care was taken in the selection of the design parameters to ensure that this quite innovative concept would in all respects likely be feasible with margins. However, a thorough feasibility demonstration will require a more detailed design including the selection of appropriate materials and the verification that these can withstand the expected mechanical, thermal, and chemical environment. So far, the predefinition work made clear that, based on conservative technology assumptions, a specific impulse of 920 s could be obtained with a thrust of 110 kN. Despite the heavy engine dry mass, a preliminary mission analysis using conservative assumptions showed that the concept was reducing the required

The Guggenheim Aeronautics Laboratory at Caltech and the creation of the modern rocket motor (1936-1946): How the dynamics of rocket theory became reality

Science.gov (United States)

Zibit, Benjamin Seth

This thesis explores and unfolds the story of discovery in rocketry at The California Institute of Technology---specifically at Caltech's Guggenheim Aeronautics Laboratory---in the 1930s and 1940s. Caltech was home to a small group of engineering students and experimenters who, beginning in the winter of 1935--1936, formed a study and research team destined to change the face of rocket science in the United States. The group, known as the Guggenheim Aeronautics Laboratory (GALCIT, for short) Rocket Research Group, invented a new type of solid-rocket propellant, made distinct and influential discoveries in the theory of rocket combustion and design, founded the Jet Propulsion Laboratory, and incorporated the first American industrial concern devoted entirely to rocket motor production: The Aerojet Corporation. The theoretical work of team members, Frank Malina, Hsueh-shen Tsien, Homer J. Stewart, and Mark Mills, is examined in this thesis in detail. The author scrutinizes Frank Malina's doctoral thesis (both its assumptions and its mathematics), and finds that, although Malina's key assertions, his formulae, hold, his work is shown to make key assumptions about rocket dynamics which only stand the test of validity if certain approximations, rather than exact measurements, are accepted. Malina studied the important connection between motor-nozzle design and thrust; in his Ph.D. thesis, he developed mathematical statements which more precisely defined the design/thrust relation. One of Malina's colleagues on the Rocket Research Team, John Whiteside Parsons, created a new type of solid propellant in the winter of 1941--1942. This propellant, known as a composite propellant (because it simply was a relatively inert amalgam of propellant and oxidizer in non-powder form), became the forerunner of all modern solid propellants, and has become one of the seminal discoveries in the field of Twentieth Century rocketry. The latter chapters of this dissertation discuss the

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.

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.

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.

Experimental characterization of cooled EGR in a gasoline direct injection engine for reducing fuel consumption and nitrogen oxide emission

Science.gov (United States)

Park, Sang-Ki; Lee, Jungkoo; Kim, Kyungcheol; Park, Seongho; Kim, Hyung-Man

2015-11-01

The emphasis on increasing fuel economy and reducing emissions is increasing. Attention has turned to how the performance of a gasoline direct injection (GDI) engine can be improved to achieve lower fuel consumption and NOx emission. Therefore, positive effects can reduce fuel consumption and NOx emission as well as knock suppression. The cooled exhaust gas recirculation (EGR) ranges within the characteristic map are characterized from the experimental results at various speeds and brake mean effective pressures in a GDI engine. The results show that the application of cooled EGR system brought in 3.63 % reduction as for the fuel consumption and 4.34 % as for NOx emission.

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.

Determination of the availability of appropriate aged flight rocket motors. [captive tests to determine case bond separation and grain bore cracking

Science.gov (United States)

Martin, P. J.

1974-01-01

A program to identify surplus solid rocket propellant engines which would be available for a program of functional integrity testing was conducted. The engines are classified as: (1) upper stage and apogee engines, (2) sounding rocket and launch vehicle engines, and (3) jato, sled, and tactical engines. Nearly all the engines were available because their age exceeds the warranted shelf life. The preference for testing included tests at nominal flight conditions, at design limits, and to establish margin limits. The principal failure modes of interest were case bond separation and grain bore cracking. Data concerning the identification and characteristics of each engine are tabulated. Methods for conducting the tests are described.

SSTO rockets. A practical possibility

Science.gov (United States)

Bekey, Ivan

1994-07-01

Most experts agree that single-stage-to-orbit (SSTO) rockets would become feasible if more advanced technologies were available to reduce the vehicle dry weight, increase propulsion system performance, or both. However, these technologies are usually judged to be very ambitious and very far off. This notion persists despite major advances in technology and vehicle design in the past decade. There appears to be four major misperceptions about SSTOs, regarding their mass fraction, their presumed inadequate performance margin, their supposedly small payloads, and their extreme sensitivity to unanticipated vehicle weight growth. These misperceptions can be dispelled for SSTO rockets using advanced technologies that could be matured and demonstrated in the near term. These include a graphite-composite primary structure, graphite-composite and Al-Li propellant tanks with integral reusable thermal protection, long-life tripropellant or LOX-hydrogen engines, and several technologies related to operational effectiveness, including vehicle health monitoring, autonomous avionics/flight control, and operable launch and ground handling systems.

Silent cooling system with a double suction type centrifugal fan with backward blades; Ryosuikomi turbo fan wo mochiita teisoon engine reikyaku system no kaihatsu

Energy Technology Data Exchange (ETDEWEB)

Funabashi, S. [Hitachi, Ltd., Tokyo (Japan); Nakamura, K. [Hitachi Construction Machinery Co. Ltd., Tokyo (Japan)

2000-04-25

We have developed a new cooling system for the engine of construction machinery. This system produces two different air flow routes by a double suction type centrifugal fan with backward blades. We first measured aerodynamic performance and sound level of the double suction type fan, which consists of two single suction type fans of different design, and compared these measurements to those of single suction type fans. Next, we installed the fan in a scale model of the new cooling system. The performance of this double suction type fan was different to that of a single suction type. It is considered that the effect of interaction of the exit flows from the two different fans caused this performance change. The test model installed in an engine compartment showed that this cooling system reduced the area of air inlets and outlets around the engine and did not cause re-circulation of cooling air. (author)

On synthesis and optimization of cooling water systems with multiple cooling towers

CSIR Research Space (South Africa)

Gololo, KV

2011-01-01

Full Text Available -1 On Synthesis and Optimization of Cooling Water Systems with Multiple Cooling Towers Khunedi Vincent Gololo?? and Thokozani Majozi*? ? Department of Chemical Engineering, University of Pretoria, Lynnwood Road, Pretoria, 0002, South Africa ? Modelling...

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

An investigation of heat recovery of submarine diesel engines for combined cooling, heating and power systems

International Nuclear Information System (INIS)

Daghigh, Roonak; Shafieian, Abdellah

2016-01-01

Highlights: • The power output of the cycle is about 53 kW in the mass flow rate of 0.6 kg/s. • The output cooling water temperature of evaporator is 3.64 °C. • The absorption chiller has a coefficient of performance equal to 0.94. - Abstract: High temperature and mass flow rate of the exhaust gases of submarine diesel engines provide an appropriate potential for their thermal recovery. The current study introduces a combined cooling, heating and power system for thermal recovery of submarine diesel engines. The cooling system is composed of a mixed effect absorption chiller with two high and low pressure generators. The exhaust of the diesel engine is used in the high pressure generator, and the low pressure generator was divided into two parts. The required heat for the first and second compartments is supplied by the cooling water of the engine and condensation of the vapor generated in the high pressure generator, respectively. The power generation system is a Rankine cycle with an organic working fluid, which is considered a normal thermal system to supply hot water. The whole system is encoded based on mass stability, condensation and energy equations. The obtained findings showed that the maximum heat recovery for the power cycle occurs in exhaust gas mass ratio of 0.23–0.29 and working fluid mass flow rate of 0.45–0.57 kg/s. Further, for each specific mass ratio of exhaust gas, only a certain range of working fluid mass flow rate is used. In the refrigerant mass flow rate of 0.6 kg/s and exhaust gas mass ratio of 0.27, the power output of the cycle is 53 kW, which can also be achieved by simultaneous increase of refrigerant mass flow rate and exhaust gas mass ratio in a certain range of higher powers. In the next section, the overall distribution diagram of output water temperature of the thermal system is obtained according to the exhaust gas mass ratio in various mass flow rates, which can increase the potential of designing and controlling the

Real-Time Rocket/Vehicle System Integrated Health Management Laboratory For Development and Testing of Health Monitoring/Management Systems

Science.gov (United States)

Aguilar, R.

2006-01-01

Pratt & Whitney Rocketdyne has developed a real-time engine/vehicle system integrated health management laboratory, or testbed, for developing and testing health management system concepts. This laboratory simulates components of an integrated system such as the rocket engine, rocket engine controller, vehicle or test controller, as well as a health management computer on separate general purpose computers. These general purpose computers can be replaced with more realistic components such as actual electronic controllers and valve actuators for hardware-in-the-loop simulation. Various engine configurations and propellant combinations are available. Fault or failure insertion capability on-the-fly using direct memory insertion from a user console is used to test system detection and response. The laboratory is currently capable of simulating the flow-path of a single rocket engine but work is underway to include structural and multiengine simulation capability as well as a dedicated data acquisition system. The ultimate goal is to simulate as accurately and realistically as possible the environment in which the health management system will operate including noise, dynamic response of the engine/engine controller, sensor time delays, and asynchronous operation of the various components. The rationale for the laboratory is also discussed including limited alternatives for demonstrating the effectiveness and safety of a flight system.

An Analysis of Rocket Propulsion Testing Costs

Science.gov (United States)

Ramirez, Carmen; Rahman, Shamim

2010-01-01

The primary mission at NASA Stennis Space Center (SSC) is rocket propulsion testing. Such testing is commonly characterized as one of two types: production testing for certification and acceptance of engine hardware, and developmental testing for prototype evaluation or research and development (R&D) purposes. For programmatic reasons there is a continuing need to assess and evaluate the test costs for the various types of test campaigns that involve liquid rocket propellant test articles. Presently, in fact, there is a critical need to provide guidance on what represents a best value for testing and provide some key economic insights for decision-makers within NASA and the test customers outside the Agency. Hence, selected rocket propulsion test databases and references have been evaluated and analyzed with the intent to discover correlations of technical information and test costs that could help produce more reliable and accurate cost projections in the future. The process of searching, collecting, and validating propulsion test cost information presented some unique obstacles which then led to a set of recommendations for improvement in order to facilitate future cost information gathering and analysis. In summary, this historical account and evaluation of rocket propulsion test cost information will enhance understanding of the various kinds of project cost information; identify certain trends of interest to the aerospace testing community.

History of Solid Rockets

Science.gov (United States)

Green, Rebecca

2017-01-01

Solid rockets are of interest to the space program because they are commonly used as boosters that provide the additional thrust needed for the space launch vehicle to escape the gravitational pull of the Earth. Larger, more advanced solid rockets allow for space launch vehicles with larger payload capacities, enabling mankind to reach new depths of space. This presentation will discuss, in detail, the history of solid rockets. The history begins with the invention and origin of the solid rocket, and then goes into the early uses and design of the solid rocket. The evolution of solid rockets is depicted by a description of how solid rockets changed and improved and how they were used throughout the 16th, 17th, 18th, and 19th centuries. Modern uses of the solid rocket include the Solid Rocket Boosters (SRBs) on the Space Shuttle and the solid rockets used on current space launch vehicles. The functions and design of the SRB and the advancements in solid rocket technology since the use of the SRB are discussed as well. Common failure modes and design difficulties are discussed as well.

Prediction of cold start hydrocarbon emissions of air cooled two wheeler spark ignition engines by simple fuzzy logic simulation

Directory of Open Access Journals (Sweden)

Samuel Raja Ayyanan

2014-01-01

Full Text Available The cold start hydrocarbon emission from the increasing population of two wheelers in countries like India is one of the research issues to be addressed. This work describes the prediction of cold start hydrocarbon emissions from air cooled spark ignition engines through fuzzy logic technique. Hydrocarbon emissions were experimentally measured from test engines of different cubic capacity, at different lubricating oil temperature and at different idling speeds with and without secondary air supply in exhaust. The experimental data were used as input for modeling average hydrocarbon emissions for 180 seconds counted from cold start and warm start of gasoline bike engines. In fuzzy logic simulation, member functions were assigned for input variables (cubic capacity and idling rpm and output variables (average hydrocarbon emission for first 180 seconds at cold start and warm start. The knowledge based rules were adopted from the analyzed experimental data and separate simulations were carried out for predicting hydrocarbon emissions from engines equipped with and without secondary air supply. The simulation yielded the average hydrocarbon emissions of air cooled gasoline engine for a set of given input data with accuracy over 90%.

Modeling Transients and Designing a Passive Safety System for a Nuclear Thermal Rocket Using Relap5

Science.gov (United States)

Khatry, Jivan

Long-term high payload missions necessitate the need for nuclear space propulsion. Several nuclear reactor types were investigated by the Nuclear Engine for Rocket Vehicle Application (NERVA) program of National Aeronautics and Space Administration (NASA). Study of planned/unplanned transients on nuclear thermal rockets is important due to the need for long-term missions. A NERVA design known as the Pewee I was selected for this purpose. The following transients were run: (i) modeling of corrosion-induced blockages on the peripheral fuel element coolant channels and their impact on radiation heat transfer in the core, and (ii) modeling of loss-of-flow-accidents (LOFAs) and their impact on radiation heat transfer in the core. For part (i), the radiation heat transfer rate of blocked channels increases while their neighbors' decreases. For part (ii), the core radiation heat transfer rate increases while the flow rate through the rocket system is decreased. However, the radiation heat transfer decreased while there was a complete LOFA. In this situation, the peripheral fuel element coolant channels handle the majority of the radiation heat transfer. Recognizing the LOFA as the most severe design basis accident, a passive safety system was designed in order to respond to such a transient. This design utilizes the already existing tie rod tubes and connects them to a radiator in a closed loop. Hence, this is basically a secondary loop. The size of the core is unchanged. During normal steady-state operation, this secondary loop keeps the moderator cool. Results show that the safety system is able to remove the decay heat and prevent the fuel elements from melting, in response to a LOFA and subsequent SCRAM.

Reducing the CO{sub 2} emission of a turbocharged DI gasoline engine by using cooled EGR and optimized cooling system control; Reduzierung der CO{sub 2}-Emissionen eines Turbo-DI-Ottomotors durch gekuehlte AGR und optimierte Regelung des Kuehlsystems

Energy Technology Data Exchange (ETDEWEB)

Mueller, Rolf; Feldhaus, Georg; Pantow, Eberhard; Edwards, Simon [Behr GmbH, Stuttgart (Germany); Finkeldei, Thomas; Trapp, Ralph [BHTC GmbH, Lippstadt (Germany); Neubauer, Matthias; Kappus, Paul [AVL List GmbH, Graz (Austria)

2007-07-01

In a joint project between the companies Behr, BHTC and AVL List different technologies for the reduction of the fuel consumption hence the CO{sub 2} emissions were studied on a turbocharged DI gasoline engine and in a vehicle. With respect to the engine, cooled external exhaust gas recirculation was examined. The fuel consumption in typical part load engine and vehicle operation was lowered by up to 5% and within the full load region, through the prevention of fuel enrichment, by up to 18%. With respect to the thermal management coolant stand still shows a fuel consumption reduction potential during the warm-up phase of the engine of about 3% in the NEDC. In engine tests it was shown how a variation of the coolant temperature by up to 10 K, with the help of a map controlled thermostat, improves the fuel efficiency by up to 1,4%. The control aspects of the cooling system (coolant stand still, temperature adjustment and indirect charge air cooling) were implemented in the vehicle with the use of a Rapid Control Prototyping system. (orig.)

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

Transient Side Load Analysis of Out-of-Round Film-Cooled Nozzle Extensions

Science.gov (United States)

Wang, Ten-See; Lin, Jeff; Ruf, Joe; Guidos, Mike

2012-01-01

There was interest in understanding the impact of out-of-round nozzle extension on the nozzle side load during transient startup operations. The out-of-round nozzle extension could be the result of asymmetric internal stresses, deformation induced by previous tests, and asymmetric loads induced by hardware attached to the nozzle. The objective of this study was therefore to computationally investigate the effect of out-of-round nozzle extension on the nozzle side loads during an engine startup transient. The rocket engine studied encompasses a regeneratively cooled chamber and nozzle, along with a film cooled nozzle extension. The computational methodology is based on an unstructured-grid, pressure-based computational fluid dynamics formulation, and transient inlet boundary flow properties derived from an engine system simulation. Six three-dimensional cases were performed with the out-of-roundness achieved by three different degrees of ovalization, elongated on lateral y and z axes: one slightly out-of-round, one more out-of-round, and one significantly out-of-round. The results show that the separation line jump was the primary source of the peak side loads. Comparing to the peak side load of the perfectly round nozzle, the peak side loads increased for the slightly and more ovalized nozzle extensions, and either increased or decreased for the two significantly ovalized nozzle extensions. A theory based on the counteraction of the flow destabilizing effect of an exacerbated asymmetrical flow caused by a lower degree of ovalization, and the flow stabilizing effect of a more symmetrical flow, created also by ovalization, is presented to explain the observations obtained in this effort.

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

Temperature control of an automotive engine cooling system utilizing a magneto-rheological fan clutch

International Nuclear Information System (INIS)

Kim, Eun-Seok; Choi, Seung-Bok; Park, Young-Gee; Lee, Soojin

2010-01-01

In this note, the temperature control of an automotive engine cooling system is undertaken using a magneto-rheological (MR) fluid-based fan clutch (MR fan clutch in short). In order to achieve this goal, an appropriate size of controllable fan clutch using an MR fluid is firstly devised by considering the design parameters of a conventional fan clutch to reflect the practical application. Then, the principal design parameters of the MR fan clutch such as the length of the disc are optimally determined through finite element analysis. The drum-type MR fan clutch is manufactured and its time response to input current is experimentally evaluated. A robust sliding mode controller is then formulated by treating the time constant of the fan clutch system as an uncertain parameter. After identifying the relationship between angular velocity of the MR fan clutch and the temperature of the cooling system, the sliding mode controller is experimentally realized for the cooling system. It has been clearly demonstrated that the proposed sliding mode controller follows well the desired temperature with a small regulating error. It is expected from this feasibility work that the proposed control system associated with an MR fan clutch can be effectively utilized for the automotive cooling system to improve the fuel efficiency. (technical note)

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

to determine what parameters drive unsteadiness in fuel films, and how these parameters affect wall temperature profiles. Parametric studies performed...temperature profiles. Parametric studies performed in 2D suggest that a Helmholtz resonator exists for simple slot geometries. Frequencies in 3D were...effect on film cooling effectiveness. In general, the heat flux exhibits complex trends and did not scale well with chamber pressure. ∗Aerospace

Laser-Induced Emissions Sensor for Soot Mass in Rocket Plumes, Phase I

Data.gov (United States)

National Aeronautics and Space Administration — A method is proposed to measure soot mass concentration non-intrusively from a distance in a rocket engine exhaust stream during ground tests using laser-induced...

A History of Welding on the Space Shuttle Main Engine (1975 to 2010)

Science.gov (United States)

Zimmerman, Frank R.; Russell, Carolyn K.

2010-01-01

The Space Shuttle Main Engine (SSME) is a high performance, throttleable, liquid hydrogen fueled rocket engine. High thrust and specific impulse (Isp) are achieved through a staged combustion engine cycle, combined with high combustion pressure (approx.3000psi) generated by the two-stage pump and combustion process. The SSME is continuously throttleable from 67% to 109% of design thrust level. The design criteria for this engine maximize performance and weight, resulting in a 7,800 pound rocket engine that produces over a half million pounds of thrust in vacuum with a specific impulse of 452/sec. It is the most reliable rocket engine in the world, accumulating over one million seconds of hot-fire time and achieving 100% flight success in the Space Shuttle program. A rocket engine with the unique combination of high reliability, performance, and reusability comes at the expense of manufacturing simplicity. Several innovative design features and fabrication techniques are unique to this engine. This is as true for welding as any other manufacturing process. For many of the weld joints it seemed mean cheating physics and metallurgy to meet the requirements. This paper will present a history of the welding used to produce the world s highest performance throttleable rocket engine.

Potential climate impact of black carbon emitted by rockets

Science.gov (United States)

Ross, Martin; Mills, Michael; Toohey, Darin

2010-12-01

A new type of hydrocarbon rocket engine is expected to power a fleet of suborbital rockets for commercial and scientific purposes in coming decades. A global climate model predicts that emissions from a fleet of 1000 launches per year of suborbital rockets would create a persistent layer of black carbon particles in the northern stratosphere that could cause potentially significant changes in the global atmospheric circulation and distributions of ozone and temperature. Tropical stratospheric ozone abundances are predicted to change as much as 1%, while polar ozone changes by up to 6%. Polar surface temperatures change as much as one degree K regionally with significant impacts on polar sea ice fractions. After one decade of continuous launches, globally averaged radiative forcing from the black carbon would exceed the forcing from the emitted CO2 by a factor of about 105 and would be comparable to the radiative forcing estimated from current subsonic aviation.

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…

Thermodynamic analysis of an absorption refrigeration system used to cool down the intake air in an Internal Combustion Engine

International Nuclear Information System (INIS)

Novella, R.; Dolz, V.; Martín, J.; Royo-Pascual, L.

2017-01-01

Highlights: • Enough power in the exhaust gases is available to operate the absorption cycle. • Three engine operating points are presented in the article. • Improvement potential up to 4% is possible in the engine indicated efficiency. • Engine indicated efficiency benefit was experimentally confirmed by direct testing. - Abstract: This paper deals with the thermodynamic analysis of an absorption refrigeration cycle used to cool down the temperature of the intake air in an Internal Combustion Engine using as a heat source the exhaust gas of the engine. The solution of ammonia-water has been selected due to the stability for a wide range of operating temperatures and pressures and the low freezing point. The effects of operating temperatures, pressures, concentrations of strong and weak solutions in the absorption refrigeration cycle were examined to achieve proper heat rejection to the ambient. Potential of increasing Internal Combustion Engine efficiency and reduce pollutant emissions was estimated by means of theoretical models and experimental tests. In order to provide boundary conditions for the absorption refrigeration cycle and to simulate its effect on engine performance, a 0D thermodynamic model was used to reproduce the engine performance when the intake air is cooled. Furthermore, a detailed experimental work was carried out to validate the results in real engine operation. Theoretical results show how the absorption refrigeration system decreases the intake air flow temperature down to a temperature around 5 °C and even lower by using the bottoming waste heat energy available in the exhaust gases in a wide range of engine operating conditions. In addition, the theoretical analysis estimates the potential of the strategy for increasing the engine indicated efficiency in levels up to 4% also at the operating conditions under evaluation. Finally, this predicted benefit in engine indicated efficiency has been experimentally confirmed by direct

Embedded expert system for space shuttle main engine maintenance

Science.gov (United States)

Pooley, J.; Thompson, W.; Homsley, T.; Teoh, W.; Jones, J.; Lewallen, P.

1987-01-01

The SPARTA Embedded Expert System (SEES) is an intelligent health monitoring system that directs analysis by placing confidence factors on possible engine status and then recommends a course of action to an engineer or engine controller. The technique can prevent catastropic failures or costly rocket engine down time because of false alarms. Further, the SEES has potential as an on-board flight monitor for reusable rocket engine systems. The SEES methodology synergistically integrates vibration analysis, pattern recognition and communications theory techniques with an artificial intelligence technique - the Embedded Expert System (EES).

Thermodynamic modelling and performance study of an engine waste heat driven adsorption cooling for automotive air-conditioning

International Nuclear Information System (INIS)

Ali, Syed Muztuza; Chakraborty, Anutosh

2015-01-01

Waste heat from engine can be utilized to drive an adsorption cooling system for air conditioning purposes in the vehicle cabin, which not only improves the fuel economy but also reduces the carbon footprint. It is also important to reduce the size of the adsorption bed to adopt the adsorption technology for air-conditioning applications in passenger cars, buses and trucks or even trains. In this article, we present a two stage indirect exhaust heat recovery system of automotive engine employing an effective lumped parameter model to simulate the dynamic behaviors of an adsorption chiller that ranges from the transient to the cyclic steady states. The thermodynamic framework of adsorption chiller is developed from the rigor of mass and energy balances of each component of the system and experimentally confirmed isotherms and kinetics data of various adsorbent–adsorbate pairs. The performance factors are calculated in terms of COP (Coefficient of Performance) and SCP (Specific Cooling Power) for different operating parameters such as cycle time, exhaust gas temperatures, cooling water temperatures and flow rates. From the simulation results, it is found that the exhaust energy of a six cylinder 3000 cc private car is able to produce nearly 3 kW of cooling power for the car cabin. It is also observed that the driving heat source temperature does not remain constant throughout the cycle time unlike the conventional adsorption chiller, and the hot water temperatures as driving source vary from 65 to 95 °C. CaCl 2 -in-silica gel–water system is found better in terms of COP and SCP as compared with other adsorbents – water systems. - Highlights: • Adsorption cooling for car air conditioning. • Thermodynamic frameworks with adsorption isotherms and kinetics. • Various adsorbents such as silica gel, zeolites (AQSOA-Z01, Z-02), CaCl 2 -in-silica gel are tested. • Cooling power for car cabin employing waste heat recovery.

Advanced Vacuum Plasma Spray (VPS) for a Robust, Longlife and Safe Space Shuttle Main Engine (SSME)

Science.gov (United States)

Holmes, Richard R.; Elam, Sandra K.; McKechnie, Timothy N.; Power, Christopher A.

2010-01-01

In 1984, the Vacuum Plasma Spray Lab was built at NASA/Marshall Space Flight Center for applying durable, protective coatings to turbine blades for the space shuttle main engine (SSME) high pressure fuel turbopump. Existing turbine blades were cracking and breaking off after five hot fire tests while VPS coated turbine blades showed no wear or cracking after 40 hot fire tests. Following that, a major manufacturing problem of copper coatings peeling off the SSME Titanium Main Fuel Valve Housing was corrected with a tenacious VPS copper coating. A patented VPS process utilizing Functional Gradient Material (FGM) application was developed to build ceramic lined metallic cartridges for space furnace experiments, safely containing gallium arsenide at 1260 degrees centigrade. The VPS/FGM process was then translated to build robust, long life, liquid rocket combustion chambers for the space shuttle main engine. A 5K (5,000 Lb. thrust) thruster with the VPS/FGM protective coating experienced 220 hot firing tests in pristine condition with no wear compared to the SSME which showed blanching (surface pulverization) and cooling channel cracks in less than 30 of the same hot firing tests. After 35 of the hot firing tests, the injector face plates disintegrated. The VPS/FGM process was then applied to spraying protective thermal barrier coatings on the face plates which showed 50% cooler operating temperature, with no wear after 50 hot fire tests. Cooling channels were closed out in two weeks, compared to one year for the SSME. Working up the TRL (Technology Readiness Level) to establish the VPS/FGM process as viable technology, a 40K thruster was built and is currently being tested. Proposed is to build a J-2X size liquid rocket engine as the final step in establishing the VPS/FGM process TRL for space flight.

Recent Experimental Efforts on High-Pressure Supercritical Injection for Liquid Rockets and Their Implications

Directory of Open Access Journals (Sweden)

Bruce Chehroudi

2012-01-01

Full Text Available Pressure and temperature of the liquid rocket thrust chambers into which propellants are injected have been in an ascending trajectory to gain higher specific impulse. It is quite possible then that the thermodynamic condition into which liquid propellants are injected reaches or surpasses the critical point of one or more of the injected fluids. For example, in cryogenic hydrogen/oxygen liquid rocket engines, such as Space Shuttle Main Engine (SSME or Vulcain (Ariane 5, the injected liquid oxygen finds itself in a supercritical condition. Very little detailed information was available on the behavior of liquid jets under such a harsh environment nearly two decades ago. The author had the opportunity to be intimately involved in the evolutionary understanding of injection processes at the Air Force Research Laboratory (AFRL, spanning sub- to supercritical conditions during this period. The information included here attempts to present a coherent summary of experimental achievements pertinent to liquid rockets, focusing only on the injection of nonreacting cryogenic liquids into a high-pressure environment surpassing the critical point of at least one of the propellants. Moreover, some implications of the results acquired under such an environment are offered in the context of the liquid rocket combustion instability problem.

Contingency power for a small turboshaft engine by using water injection into turbine cooling air

Science.gov (United States)

Biesiadny, Thomas J.; Klann, Gary A.

1992-01-01

Because of one-engine-inoperative (OEI) requirements, together with hot-gas reingestion and hot-day, high-altitude take-off situations, power augmentation for multiengine rotorcraft has always been of critical interest. However, power augmentation by using overtemperature at the turbine inlet will shorten turbine life unless a method of limiting thermal and mechanical stress is found. A possible solution involves allowing the turbine inlet temperature to rise to augment power while injecting water into the turbine cooling air to limit hot-section metal temperatures. An experimental water injection device was installed in an engine and successfully tested. Although concern for unprotected subcomponents in the engine hot section prevented demonstration of the technique's maximum potential, it was still possible to demonstrate increases in power while maintaining nearly constant turbine rotor blade temperature.

A very cool cooling system

CERN Multimedia

Antonella Del Rosso

2015-01-01

The NA62 Gigatracker is a jewel of technology: its sensor, which delivers the time of the crossing particles with a precision of less than 200 picoseconds (better than similar LHC detectors), has a cooling system that might become the precursor to a completely new detector technique.   The 115 metre long vacuum tank of the NA62 experiment. The NA62 Gigatracker (GTK) is composed of a set of three innovative silicon pixel detectors, whose job is to measure the arrival time and the position of the incoming beam particles. Installed in the heart of the NA62 detector, the silicon sensors are cooled down (to about -20 degrees Celsius) by a microfluidic silicon device. “The cooling system is needed to remove the heat produced by the readout chips the silicon sensor is bonded to,” explains Alessandro Mapelli, microsystems engineer working in the Physics department. “For the NA62 Gigatracker we have designed a cooling plate on top of which both the silicon sensor and the...

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)

Karl Poggensee - A widely unknown German rocket pioneer - The early years 1930-1934 - A chronology

Science.gov (United States)

Rohrwild, Karlheinz

2017-09-01

The rediscovered estate of Karl Poggensee allows to reproduce chronologically his rocket tests of the period 1930-1934 almost completely for the first time. Thrilled by the movie ;The Woman in the Moon; for the idea of space travel, he started as a student of Hinderburg-Polytechnikum (IAO), Oldenburg, to build his first solid-fuel rocket, producing his own propellant charges. Being a coming electrical engineer his main goal was not set up new record heights, but to provide his rockets with automatic measuring instruments, camera and parachute release systems. The optimization of this sequence was his main focus.

Rocket observations

Science.gov (United States)

1984-05-01

The Institute of Space and Astronautical Science (ISAS) sounding rocket experiments were carried out during the periods of August to September, 1982, January to February and August to September, 1983 and January to February, 1984 with sounding rockets. Among 9 rockets, 3 were K-9M, 1 was S-210, 3 were S-310 and 2 were S-520. Two scientific satellites were launched on February 20, 1983 for solar physics and on February 14, 1984 for X-ray astronomy. These satellites were named as TENMA and OHZORA and designated as 1983-011A and 1984-015A, respectively. Their initial orbital elements are also described. A payload recovery was successfully carried out by S-520-6 rocket as a part of MINIX (Microwave Ionosphere Non-linear Interaction Experiment) which is a scientific study of nonlinear plasma phenomena in conjunction with the environmental assessment study for the future SPS project. Near IR observation of the background sky shows a more intense flux than expected possibly coming from some extragalactic origin and this may be related to the evolution of the universe. US-Japan cooperative program of Tether Experiment was done on board US rocket.

Nonlinear Longitudinal Mode Instability in Liquid Propellant Rocket Engine Preburners

Science.gov (United States)

Sims, J. D. (Technical Monitor); Flandro, Gary A.; Majdalani, Joseph; Sims, Joseph D.

2004-01-01

Nonlinear pressure oscillations have been observed in liquid propellant rocket instability preburner devices. Unlike the familiar transverse mode instabilities that characterize primary combustion chambers, these oscillations appear as longitudinal gas motions with frequencies that are typical of the chamber axial acoustic modes. In several respects, the phenomenon is similar to longitudinal mode combustion instability appearing in low-smoke solid propellant motors. An important feature is evidence of steep-fronted wave motions with very high amplitude. Clearly, gas motions of this type threaten the mechanical integrity of associated engine components and create unacceptably high vibration levels. This paper focuses on development of the analytical tools needed to predict, diagnose, and correct instabilities of this type. For this purpose, mechanisms that lead to steep-fronted, high-amplitude pressure waves are described in detail. It is shown that such gas motions are the outcome of the natural steepening process in which initially low amplitude standing acoustic waves grow into shock-like disturbances. The energy source that promotes this behavior is a combination of unsteady combustion energy release and interactions with the quasi-steady mean chamber flow. Since shock waves characterize the gas motions, detonation-like mechanisms may well control the unsteady combustion processes. When the energy gains exceed the losses (represented mainly by nozzle and viscous damping), the waves can rapidly grow to a finite amplitude limit cycle. Analytical tools are described that allow the prediction of the limit cycle amplitude and show the dependence of this wave amplitude on the system geometry and other design parameters. This information can be used to guide corrective procedures that mitigate or eliminate the oscillations.

Numerical simulation of divergent rocket-based-combined-cycle performances under the flight condition of Mach 3

Science.gov (United States)

Cui, Peng; Xu, WanWu; Li, Qinglian

2018-01-01

Currently, the upper operating limit of the turbine engine is Mach 2+, and the lower limit of the dual-mode scramjet is Mach 4. Therefore no single power systems can operate within the range between Mach 2 + and Mach 4. By using ejector rockets, Rocket-based-combined-cycle can work well in the above scope. As the key component of Rocket-based-combined-cycle, the ejector rocket has significant influence on Rocket-based-combined-cycle performance. Research on the influence of rocket parameters on Rocket-based-combined-cycle in the speed range of Mach 2 + to Mach 4 is scarce. In the present study, influences of Mach number and total pressure of the ejector rocket on Rocket-based-combined-cycle were analyzed numerically. Due to the significant effects of the flight conditions and the Rocket-based-combined-cycle configuration on Rocket-based-combined-cycle performances, flight altitude, flight Mach number, and divergence ratio were also considered. The simulation results indicate that matching lower altitude with higher flight Mach numbers can increase Rocket-based-combined-cycle thrust. For another thing, with an increase of the divergent ratio, the effect of the divergent configuration will strengthen and there is a limit on the divergent ratio. When the divergent ratio is greater than the limit, the effect of divergent configuration will gradually exceed that of combustion on supersonic flows. Further increases in the divergent ratio will decrease Rocket-based-combined-cycle thrust.

Stennis engineer part of LCROSS moon mission

Science.gov (United States)

2009-01-01

Karma Snyder, a project manager at NASA's John C. Stennis Space Center, was a senior design engineer on the RL10 liquid rocket engine that powered the Centaur, the upper stage of the rocket used in NASA's Lunar CRater Observation and Sensing Satellite (LCROSS) mission in October 2009. Part of the LCROSS mission was to search for water on the moon by striking the lunar surface with a rocket stage, creating a plume of debris that could be analyzed for water ice and vapor. Snyder's work on the RL10 took place from 1995 to 2001 when she was a senior design engineer with Pratt & Whitney Rocketdyne. Years later, she sees the project as one of her biggest accomplishments in light of the LCROSS mission. 'It's wonderful to see it come into full service,' she said. 'As one of my co-workers said, the original dream was to get that engine to the moon, and we're finally realizing that dream.'

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

Unsteady response of flow system around balance piston in a rocket pump

Science.gov (United States)

Kawasaki, S.; Shimura, T.; Uchiumi, M.; Hayashi, M.; Matsui, J.

2013-03-01

In the rocket engine turbopump, a self-balancing type of axial thrust balancing system using a balance piston is often applied. In this study, the balancing system in liquid-hydrogen (LH2) rocket pump was modeled combining the mechanical structure and the flow system, and the unsteady response of the balance piston was investigated. The axial vibration characteristics of the balance piston with a large amplitude were determined, sweeping the frequency of the pressure fluctuation on the inlet of the balance piston. This vibration was significantly affected by the compressibility of LH2.

Actively Cooled Ceramic Composite Nozzle Material, Phase I

Data.gov (United States)

National Aeronautics and Space Administration — For Next Generation Launch Vehicles (NGLV), Either a Rocket-based or Turbine-based Combined Cycle (RBCC or TBCC) engine will power the Next Generation Launch Vehicle...

Investigation of the loss of forced cooling test by using the high temperature engineering test reactor (HTTR) (Contract research)

International Nuclear Information System (INIS)

Nakagawa, Shigeaki; Takamatsu, Kuniyoshi; Inaba, Yoshitomo; Goto, Minoru; Tochio, Daisuke

2007-09-01

The three gas circulators trip test and the vessel cooling system stop test as the safety demonstration test by using the High Temperature engineering Test Reactor (HTTR) are under planning to demonstrate inherent safety features of High Temperature Gas-cooled Reactor. All three gas circulators to circulate the helium gas as the coolant are stopped to simulate the loss of forced cooling in the three gas circulators trip test. The stop of the vessel cooling system located outside the reactor pressure vessel to remove the residual heat of the reactor core follows the stop of all three gas circulators in the vessel cooling system stop test. The analysis of the reactor transient for such tests and abnormal events postulated during the test was performed. From the result of analysis, it was confirmed that the three gas circulators trip test and the vessel cooling system stop test can be performed within the region of the normal operation in the HTTR and the safety of the reactor facility is ensured even if the abnormal events would occur. (author)

Transpiration and film cooling boundary layer computer program. Volume 1: Numerical solutions of the turbulent boundary layer equations with equilibrium chemistry

Science.gov (United States)

Levine, J. N.

1971-01-01

A finite difference turbulent boundary layer computer program has been developed. The program is primarily oriented towards the calculation of boundary layer performance losses in rocket engines; however, the solution is general, and has much broader applicability. The effects of transpiration and film cooling as well as the effect of equilibrium chemical reactions (currently restricted to the H2-O2 system) can be calculated. The turbulent transport terms are evaluated using the phenomenological mixing length - eddy viscosity concept. The equations of motion are solved using the Crank-Nicolson implicit finite difference technique. The analysis and computer program have been checked out by solving a series of both laminar and turbulent test cases and comparing the results to data or other solutions. These comparisons have shown that the program is capable of producing very satisfactory results for a wide range of flows. Further refinements to the analysis and program, especially as applied to film cooling solutions, would be aided by the acquisition of a firm data base.

A direct numerical simulation of cool-flame affected autoignition in diesel engine-relevant conditions

Energy Technology Data Exchange (ETDEWEB)

Krisman, Alexander; Hawkes, Evatt Robert.; Talei, Mohsen; Bhagatwala, Ankit; Chen, Jacqueline H.

2016-11-11

In diesel engines, combustion is initiated by a two-staged autoignition that includes both low- and high-temperature chemistry. The location and timing of both stages of autoignition are important parameters that influence the development and stabilisation of the flame. In this study, a two-dimensional direct numerical simulation (DNS) is conducted to provide a fully resolved description of ignition at diesel engine-relevant conditions. The DNS is performed at a pressure of 40 atmospheres and at an ambient temperature of 900 K using dimethyl ether (DME) as the fuel, with a 30 species reduced chemical mechanism. At these conditions, similar to diesel fuel, DME exhibits two-stage ignition. The focus of this study is on the behaviour of the low-temperature chemistry (LTC) and the way in which it influences the high-temperature ignition. The results show that the LTC develops as a “spotty” first-stage autoignition in lean regions which transitions to a diffusively supported cool-flame and then propagates up the local mixture fraction gradient towards richer regions. The cool-flame speed is much faster than can be attributed to spatial gradients in first-stage ignition delay time in homogeneous reactors. The cool-flame causes a shortening of the second-stage ignition delay times compared to a homogeneous reactor and the shortening becomes more pronounced at richer mixtures. Multiple high-temperature ignition kernels are observed over a range of rich mixtures that are much richer than the homogeneous most reactive mixture and most kernels form much earlier than suggested by the homogeneous ignition delay time of the corresponding local mixture. Altogether, the results suggest that LTC can strongly influence both the timing and location in composition space of the high-temperature ignition.

Two stage turbine for rockets

Science.gov (United States)

Veres, Joseph P.

1993-01-01

The aerodynamic design and rig test evaluation of a small counter-rotating turbine system is described. The advanced turbine airfoils were designed and tested by Pratt & Whitney. The technology represented by this turbine is being developed for a turbopump to be used in an advanced upper stage rocket engine. The advanced engine will use a hydrogen expander cycle and achieve high performance through efficient combustion of hydrogen/oxygen propellants, 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 mass flow rates and high operating pressures result in very small airfoil heights and diameters. The high efficiency and small size requirements present a challenging turbine design problem. The shrouded axial turbine blades are 50 percent reaction with a maximum thickness to chord ratio near 1. At 6 deg from the tangential direction, the nozzle and blade exit flow angles are well below the traditional design minimum limits. The blade turning angle of 160 deg also exceeds the maximum limits used in traditional turbine designs.

Three Dimensional Numerical Simulation of Rocket-based Combined-cycle Engine Response During Mode Transition Events

Science.gov (United States)

Edwards, Jack R.; McRae, D. Scott; Bond, Ryan B.; Steffan, Christopher (Technical Monitor)

2003-01-01

The GTX program at NASA Glenn Research Center is designed to develop a launch vehicle concept based on rocket-based combined-cycle (RBCC) propulsion. Experimental testing, cycle analysis, and computational fluid dynamics modeling have all demonstrated the viability of the GTX concept, yet significant technical issues and challenges still remain. Our research effort develops a unique capability for dynamic CFD simulation of complete high-speed propulsion devices and focuses this technology toward analysis of the GTX response during critical mode transition events. Our principal attention is focused on Mode 1/Mode 2 operation, in which initial rocket propulsion is transitioned into thermal-throat ramjet propulsion. A critical element of the GTX concept is the use of an Independent Ramjet Stream (IRS) cycle to provide propulsion at Mach numbers less than 3. In the IRS cycle, rocket thrust is initially used for primary power, and the hot rocket plume is used as a flame-holding mechanism for hydrogen fuel injected into the secondary air stream. A critical aspect is the establishment of a thermal throat in the secondary stream through the combination of area reduction effects and combustion-induced heat release. This is a necessity to enable the power-down of the rocket and the eventual shift to ramjet mode. Our focus in this first year of the grant has been in three areas, each progressing directly toward the key initial goal of simulating thermal throat formation during the IRS cycle: CFD algorithm development; simulation of Mode 1 experiments conducted at Glenn's Rig 1 facility; and IRS cycle simulations. The remainder of this report discusses each of these efforts in detail and presents a plan of work for the next year.

The comparative analysis of the forecasts of development of rocket propulsion in past and now

Science.gov (United States)

Nedaivoda, A.; Prisniakov, V.

2001-03-01

Consideration is being given to use the known long and short forecasts of development of rocket engines in past - at the beginning of development of a missile engineering (K. Tsiolkovsky etc. pioneers of rocket propulsion); on the eve of launching of the artificial satellite of Earth (A. Blagonravov); after manned flight of Yu. Gagarin (V. Gluchko); after manned flight on Moon (" The Forecasts on 2001 " on materials of readings R. Goddard in USA); in middle of 70-s' years (D. Sevruk, V. Prisniakov) and at the end of 20 centure. Last years under the initiative R. Beichel and M. Pouliquen IAA. Advanced Propulsion Working Group carries out large researches on definition of the tendencies of development of rocket propulsion for the next forty years, the outcomes which one will be used in the report. The comparison of development of rocket propulsion expected to the end of 20 century and real-life is given. The report analyses the errors of the forecasts of the past - the absence reliable prognostic procedure; the euphoria of the maiden successes of conquest of space; dominance of military and political- propaganda motives of implementation of the space programs before economical; to keep developments secret; competition of two super-powers USSR and USA etc.

Modal Survey of ETM-3, A 5-Segment Derivative of the Space Shuttle Solid Rocket Booster

Science.gov (United States)

Nielsen, D.; Townsend, J.; Kappus, K.; Driskill, T.; Torres, I.; Parks, R.

2005-01-01

The complex interactions between internal motor generated pressure oscillations and motor structural vibration modes associated with the static test configuration of a Reusable Solid Rocket Motor have potential to generate significant dynamic thrust loads in the 5-segment configuration (Engineering Test Motor 3). Finite element model load predictions for worst-case conditions were generated based on extrapolation of a previously correlated 4-segment motor model. A modal survey was performed on the largest rocket motor to date, Engineering Test Motor #3 (ETM-3), to provide data for finite element model correlation and validation of model generated design loads. The modal survey preparation included pretest analyses to determine an efficient analysis set selection using the Effective Independence Method and test simulations to assure critical test stand component loads did not exceed design limits. Historical Reusable Solid Rocket Motor modal testing, ETM-3 test analysis model development and pre-test loads analyses, as well as test execution, and a comparison of results to pre-test predictions are discussed.

On fundamentally new sources of energy for rockets in the early works of the pioneers of astronautics

Science.gov (United States)

Melkumov, T. M.

1977-01-01

The research for more efficient methods of propelling a spacecraft, than can be achieved with chemical energy, was studied. During a time when rockets for space flight had not actually been built pioneers in rocket technology were already concerned with this problem. Alternative sources proposed at that time, were nuclear and solar energy. Basic engineering problems of each source were investigated.

Solid Rocket Testing at AFRL (Briefing Charts)

Science.gov (United States)

2016-10-21

Distribution Unlimited. PA#16492 2 Agenda • Solid Rocket Motors • History of Sea Level Testing • Small Component Testing • Full-scale Testing • Altitude...Facility • History of Testing • Questions -Distribution A: Approved for Public Release; Distribution Unlimited. PA#16492 3 RQ-West • AFRL/RQ...INTEGRATION FACILITY NATIONAL HOVER TEST FACILITY TITAN SRM TEST FACILITY TS-1C1-125 LARGE ENGINE/COMPONENT TEST FACILITY TS-1A 1-120 1-115 X-33 LAUNCH

Cooling systems for waste heat. Cooling systems, review and selection criteria. Kuehlsysteme fuer Abwaerme. Kuehlsysteme, Ueberblick und Auswahlkriterien

Energy Technology Data Exchange (ETDEWEB)

Mueller, W. (Jaeggi, Wallisellen (Switzerland))

1990-05-01

In many areas of ventilation, air-conditioning and refrigeration engineering, chemical and process engineering and energy production waste heat occurs. If a reduction in energy losses or heat recovery is not possible waste heat has to be drawn off through cooling systems. For this the following systems can be used: dry cooling systems, dry cooler with spray system, open-cycle wet cooler, hybrid dry cooler, and closed-cycle wet cooler. Particularly hybrid cooling systems can give acceptable solutions when the results with other systems are only unsatisfactory. (BWI).

Effects of Engine Cooling Water Temperature on Performance and Emission Characteristics of a Ci Engine Operated with Biofuel Blend

Directory of Open Access Journals (Sweden)

Abul Hossain

2017-03-01

Full Text Available The temperature of the coolant is known to have significant influence on engine performance and emissions. Whereas existing literature describes the effects of coolant temperature in engines using fossil derived fuels, very few studies have investigated these effects when biofuel is used. In this study, Jatropha oil was blended separately with ethanol and butanol. It was found that the 80% jatropha oil + 20% butanol blend was the most suitable alternative, as its properties were closest to that of fossil diesel. The coolant temperature was varied between 50°C and 95°C. The combustion process enhanced for both diesel and biofuel blend, when the coolant temperature was increased. The carbon dioxide emissions for both diesel and biofuel blend were observed to increase with temperature. The carbon monoxide, oxygen and lambda values were observed to decrease with temperature. When the engine was operated using diesel, nitrogen oxides emissions correlated in an opposite manner to smoke opacity; however, nitrogen oxides emissions and smoke opacity correlated in an identical manner for biofuel blend. Brake specific fuel consumption was observed to decrease as the temperature was increased and was higher on average when the biofuel was used. The study concludes that both biofuel blend and fossil diesel produced identical correlations between coolant temperature and engine performance. The trends of nitrogen oxides and smoke emissions with cooling temperatures were not identical to fossil diesel when biofuel blend was used in the engine.

Bibliography of Books and Published Reports on Gas Turbines, Jet Propulsion, and Rocket Power Plants

Science.gov (United States)

1951-06-01

Ink , New York, 1945. W. Ley, Rockets. Viking Press. New York. 1945. LI. S. Zim, Rockets and jets. Harcourt Brace, New York, 1945. Jet propulsion...Hausenblas, Design nomograms for turbine stages. Motortechnische Zeit. 11, 96 (Aug. 1950). S. L. Koutz et al., Effect of beat and power extraction on...Edelman, The pulsating engine-its evolution and future prospects. SAE Quart. Trans. 1, 204 (1947). R. McLarren, Project Squid probes pulsejet. Aviation

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

International Nuclear Information System (INIS)

O'Brien, Robert C.

2001-01-01

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.

Multidisciplinary design of a rocket-based combined cycle SSTO launch vehicle using Taguchi methods

Science.gov (United States)

Olds, John R.; Walberg, Gerald D.

1993-01-01

Results are presented from the optimization process of a winged-cone configuration SSTO launch vehicle that employs a rocket-based ejector/ramjet/scramjet/rocket operational mode variable-cycle engine. The Taguchi multidisciplinary parametric-design method was used to evaluate the effects of simultaneously changing a total of eight design variables, rather than changing them one at a time as in conventional tradeoff studies. A combination of design variables was in this way identified which yields very attractive vehicle dry and gross weights.

Eddie Rocket's Franchise

OpenAIRE

Vahter, Jenni

2008-01-01

Eddie Rocket's Franchise - Setting up a franchise restaurant in Helsinki. TIIVISTELMÄ: Eddie Rocket's on menestynyt amerikkalaistyylinen 1950-luvun ”diner” franchiseravintolaketju Irlannista. Ravintoloita on perustettu viimeisen 18 vuoden aikana 28 kappaletta Irlantiin ja Isoon Britanniaan sekä yksi Espanjaan. Tämän tutkimuksen tarkoitus on tutkia onko Eddie Rocket'silla potentiaalia menestyä Helsingissä, Suomessa. Tutkimuskysymystä on lähestytty toimiala-analyysin, markkinatutkimuksen j...

State Machine Modeling of the Space Launch System Solid Rocket Boosters

Science.gov (United States)

Harris, Joshua A.; Patterson-Hine, Ann

2013-01-01

The Space Launch System is a Shuttle-derived heavy-lift vehicle currently in development to serve as NASA's premiere launch vehicle for space exploration. The Space Launch System is a multistage rocket with two Solid Rocket Boosters and multiple payloads, including the Multi-Purpose Crew Vehicle. Planned Space Launch System destinations include near-Earth asteroids, the Moon, Mars, and Lagrange points. The Space Launch System is a complex system with many subsystems, requiring considerable systems engineering and integration. To this end, state machine analysis offers a method to support engineering and operational e orts, identify and avert undesirable or potentially hazardous system states, and evaluate system requirements. Finite State Machines model a system as a finite number of states, with transitions between states controlled by state-based and event-based logic. State machines are a useful tool for understanding complex system behaviors and evaluating "what-if" scenarios. This work contributes to a state machine model of the Space Launch System developed at NASA Ames Research Center. The Space Launch System Solid Rocket Booster avionics and ignition subsystems are modeled using MATLAB/Stateflow software. This model is integrated into a larger model of Space Launch System avionics used for verification and validation of Space Launch System operating procedures and design requirements. This includes testing both nominal and o -nominal system states and command sequences.

Modeling of Nonlinear Marine Cooling Systems with Closed Circuit Flow

DEFF Research Database (Denmark)

Hansen, Michael; Stoustrup, Jakob; Bendtsen, Jan Dimon

2011-01-01

We consider the problem of constructing a mathematical model for a specific type of marine cooling system. The system in question is used for cooling the main engine and main engine auxiliary components, such as diesel generators, turbo chargers and main engine air coolers for certain classes...

Metallic hydrogen: The most powerful rocket fuel yet to exist

Energy Technology Data Exchange (ETDEWEB)

Silvera, Isaac F [Lyman Laboratory of Physics, Harvard University, Cambridge MA 02138 (United States); Cole, John W, E-mail: silvera@physics.harvard.ed [NASA MSFC, Huntsville, AL 35801 (United States)

2010-03-01

Wigner and Huntington first predicted that pressures of order 25 GPa were required for the transition of solid molecular hydrogen to the atomic metallic phase. Later it was predicted that metallic hydrogen might be a metastable material so that it remains metallic when pressure is released. Experimental pressures achieved on hydrogen have been more than an order of magnitude higher than the predicted transition pressure and yet it remains an insulator. We discuss the applications of metastable metallic hydrogen to rocketry. Metastable metallic hydrogen would be a very light-weight, low volume, powerful rocket propellant. One of the characteristics of a propellant is its specific impulse, I{sub sp}. Liquid (molecular) hydrogen-oxygen used in modern rockets has an Isp of {approx}460s; metallic hydrogen has a theoretical I{sub sp} of 1700s. Detailed analysis shows that such a fuel would allow single-stage rockets to enter into orbit or carry economical payloads to the moon. If pure metallic hydrogen is used as a propellant, the reaction chamber temperature is calculated to be greater than 6000 K, too high for currently known rocket engine materials. By diluting metallic hydrogen with liquid hydrogen or water, the reaction temperature can be reduced, yet there is still a significant performance improvement for the diluted mixture.

Nuclear thermal rockets using indigenous Martian propellants

International Nuclear Information System (INIS)

Zubrin, R.M.

1989-01-01

This paper considers a novel concept for a Martian descent and ascent vehicle, called NIMF (for nuclear rocket using indigenous Martian fuel), the propulsion for which will be provided by a nuclear thermal reactor which will heat an indigenous Martian propellant gas to form a high-thrust rocket exhaust. The performance of each of the candidate Martian propellants, which include CO2, H2O, CH4, N2, CO, and Ar, is assessed, and the methods of propellant acquisition are examined. Attention is also given to the issues of chemical compatibility between candidate propellants and reactor fuel and cladding materials, and the potential of winged Mars supersonic aircraft driven by this type of engine. It is shown that, by utilizing the nuclear landing craft in combination with a hydrogen-fueled nuclear thermal interplanetary vehicle and a heavy lift booster, it is possible to achieve a manned Mars mission in one launch. 6 refs

How Does Rocket Propulsion Work? The most common answer to ...

Indian Academy of Sciences (India)

internal combustion engines. The fuel or propellant is stored in the fuel tank. Here we will consider liquid hydrogen as the fuel. For the combustion to take place in outer space or in the absence of atmospheric oxygen the rocket carries along an oxidizer; here we will consider liquid oxygen as the oxidizer. The oxidizer or in.

Two-Rockets Thought Experiment

Science.gov (United States)

Smarandache, Florentin

2014-03-01

Let n>=2 be identical rockets: R1 ,R2 , ..., Rn. Each of them moving at constant different velocities respectively v1, v2, ..., vn on parallel directions in the same sense. In each rocket there is a light clock, the observer on earth also has a light clock. All n + 1 light clocks are identical and synchronized. The proper time Δt' in each rocket is the same. Let's focus on two arbitrary rockets Ri and Rjfrom the previous n rockets. Let's suppose, without loss of generality, that their speeds verify virocket Rj is contracted with the factor C(vj -vi) , i.e. Lj =Lj' C(vj -vi) .(2) But in the reference frame of the astronaut in Rjit is like rocket Rjis stationary andRi moves with the speed vj -vi in opposite direction. Therefore, similarly, the non-proper time interval as measured by the astronaut inRj with respect to the event inRi is dilated with the same factor D(vj -vi) , i.e. Δtj . i = Δt' D(vj -vi) , and rocketRi is contracted with the factor C(vj -vi) , i.e. Li =Li' C(vj -vi) .But it is a contradiction to have time dilations in both rockets. (3) Varying i, j in {1, 2, ..., n} in this Thought Experiment we get again other multiple contradictions about time dilations. Similarly about length contractions, because we get for a rocket Rj, n-2 different length contraction factors: C(vj -v1) , C(vj -v2) , ..., C(vj -vj - 1) , C(vj -vj + 1) , ..., C(vj -vn) simultaneously! Which is abnormal.

The Development of Rocketry Capability in New Zealand—World Record Rocket and First of Its Kind Rocketry Course

Directory of Open Access Journals (Sweden)

George Buchanan

2015-02-01

Full Text Available The University of Canterbury has developed a rocket research group, UC Rocketry, which recently broke the world altitude record for an I-class motor (impulse of 320–640 Ns and has run a rocketry course for the first time in New Zealand. This paper discusses the development and results of the world record rocket “Milly” and details all the fundamental elements of the rocketry final year engineering course, including the manufacturing processes, wind tunnel testing, avionics, control and the final rocket launch of “Smokey”. The rockets Milly and Smokey are an example of the design, implementation and testing methodologies that have significantly contributed to research and graduates for New Zealand’s space program.

Rocketing into the future the history and technology of rocket planes

CERN Document Server

van Pelt, Michel

2012-01-01

Rocketing into the Future journeys into the exciting world of rocket planes, examining the exotic concepts and actual flying vehicles that have been devised over the last one hundred years. Lavishly illustrated with over 150 photographs, it recounts the history of rocket planes from the early pioneers who attached simple rockets on to their wooden glider airplanes to the modern world of high-tech research vehicles. The book then looks at the possibilities for the future. The technological and economic challenges of the Space Shuttle proved insurmountable, and thus the program was unable to fulfill its promise of low-cost access to space. However, the burgeoning market of suborbital space tourism may yet give the necessary boost to the development of a truly reusable spaceplane.

An example of successful international cooperation in rocket motor technology

Science.gov (United States)

Ellis, Russell A.; Berdoyes, Michel

2002-07-01

The history of over 25 years of cooperation between Pratt & Whitney, San Jose, CA, USA and Snecma Moteurs, Le Haillan, France in solid rocket motor and, in one case, liquid rocket engine technology is presented. Cooperative efforts resulted in achievements that likely would not have been realized individually. The combination of resources and technologies resulted in synergistic benefits and advancement of the state of the art in rocket motors and components. Discussions begun between the two companies in the early 1970's led to the first cooperative project, demonstration of an advanced apogee motor nozzle, during the mid 1970's. Shortly thereafter advanced carboncarbon (CC) throat materials from Snecma were comparatively tested with other materials in a P&W program funded by the USAF. Use of Snecma throat materials in CSD Tomahawk boosters followed. Advanced space motors were jointly demonstrated in company-funded joint programs in the late 1970's and early 1980's: an advanced space motor with an extendible exit cone and an all-composite advanced space motor that included a composite chamber polar adapter. Eight integral-throat entrances (ITEs) of 4D and 6D construction were tested by P&W for Snecma in 1982. Other joint programs in the 1980's included test firing of a "membrane" CC exit cone, and integral throat and exit cone (ITEC) nozzle incorporating NOVOLTEX® SEPCARB® material. A variation of this same material was demonstrated as a chamber aft polar boss in motor firings that included demonstration of composite material hot gas valve thrust vector control (TVC). In the 1990's a supersonic splitline flexseal nozzle was successfully demonstrated by the two companies as part of a US Integrated High Payoff Rocket Propulsion Technology (IHPRPT) program effort. Also in the mid-1990s the NOVOLTEX® SEPCARB® material, so successful in solid rocket motor application, was successfully applied to a liquid engine nozzle extension. The first cooperative

Lunar mission design using nuclear thermal rockets

International Nuclear Information System (INIS)

Stancati, M.L.; Collins, J.T.; Borowski, S.K.

1991-01-01

The NERVA-class Nuclear Thermal Rocket (NTR), with performance nearly double that of advanced chemical engines, has long been considered an enabling technology for human missions to Mars. NTR engines address the demanding trip time and payload delivery needs of both cargo-only and piloted flights. But NTR can also reduce the Earth launch requirements for manned lunar missions. First use of NTR for the Moon would be less demanding and would provide a test-bed for early operations experience with this powerful technology. Study of application and design options indicates that NTR propulsion can be integrated with the Space Exploration Initiative scenarios to deliver performance gains while managing controlled, long-term disposal of spent reactors to highly stable orbits

Preliminary Evaluation of a Turbine/Rotary Combustion Compound Engine for a Subsonic Transport. [fuel consumption and engine tests of turbofan engines

Science.gov (United States)

Civinskas, K. C.; Kraft, G. A.

1976-01-01

The fuel consumption of a modern compound engine with that of an advanced high pressure ratio turbofan was compared. The compound engine was derived from a turbofan engine by replacing the combustor with a rotary combustion (RC) engine. A number of boost pressure ratios and compression ratios were examined. Cooling of the RC engine was accomplished by heat exchanging to the fan duct. Performance was estimated with an Otto-cycle for two levels of energy lost to cooling. The effects of added complexity on cost and maintainability were not examined and the comparison was solely in terms of cruise performance and weight. Assuming a 25 percent Otto-cycle cooling loss (representative of current experience), the best compound engine gave a 1.2 percent improvement in cruise. Engine weight increased by 23 percent. For a 10 percent Otto-cycle cooling loss (representing advanced insulation/high temperature materials technology), a compound engine with a boost PR of 10 and a compression ratio of 10 gave an 8.1 percent lower cruise than the reference turbofan.

Temperature control system with computer mapping for engine cooling circuits; Kennfeldgesteuertes Temperaturregelsystem fuer Motorkuehlkreislaeufe

Energy Technology Data Exchange (ETDEWEB)

Saur, R.; Leu, P.; Lemberger, H.; Heumer, G.

1996-07-01

Thermomanagement of the vehicles powered by internal combustion engines is one of the prerequisites needed to fulfil the German automobil industry`s commitment to reduce fuel consumption by 25% as compared with 1990 before the year 2005. Thermomanagement improves comfort, and reduces fuel consumption and toxic emissions. BMW and Behr Thermot-Tronik have jointly developed the first component of such a thermomanagement system: An engine cooling system with computer mapping. BMW is the first manufacturer worldwide to install this system as standard equipment, as it is doing in its refinde eight-cylinder engine series (M62). (orig.) [Deutsch] Die Verpflichtung der deutschen Automobilindustrie, den Kraftstoffverbrauch bis zum Jahre 2005 zur Basis des Jahres 1990 um 25% zu reduzieren, fuehrt unter anderem zwingend zum Thermomanagement der mit Verbrennungsmotoren betriebenen Fahrzeuge. Ein Thermomanagement verbessert den Komfort, reduziert den Verbrauch und vermindert die Schadstoffemissionen. BMW und Behr Thermot-Tronik haben gemeinsam den ersten Baustein des Thermomanagements - das kennfeldgesteuerte Motorkuehlungssystem - entwickelt. Dieses System wird seit Januar 1996 weltweit erstmalig von BMW serienmaessig in der ueberarbeiteten Achtzylinder-Motorbaureihe (M62) eingesetzt. (orig.)

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.

Nuclear-powered rocket of the future

Energy Technology Data Exchange (ETDEWEB)

Yunqiao, B

1979-06-01

A possible manned mission to Mars with a crew of 7 in an 80-meter-long nuclear-powered rocket will take 180 days to reach its destination, will spend 10 to 14 days on the surface, and will take 200 days to return. A nuclear-powered engine (using U-235 or U-239) is the most likely means of propulsion. Four designs are described. The superheated exhaust engine will use a reactor to heat liquid hydrogen to over 4000/sup 0/C, after which it will be ejected from the exhaust. A plasma compression engine will use electric current produced by a reactor to heat hydrogen to plasma temperature (70,000/sup 0/C), after which it will be ejected through the exhaust by a magnetic field. In a gaseous-core reactor engine, gaseous fuel will heat liquid hydrogen to over 9,000/sup 0/C and use it as the propellant. The boldest solution is a proposal to use small nuclear explosions as the propulsive force. The first alternative will probably not produce enough thrust, while there will be a difficulty producing sufficient electricity in the second alternative. The other two alternatives seem promising.

Technology Development of a Fiber Optic-Coupled Laser Ignition System for Multi-Combustor Rocket Engines

Science.gov (United States)

Trinh, Huu P.; Early, Jim; Osborne, Robin; Thomas, Matthew E.; Bossard, John A.

2002-01-01

This paper addresses the progress of technology development of a laser ignition system at NASA Marshall Space Flight Center (MSFC). The first two years of the project focus on comprehensive assessments and evaluations of a novel dual-pulse laser concept, flight- qualified laser system, and the technology required to integrate the laser ignition system to a rocket chamber. With collaborations of the Department of Energy/Los Alamos National Laboratory (LANL) and CFD Research Corporation (CFDRC), MSFC has conducted 26 hot fire ignition tests with lab-scale laser systems. These tests demonstrate the concept feasibility of dual-pulse laser ignition to initiate gaseous oxygen (GOX)/liquid kerosene (RP-1) combustion in a rocket chamber. Presently, a fiber optic- coupled miniaturized laser ignition prototype is being implemented at the rocket chamber test rig for future testing. Future work is guided by a technology road map that outlines the work required for maturing a laser ignition system. This road map defines activities for the next six years, with the goal of developing a flight-ready laser ignition system.

A study of air breathing rockets. 3: Supersonic mode combustors

Science.gov (United States)

Masuya, G.; Chinzel, N.; Kudo, K.; Murakami, A.; Komuro, T.; Ishii, S.

An experimental study was made on supersonic mode combustors of an air breathing rocket engine. Supersonic streams of room-temperature air and hot fuel-rich rocket exhaust were coaxially mixed and burned in a concially diverging duct of 2 deg half-angle. The effect of air inlet Mach number and excess air ratio was investigated. Axial wall pressure distribution was measured to calculate one dimensional change of Mach number and stagnation temperature. Calculated results showed that supersonic combustion occurred in the duct. At the exit of the duct, gas sampling and Pitot pressure measurement was made, from which radial distributions of various properties were deduced. The distribution of mass fraction of elements from rocket exhaust showed poor mixing performance in the supersonic mode combustors compared with the previously investigated cylindrical subsonic mode combustors. Secondary combustion efficiency correlated well with the centerline mixing parameter, but not with Annushkin's non-dimensional combustor length. No major effect of air inlet Mach number or excess air ratio was seen within the range of conditions under which the experiment was conducted.

Combining MHD Airbreathing and Fusion Rocket Propulsion for Earth-to-Orbit Flight

International Nuclear Information System (INIS)

Froning, H. D. Jr; Yang, Yang; Momota, H.; Burton, E.; Miley, G. H.; Luo, Nie

2005-01-01

Previous studies have shown that Single-State-to-Orbit (SSTO) vehicle propellant can be reduced by Magnets-Hydro-Dynamic (MHD) processes that minimize airbreathing propulsion losses and propellant consumption during atmospheric flight. Similarly additional reduction in SSTO propellant is enabled by Inertial Electrostatic Confinement (IEC) fusion, whose more energetic reactions reduce rocket propellant needs. MHD airbreathing propulsion during an SSTO vehicle's initial atmospheric flight phase and IEC fusion propulsion during its final exo-atmospheric flight phase is therefore being explored. Accomplished work is not yet sufficient for claiming such a vehicle's feasibility. But takeoff and propellant mass for an MHD airbreathing and IEC fusion vehicle could be as much as 25 and 40 percent less than one with ordinary airbreathing and IEC fusion; and as much as 50 and 70 percent less than SSTO takeoff and propellant mass with MHD airbreathing and chemical rocket propulsion. Thus this unusual combined cycle engine shows great promise for performance gains beyond contemporary combined-cycle airbreathing engines

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.

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…

Heated Tube Facility

Data.gov (United States)

Federal Laboratory Consortium — The Heated Tube Facility at NASA GRC investigates cooling issues by simulating conditions characteristic of rocket engine thrust chambers and high speed airbreathing...

Ablative Material Testing at Lewis Rocket Lab

Science.gov (United States)

1997-01-01

The increasing demand for a low-cost, reliable way to launch commercial payloads to low- Earth orbit has led to the need for inexpensive, expendable propulsion systems for new launch vehicles. This, in turn, has renewed interest in less complex, uncooled rocket engines that have combustion chambers and exhaust nozzles fabricated from ablative materials. A number of aerospace propulsion system manufacturers have utilized NASA Lewis Research Center's test facilities with a high degree of success to evaluate candidate materials for application to new propulsion devices.

South Pole rockets, (1)

International Nuclear Information System (INIS)

Kimura, Iwane

1977-01-01

Wave-particle interaction was observed, using three rockets, S-210 JA-20, -21 and S-310 JA-2, launched from the South Pole into aurora. Electron density and temperature were measured with these rockets. Simultaneous observations of waves were also made from a satellite (ISIS-II) and at two ground bases (Showa base and Mizuho base). Observed data are presented in this paper. These include electron density and temperature in relation to altitude; variation of electron (60 - 80 keV) count rate with altitude; VLF spectra measured by the PWL of S-210 JA-20 and -21 rockets and the corresponding VLF spectra at the ground bases; low-energy (<10 keV) electron flux measured by S-310 JA-2 rocket; and VLF spectrum measured with S-310 JA-2 rocket. Scheduled measurements for the next project are also briefly described. (Aoki, K.)

Development of Engine Loads Methodology, Phase I

Data.gov (United States)

National Aeronautics and Space Administration — This SBIR seeks to improve the definition of design loads for rocket engine components such that higher performing, lighter weight engines can be developed more...

Micro-Rockets for the Classroom.

Science.gov (United States)

Huebner, Jay S.; Fletcher, Alice S.; Cato, Julia A.; Barrett, Jennifer A.

1999-01-01

Compares micro-rockets to commercial models and water rockets. Finds that micro-rockets are more advantageous because they are constructed with inexpensive and readily available materials and can be safely launched indoors. (CCM)

Hot rocket plume experiment - Survey and conceptual design. [of rhenium-iridium bipropellants

Science.gov (United States)

Millard, Jerry M.; Luan, Taylor W.; Dowdy, Mack W.

1992-01-01

Attention is given to a space-borne engine plume experiment study to fly an experiment which will both verify and quantify the reduced contamination from advanced rhenium-iridium earth-storable bipropellant rockets (hot rockets) and provide a correlation between high-fidelity, in-space measurements and theoretical plume and surface contamination models. The experiment conceptual design is based on survey results from plume and contamination technologists throughout the U.S. With respect to shuttle use, cursory investigations validate Hitchhiker availability and adaptability, adequate remote manipulator system (RMS) articulation and dynamic capability, acceptable RMS attachment capability, adequate power and telemetry capability, and adequate flight altitude and attitude/orbital capability.

Blade Surface Pressure Distributions in a Rocket Engine Turbine: Experimental Work With On-Blade Pressure Transducers

Science.gov (United States)

Hudson, Susan T.; Zoladz, Thomas F.; Griffin, Lisa W.; Turner, James E. (Technical Monitor)

2000-01-01

Understanding the unsteady aspects of turbine rotor flowfields is critical to successful future turbine designs. A technology program was conducted at NASA's Marshall Space Flight Center to increase the understanding of unsteady environments for rocket engine turbines. The experimental program involved instrumenting turbine rotor blades with surface-mounted high frequency response pressure transducers. The turbine model was then tested to measure the unsteady pressures on the rotor blades. The data obtained from the experimental program is unique in three respects. First, much more unsteady data was obtained (several minutes per set point) than has been possible in the past. Also, two independent unsteady data acquisition systems and fundamental signal processing approaches were used. Finally, an extensive steady performance database existed for the turbine model. This allowed an evaluation of the effect of the on-blade instrumentation on the turbine's performance. This unique data set, the lessons learned for acquiring this type of data, and the improvements made to the data analysis and prediction tools will contribute to future turbine programs such as those for reusable launch vehicles.

Study of Scramjet Engine System

OpenAIRE

苅田, 丈士; KANDA, Takeshi

2001-01-01

1. Introduction The scramjet engine for the single-stage-to-orbit (SSTO) aerospace plane has been studied in the ramjet propulsion research division. The problems of the scramjet are (1) combustion, (2) light structure, (3) startability of the inlet, (4) integration of engines, and (5) cooling. The construction of the cooling system is important for the scramjet engine, because of high heat flux during operation. Cooling is not only a problem for the engine itself, but also for the airframe. ...

T-Minus engineering : A company in rocketry

NARCIS (Netherlands)

Olthof, H.; Uitendaal, M.

2013-01-01

The idea started when a group of students from Delft Aerospace Rocket Engineering were sitting at the breakfast table of the Esrange launch base canteen, in the far north of Sweden. It was the day after the successful launch of the self-built Stratos rocket, that broke the European altitude record

Helical muon beam cooling channel engineering design

International Nuclear Information System (INIS)

Johnson, Rolland

2015-01-01

The Helical Cooling Channel (HCC) achieves effective ionization cooling of the six-dimensional (6d) phase space of a muon beam by means of a series of 21st century inventions. In the HCC, hydrogen-pressurized RF cavities enable high RF gradients in strong external magnetic fields. The theory of the HCC, which requires a magnetic field with solenoid, helical dipole, and helical quadrupole components, demonstrates that dispersion in the gaseous hydrogen energy absorber provides effective emittance exchange to enable longitudinal ionization cooling. The 10-year development of a practical implementation of a muon-beam cooling device has involved a series of technical innovations and experiments that imply that an HCC of less than 300 m length can cool the 6d emittance of a muon beam by six orders of magnitude. We describe the design and construction plans for a prototype HCC module based on oxygen-doped hydrogen-pressurized RF cavities that are loaded with dielectric, fed by magnetrons, and operate in a superconducting helical solenoid magnet. The first phase of this project saw the development of a conceptual design for the integration of 805 MHz RF cavities into a 10 T Nb 3 Sn-based HS test section. Two very novel ideas are required to realize the design. The first idea is the use of dielectric inserts in the RF cavities to make them smaller for a given frequency so that the cavities and associated plumbing easily fit inside the magnet cryostat. Calculations indicate that heat loads will be tolerable, while RF breakdown of the dielectric inserts will be suppressed by the pressurized hydrogen gas. The second new idea is the use of a multi-layer Nb 3 Sn helical solenoid. The technology demonstrations for the two aforementioned key components of a 10T, 805 MHz HCC were begun in this project. The work load in the Fermilab Technical Division made it difficult to test a multi-layer Nb 3 Sn solenoid as originally planned. Instead, a complementary project was approved by the

Helical muon beam cooling channel engineering design

Energy Technology Data Exchange (ETDEWEB)

Johnson, Rolland [Muons, Inc., Batavia, IL (United States)

2015-08-07

The Helical Cooling Channel (HCC) achieves effective ionization cooling of the six-dimensional (6d) phase space of a muon beam by means of a series of 21st century inventions. In the HCC, hydrogen-pressurized RF cavities enable high RF gradients in strong external magnetic fields. The theory of the HCC, which requires a magnetic field with solenoid, helical dipole, and helical quadrupole components, demonstrates that dispersion in the gaseous hydrogen energy absorber provides effective emittance exchange to enable longitudinal ionization cooling. The 10-year development of a practical implementation of a muon-beam cooling device has involved a series of technical innovations and experiments that imply that an HCC of less than 300 m length can cool the 6d emittance of a muon beam by six orders of magnitude. We describe the design and construction plans for a prototype HCC module based on oxygen-doped hydrogen-pressurized RF cavities that are loaded with dielectric, fed by magnetrons, and operate in a superconducting helical solenoid magnet. The first phase of this project saw the development of a conceptual design for the integration of 805 MHz RF cavities into a 10 T Nb₃Sn-based HS test section. Two very novel ideas are required to realize the design. The first idea is the use of dielectric inserts in the RF cavities to make them smaller for a given frequency so that the cavities and associated plumbing easily fit inside the magnet cryostat. Calculations indicate that heat loads will be tolerable, while RF breakdown of the dielectric inserts will be suppressed by the pressurized hydrogen gas. The second new idea is the use of a multi-layer Nb₃Sn helical solenoid. The technology demonstrations for the two aforementioned key components of a 10T, 805 MHz HCC were begun in this project. The work load in the Fermilab Technical Division made it difficult to test a multi-layer Nb₃Sn solenoid as originally planned. Instead, a complementary

Real-Time Closed Loop Modulated Turbine Cooling

Science.gov (United States)

Shyam, Vikram; Culley, Dennis E.; Eldridge, Jeffrey; Jones, Scott; Woike, Mark; Cuy, Michael

2014-01-01

It has been noted by industry that in addition to dramatic variations of temperature over a given blade surface, blade-to-blade variations also exist despite identical design. These variations result from manufacturing variations, uneven wear and deposition over the life of the part as well as limitations in the uniformity of coolant distribution in the baseline cooling design. It is proposed to combine recent advances in optical sensing, actuation, and film cooling concepts to develop a workable active, closed-loop modulated turbine cooling system to improve by 10 to 20 the turbine thermal state over the flight mission, to improve engine life and to dramatically reduce turbine cooling air usage and aircraft fuel burn. A reduction in oxides of nitrogen (NOx) can also be achieved by using the excess coolant to improve mixing in the combustor especially for rotorcraft engines. Recent patents filed by industry and universities relate to modulating endwall cooling using valves. These schemes are complex, add weight and are limited to the endwalls. The novelty of the proposed approach is twofold 1) Fluidic diverters that have no moving parts are used to modulate cooling and can operate under a wide range of conditions and environments. 2) Real-time optical sensing to map the thermal state of the turbine has never been attempted in realistic engine conditions.

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.

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

International Nuclear Information System (INIS)

Powell, J.R.; Ludewig, H.; Maise, G.

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

Design and qualification of an UHV system for operation on sounding rockets

Energy Technology Data Exchange (ETDEWEB)

Grosse, Jens, E-mail: jens.grosse@dlr.de; Braxmaier, Claus [Center of Applied Space Technology and Microgravity (ZARM), University of Bremen, Bremen, 28359, Germany and German Aerospace Center (DLR) Bremen, Bremen, 28359 (Germany); Seidel, Stephan Tobias; Becker, Dennis; Lachmann, Maike Diana [Institute of Quantum Optics, Leibniz University Hanover, Hanover, 30167 (Germany); Scharringhausen, Marco [German Aerospace Center (DLR) Bremen, Bremen, 28359 (Germany); Rasel, Ernst Maria [Institute of Quantum Optics, Leibniz University Hanover, Hanover, 30167, Bremen (Germany)

2016-05-15

The sounding rocket mission MAIUS-1 has the objective to create the first Bose–Einstein condensate in space; therefore, its scientific payload is a complete cold atom experiment built to be launched on a VSB-30 sounding rocket. An essential part of the setup is an ultrahigh vacuum system needed in order to sufficiently suppress interactions of the cooled atoms with the residual background gas. Contrary to vacuum systems on missions aboard satellites or the international space station, the required vacuum environment has to be reached within 47 s after motor burn-out. This paper contains a detailed description of the MAIUS-1 vacuum system, as well as a description of its qualification process for the operation under vibrational loads of up to 8.1 g{sub RMS} (where RMS is root mean square). Even though a pressure rise dependent on the level of vibration was observed, the design presented herein is capable of regaining a pressure of below 5 × 10{sup −10} mbar in less than 40 s when tested at 5.4 g{sub RMS}. To the authors' best knowledge, it is the first UHV system qualified for operation on a sounding rocket.

Fundamentals and applications of neutron imaging. Application part 3. Application of neutron imaging in aircraft, space rocket, car and gunpowder industries

International Nuclear Information System (INIS)

Ikeda, Yasushi

2007-01-01

Neutron imaging is applied to nondestructive test. Four neutron imaging facilities are used in Japan. The application examples of industries are listed in the table: space rocket, aircraft, car, liquid metal, and works of art. Neutron imaging of transportation equipments are illustrated as an application to industry. X-ray radiography testing (XRT) image and neutron radiography testing (NRT) image of turbine blade of aircraft engine, honeycomb structure of aircraft, helicopter rotor blade, trigger tube, separation nut of space rocket, carburetor of car, BMW engine, fireworks and ammunitions are illustrated. (S.Y.)

Particle bed reactor nuclear rocket concept

International Nuclear Information System (INIS)

Ludewig, H.

1991-01-01

The particle bed reactor nuclear rocket concept consists of fuel particles (in this case (U,Zr)C with an outer coat of zirconium carbide). These particles are packed in an annular bed surrounded by two frits (porous tubes) forming a fuel element; the outer one being a cold frit, the inner one being a hot frit. The fuel element are cooled by hydrogen passing in through the moderator. These elements are assembled in a reactor assembly in a hexagonal pattern. The reactor can be either reflected or not, depending on the design, and either 19 or 37 elements, are used. Propellant enters in the top, passes through the moderator fuel element and out through the nozzle. Beryllium used for the moderator in this particular design to withstand the high radiation exposure implied by the long run times

Engineering Design Handbook. Military Vehicle Power Plant Cooling

Science.gov (United States)

1975-06-01

conductivity of steel and, Indirect cooling is achieved by an interme- therefore, has a much higher heat transfer diary fluid that absorbs heat from...PROBABLZ - -- - - N PERFOEUMACE 0 t0 4 8 21 is s0 24 AIR DELIVEY,. CPU flo 1000-0- U.=rWIM-YEC0*4CAL " AYA OM". 7-6-59 F-igure 8-40. Cooling Fin

Integration of rocket turbine design and analysis through computer graphics

Science.gov (United States)

Hsu, Wayne; Boynton, Jim

1988-01-01

An interactive approach with engineering computer graphics is used to integrate the design and analysis processes of a rocket engine turbine into a progressive and iterative design procedure. The processes are interconnected through pre- and postprocessors. The graphics are used to generate the blade profiles, their stacking, finite element generation, and analysis presentation through color graphics. Steps of the design process discussed include pitch-line design, axisymmetric hub-to-tip meridional design, and quasi-three-dimensional analysis. The viscous two- and three-dimensional analysis codes are executed after acceptable designs are achieved and estimates of initial losses are confirmed.

Numerical Study on Similarity of Plume’s Infrared Radiation from Reduced Scaling Solid Rocket

Directory of Open Access Journals (Sweden)

Xiaoying Zhang

2015-01-01

Full Text Available Similarity of plume radiation between reduced scaling solid rocket models and full scale ones in ground conditions has been taken for investigation. Flow and radiation of plume from solid rockets with scaling ratio from 0.1 to 1 have been computed. The radiative transfer equation (RTE is solved by the finite volume method (FVM in infrared band 2~6 μm. The spectral characteristics of plume gases have been calculated with the weighted-sum-of-gray-gas (WSGG model, and those of the Al2O3 particles have been solved by the Mie scattering model. Our research shows that, with the decreasing scaling ratio of the rocket engine, the radiation intensity of the plume decreases with 1.5~2.5 power of the scaling ratio. The infrared radiation of the plume gases shows a strong spectral dependency, while that of the Al2O3 particles shows grey property. Spectral radiation intensity of the high temperature core of the solid rocket plume increases greatly in the peak absorption spectrum of plume gases. Al2O3 particle is the major radiation composition in the rocket plume, whose scattering coefficient is much larger than its absorption coefficient. There is good similarity between spectral variations of plumes from different scaling solid rockets. The directional plume radiation rises with the increasing azimuth angle.

A study of performance and cost improvement potential of the 120 inch (3.05 m) diameter solid rocket motor. Volume 1: Summary report

Science.gov (United States)

Backlund, S. J.; Rossen, J. N.

1971-01-01

A parametric study of ballistic modifications to the 120 inch diameter solid propellant rocket engine which forms part of the Air Force Titan 3 system is presented. 576 separate designs were defined and 24 were selected for detailed analysis. Detailed design descriptions, ballistic performance, and mass property data were prepared for each design. It was determined that a relatively simple change in design parameters could provide a wide range of solid propellant rocket engine ballistic characteristics for future launch vehicle applications.

14 CFR 101.25 - Operating limitations for Class 2-High Power Rockets and Class 3-Advanced High Power Rockets.

Science.gov (United States)

2010-01-01

... Power Rockets and Class 3-Advanced High Power Rockets. 101.25 Section 101.25 Aeronautics and Space... OPERATING RULES MOORED BALLOONS, KITES, AMATEUR ROCKETS AND UNMANNED FREE BALLOONS Amateur Rockets § 101.25 Operating limitations for Class 2-High Power Rockets and Class 3-Advanced High Power Rockets. When operating...

Influence of cooled exhaust gas recirculation on performance, emissions and combustion characteristics of LPG fuelled lean burn SI engine

Science.gov (United States)

Ravi, K.; Pradeep Bhasker, J.; Alexander, Jim; Porpatham, E.

2017-11-01

On fuel perspective, Liquefied Petroleum Gas (LPG) provides cleaner emissions and also facilitates lean burn signifying less fuel consumption and emissions. Lean burn technology can attain better efficiencies and lesser combustion temperatures but this temperature is quite sufficient to facilitate formation of nitrogen oxide (NOx). Exhaust Gas Recirculation (EGR) for NOx reduction has been considered allover but extremely little literatures exist on the consequence of EGR on lean burn LPG fuelled spark ignition (SI) engine. The following research is carried out to find the optimal rate of EGR addition to reduce NOx emissions without settling on performance and combustion characteristics. A single cylinder diesel engine is altered to operate as LPG fuelled SI engine at a compression ratio of 10.5:1 and arrangements to provide different ratios of cooled EGR in the intake manifold. Investigations are done to arrive at optimum ratio of the EGR to reduce emissions without compromising on performance. Significant reductions in NOx emissions alongside HC and CO emissions were seen. Higher percentages of EGR further diluted the charge and lead to improper combustion and thus increased hydrocarbon emissions. Cooled EGR reduced the peak in-cylinder temperature which reduced NOx emissions but lead to misfire at lower lean limits.

Particle bed reactor propulsion vehicle performance and characteristics as an orbital transfer rocket

International Nuclear Information System (INIS)

Horn, F.L.; Powell, J.R.; Lazareth, O.W.

1986-01-01

The particle bed reactor designed for 100 to 300 MW power output using hydrogen as a coolant is capable of specific impulses up to 1000 seconds as a nuclear rocket. A single space shuttle compatible vehicle can perform extensive missions from LEO to 3 times GEO and return with multi-ton payloads. The use of hydrogen to directly cool particulate reactor fuel results in a compact, lightweight rocket vehicle, whose duration of usefulness is dependent only upon hydrogen resupply availability. The LEO to GEO mission had a payload capability of 15.4 metric tons with 3.4 meters of shuttle bay. To increase the volume limitation of the shuttle bay, the use of ammonia in the initial boost phase from LEO is used to give greater payload volume with a small decrease in payload mass, 8.7 meters and 12.7 m-tons. 5 refs., 15 figs

Nuclear thermal rocket workshop reference system Rover/NERVA

International Nuclear Information System (INIS)

Borowski, S.K.

1991-01-01

The Rover/NERVA engine system is to be used as a reference, against which each of the other concepts presented in the workshop will be compared. The following topics are reviewed: the operational characteristics of the nuclear thermal rocket (NTR); the accomplishments of the Rover/NERVA programs; and performance characteristics of the NERVA-type systems for both Mars and lunar mission applications. Also, the issues of ground testing, NTR safety, NASA's nuclear propulsion project plans, and NTR development cost estimates are briefly discussed

Cooling Effectiveness Measurements for Air Film Cooling of Thermal Barrier Coated Surfaces in a Burner Rig Environment Using Phosphor Thermometry

Science.gov (United States)

Eldridge, Jeffrey I.; Shyam, Vikram; Wroblewski, Adam C.; Zhu, Dongming; Cuy, Michael D.; Wolfe, Douglas E.

2016-01-01

While the effects of thermal barrier coating (TBC) thermal protection and air film cooling effectiveness are usually studied separately, their contributions to combined cooling effectiveness are interdependent and are not simply additive. Therefore, combined cooling effectiveness must be measured to achieve an optimum balance between TBC thermal protection and air film cooling. In this investigation, surface temperature mapping was performed using recently developed Cr-doped GdAlO3 phosphor thermometry. Measurements were performed in the NASA GRC Mach 0.3 burner rig on a TBC-coated plate using a scaled up cooling hole geometry where both the mainstream hot gas temperature and the blowing ratio were varied. Procedures for surface temperature and cooling effectiveness mapping of the air film-cooled TBC-coated surface are described. Applications are also shown for an engine component in both the burner rig test environment as well as an engine afterburner environment. The effects of thermal background radiation and flame chemiluminescence on the measurements are investigated, and advantages of this method over infrared thermography as well as the limitations of this method for studying air film cooling are discussed.

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

Transient Three-Dimensional Side Load Analysis of Out-of-Round Film Cooled Nozzles

Science.gov (United States)

Wang, Ten-See; Lin, Jeff; Ruf, Joe; Guidos, Mike

2010-01-01

The objective of this study is to investigate the effect of nozzle out-of-roundness on the transient startup side loads at a high altitude, with an anchored computational methodology. The out-of-roundness could be the result of asymmetric loads induced by hardware attached to the nozzle, asymmetric internal stresses induced by previous tests, and deformation, such as creep, from previous tests. The rocket engine studied encompasses a regeneratively cooled thrust chamber and a film cooled nozzle extension with film coolant distributed from a turbine exhaust manifold. The computational methodology is based on an unstructured-grid, pressure-based computational fluid dynamics formulation, and a transient inlet history based on an engine system simulation. Transient startup computations were performed with the out-of-roundness achieved by four different degrees of ovalization: one perfectly round, one slightly out-of-round, one more out-of-round, and one significantly out-of-round. The results show that the separation-line-jump is the peak side load physics for the round, slightly our-of-round, and more out-of-round cases, and the peak side load increases as the degree of out-of-roundness increases. For the significantly out-of-round nozzle, however, the peak side load reduces to comparable to that of the round nozzle and the separation line jump is not the peak side load physics. The counter-intuitive result of the significantly out-of-round case is found to be related to a side force reduction mechanism that splits the effect of the separation-line-jump into two parts, not only in the circumferential direction and most importantly in time.

CoolPack – Simulation tools for refrigeration systems

DEFF Research Database (Denmark)

Jakobsen, Arne; Rasmussen, Bjarne D.; Andersen, Simon Engedal

1999-01-01

CoolPack is a collection of programs used for energy analysis and optimisation of refrigeration systems. CoolPack is developed at the Department of Energy Engineering at the Technical University of Denmark. The Danish Energy Agency finances the project. CoolPack is freeware and can be downloaded...

Design for micro-combined cooling, heating and power systems stirling engines and renewable power systems

CERN Document Server

2015-01-01

‘Design for Micro-Combined Cooling, Heating & Power Systems’ provides a manual for the technical and structural design of systems for supplying decentralised energy in residential buildings. It presents the micro-combined cooling, heating & power systems Stirling engines & renewable energy sources (mCCHP-SE-RES) systems in an accessible manner both for the public at large, and for professionals who conceive, design or commercialise such systems or their components.  The high performance levels of these systems are demonstrated within the final chapter by the results of an experiment in which a house is equipped with a mCCHP-SE-RES system. The reader is also familiarized with the conceptual, technical and legal aspects of modern domestic energy systems; the components that constitute these systems; and advanced algorithms for achieving the structural and technical design of such systems. In residential buildings, satisfying demands of durable development has gradually evolved from necessity to...

Helium-Hydrogen Recovery System, Phase I

Data.gov (United States)

National Aeronautics and Space Administration — Immense quantities of expensive liquefied helium are required at Stennis and Kennedy Space Centers for pre-cooling rocket engine propellant systems prior to filling...

Computer Program for Analysis, Design and Optimization of Propulsion, Dynamics, and Kinematics of Multistage Rockets

Science.gov (United States)

Lali, Mehdi

2009-03-01

A comprehensive computer program is designed in MATLAB to analyze, design and optimize the propulsion, dynamics, thermodynamics, and kinematics of any serial multi-staging rocket for a set of given data. The program is quite user-friendly. It comprises two main sections: "analysis and design" and "optimization." Each section has a GUI (Graphical User Interface) in which the rocket's data are entered by the user and by which the program is run. The first section analyzes the performance of the rocket that is previously devised by the user. Numerous plots and subplots are provided to display the performance of the rocket. The second section of the program finds the "optimum trajectory" via billions of iterations and computations which are done through sophisticated algorithms using numerical methods and incremental integrations. Innovative techniques are applied to calculate the optimal parameters for the engine and designing the "optimal pitch program." This computer program is stand-alone in such a way that it calculates almost every design parameter in regards to rocket propulsion and dynamics. It is meant to be used for actual launch operations as well as educational and research purposes.

Development of small solid rocket boosters for the ILR-33 sounding rocket

Science.gov (United States)

Nowakowski, Pawel; Okninski, Adam; Pakosz, Michal; Cieslinski, Dawid; Bartkowiak, Bartosz; Wolanski, Piotr

2017-09-01

This paper gives an overview of the development of a 6000 Newton-class solid rocket motor for suborbital applications. The design configuration and results of interior ballistics calculations are given. The initial use of the motor as the main propulsion system of the H1 experimental in-flight test platform, within the Polish Small Sounding Rocket Program, is presented. Comparisons of theoretical and experimental performance are shown. Both on-ground and in-flight tests are discussed. A novel composite-case manufacturing technology, which enabled to reach high propellant mass fractions, was validated and significant cost-reductions were achieved. This paper focuses on the process of adapting the design for use as the booster stage of the ILR-33 sounding rocket, under development at the Institute of Aviation in Warsaw, Poland. Parallel use of two of the flight-proven rocket motors along with the main stage is planned. The process of adapting the rocket motor for booster application consists of stage integration, aerothermodynamics and reliability analyses. The separation mechanism and environmental impact are also discussed within this paper. Detailed performance analysis with focus on propellant grain geometry is provided. The evolution of the design since the first flights of the H1 rocket is covered and modifications of the manufacturing process are described. Issues of simultaneous ignition of two motors and their non-identical performance are discussed. Further applications and potential for future development are outlined. The presented results are based on the initial work done by the Rocketry Group of the Warsaw University of Technology Students' Space Association. The continuation of the Polish Small Sounding Rocket Program on a larger scale at the Institute of Aviation proves the value of the outcomes of the initial educational project.

Baking Soda and Vinegar Rockets

Science.gov (United States)

Claycomb, James R.; Zachary, Christopher; Tran, Quoc

2009-01-01

Rocket experiments demonstrating conservation of momentum will never fail to generate enthusiasm in undergraduate physics laboratories. In this paper, we describe tests on rockets from two vendors that combine baking soda and vinegar for propulsion. The experiment compared two analytical approximations for the maximum rocket height to the…

Design of a High Temperature Radiator for the Variable Specific Impulse Magnetoplasma Rocket

Science.gov (United States)

Sheth, Rubik B.; Ungar, Eugene K.; Chambliss, Joe P.

2012-01-01

The Variable Specific Impulse Magnetoplasma Rocket (VASIMR), currently under development by Ad Astra Rocket Company (Webster, TX), is a unique propulsion system that could change the way space propulsion is performed. VASIMR's efficiency, when compared to that of a conventional chemical rocket, reduces the propellant needed for exploration missions by a factor of 10. Currently plans include flight tests of a 200 kW VASIMR system, titled VF-200, on the International Space Station (ISS). The VF-200 will consist of two 100 kW thruster units packaged together in one engine bus. Each thruster core generates 27 kW of waste heat during its 15 minute firing time. The rocket core will be maintained between 283 and 573 K by a pumped thermal control loop. The design of a high temperature radiator is a unique challenge for the vehicle design. This paper will discuss the path taken to develop a steady state and transient-based radiator design. The paper will describe the radiator design option selected for the VASIMR thermal control system for use on ISS, and how the system relates to future exploration vehicles.

Multi-Rocket Thought Experiment

Science.gov (United States)

Smarandache, Florentin

2014-03-01

We consider n>=2 identical rockets: R1 ,R2 , ..., Rn. Each of them moving at constant different velocities respectively v1 ,v2 , ..., vn on parallel directions in the same sense. In each rocket there is a light clock, the observer on earth also has a light clock. All n + 1 light clocks are identical and synchronized. The proper time Δt' in each rocket is the same. (1) If we consider the observer on earth and the first rocket R1, then the non-proper time Δt of the observer on earth is dilated with the factor D(v1) : or Δt = Δt' D(v1) (1) But if we consider the observer on earth and the second rocket R2 , then the non-proper time Δt of the observer on earth is dilated with a different factor D(v2) : or Δt = Δt' D(v2) And so on. Therefore simultaneously Δt is dilated with different factors D(v1) , D(v2), ..., D(vn) , which is a multiple contradiction.

Rocket Science 101 Interactive Educational Program

Science.gov (United States)

Armstrong, Dennis; Funkhouse, Deborah; DiMarzio, Donald

2007-01-01

To better educate the public on the basic design of NASA s current mission rockets, Rocket Science 101 software has been developed as an interactive program designed to retain a user s attention and to teach about basic rocket parts. This program also has helped to expand NASA's presence on the Web regarding educating the public about the Agency s goals and accomplishments. The software was designed using Macromedia s Flash 8. It allows the user to select which type of rocket they want to learn about, interact with the basic parts, assemble the parts to create the whole rocket, and then review the basic flight profile of the rocket they have built.

Droplet behaviour in an acoustic field: application to high frequency instability in liquid propellant rocket engines; Comportement de gouttes dans un champ acoustique: applications aux instabilites hautes-frequences dans les moteurs de fusees a ergols liquides

Energy Technology Data Exchange (ETDEWEB)

Boisneau, O.; Lecourt, R.; Grisch, F.; Orain, M.

2002-07-01

A setup has been developed at ONERA in the scope of studying interaction between calibrated droplets and a transversal acoustic wave in the scope of high frequency instabilities in liquid rocket engines. First, the setup has been checked acoustically by hot-wire anemometer and microphone. We present an analytical solution of the Stokes' droplet motion equation in an acoustic field. The trajectory equation can be split into three different parts: a sinusoidal part (negligible in liquid rocket engines), a transient part and a final mean position (only function of the loudspeaker characteristics but never reached). Some kind of vibrational breakup at low Weber's number has been observed using line-of-sight visualization of acoustic/droplet interactions. However, preponderant phenomena observed were jet oscillations and droplet coalescence. For ambient temperature, PLIF visualization has shown a coupling between the created vapor cylinder and the acoustic induced jet position. For hot temperature, some unsteady phenomena seem to appear but further processing are needed. (authors)

Numerical Optimisation in Non Reacting Conditions of the Injector Geometry for a Continuous Detonation Wave Rocket Engine

Science.gov (United States)

Gaillard, T.; Davidenko, D.; Dupoirieux, F.

2015-06-01

The paper presents the methodology and the results of a numerical study, which is aimed at the investigation and optimisation of different means of fuel and oxidizer injection adapted to rocket engines operating in the rotating detonation mode. As the simulations are achieved at the local scale of a single injection element, only one periodic pattern of the whole geometry can be calculated so that the travelling detonation waves and the associated chemical reactions can not be taken into account. Here, separate injection of fuel and oxidizer is considered because premixed injection is handicapped by the risk of upstream propagation of the detonation wave. Different associations of geometrical periodicity and symmetry are investigated for the injection elements distributed over the injector head. To analyse the injection and mixing processes, a nonreacting 3D flow is simulated using the LES approach. Performance of the studied configurations is analysed using the results on instantaneous and mean flowfields as well as by comparing the mixing efficiency and the total pressure recovery evaluated for different configurations.

Study on the Modifications Required to Re-Engine the Lockheed D-21 Drone

Science.gov (United States)

1999-01-01

This report was prepared by Lockheed Martin (LM). The purpose of this 45 day study contract was to investigate the feasibility of using the D-21 as a Rocket Based Combined Cycle engine test-bed. The new NASA engine is entitled "Demonstration of Rocket Combined Cycle Operations (DRACO)". Four objectives were defined and modification study provide an estimation of the: (1) mudified vehicle performance; (2) required engine performance; (3) required vehicle modification; and (4) modification cost and schedule.

A Reusable, Oxidizer-Cooled, Hybrid Aerospike Rocket Motor for Flight Test, Phase I

Data.gov (United States)

National Aeronautics and Space Administration — The proposed innovation is to use the refrigerant capabilities of nitrous oxide (N2O) to provide the cooling required for reusable operation of an aerospike nozzle...

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

Cooling towers principles and practice

CERN Document Server

Hill, G B; Osborn, Peter D

1990-01-01

Cooling Towers: Principles and Practice, Third Edition, aims to provide the reader with a better understanding of the theory and practice, so that installations are correctly designed and operated. As with all branches of engineering, new technology calls for a level of technical knowledge which becomes progressively higher; this new edition seeks to ensure that the principles and practice of cooling towers are set against a background of up-to-date technology. The book is organized into three sections. Section A on cooling tower practice covers topics such as the design and operation of c

Linac drift tube tank upgrade engineering - cooling solution

International Nuclear Information System (INIS)

Li, G.; Heilbrunn, W.; Potter, J.

1999-01-01

Components from the injector of the canceled SSC project are being modified by JPAW to make a commercial radioisotope production linac for I 3 in Denton, TX. The biggest challenge of the upgraded design is the increased average power of the DTL, 40 times the original. With the thermo-mechanical analysis backed by a thorough understanding of the thermal physics, 156 drift tubes have been redesigned according to the RF power deposition. Increasing flow rate in the original cooling channels and adding four flow paths reduces the average tank temperature to an acceptable level. The Δf tolerance budget is controlled without the use of additional temperature control units. The unfinished SSC endwall parts have been modified for additional cooling of the nose and the wall. The different LINAC cooling subsystems are connected to a manifold in parallel through independent flow control valves to balance the required flow rate for each branch

Design study of laser fusion rocket

International Nuclear Information System (INIS)

Nakashima, Hideki; Shoyama, Hidetoshi; Kanda, Yukinori

1991-01-01

A design study was made on a rocket powered by laser fusion. Dependence of its flight performance on target gain, driver repetition rate and fuel composition was analyzed to obtain optimal design parameters of the laser fusion rocket. The results indicate that the laser fusion rocket fueled with DT or D 3 He has the potential advantages over other propulsion systems such as fission rocket for interplanetary travel. (author)

Solid propellant processing factor in rocket motor design

Science.gov (United States)

1971-01-01

The ways are described by which propellant processing is affected by choices made in designing rocket engines. Tradeoff studies, design proof or scaleup studies, and special design features are presented that are required to obtain high product quality, and optimum processing costs. Processing is considered to include the operational steps involved with the lining and preparation of the motor case for the grain; the procurement of propellant raw materials; and propellant mixing, casting or extrusion, curing, machining, and finishing. The design criteria, recommended practices, and propellant formulations are included.

The Chameleon Solid Rocket Propulsion Model

International Nuclear Information System (INIS)

Robertson, Glen A.

2010-01-01

The Khoury and Weltman (2004a and 2004b) Chameleon Model presents an addition to the gravitation force and was shown by the author (Robertson, 2009a and 2009b) to present a new means by which one can view other forces in the Universe. The Chameleon Model is basically a density-dependent model and while the idea is not new, this model is novel in that densities in the Universe to include the vacuum of space are viewed as scalar fields. Such an analogy gives the Chameleon scalar field, dark energy/dark matter like characteristics; fitting well within cosmological expansion theories. In respect to this forum, in this paper, it is shown how the Chameleon Model can be used to derive the thrust of a solid rocket motor. This presents a first step toward the development of new propulsion models using density variations verse mass ejection as the mechanism for thrust. Further, through the Chameleon Model connection, these new propulsion models can be tied to dark energy/dark matter toward new space propulsion systems utilizing the vacuum scalar field in a way understandable by engineers, the key toward the development of such systems. This paper provides corrections to the Chameleon rocket model in Robertson (2009b).

Solar-Thermal Engine Testing

Science.gov (United States)

Tucker, Stephen; Salvail, Pat; Haynes, Davy (Technical Monitor)

2001-01-01

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 (I(sub sp)). The Shooting Star Experiment (SSE) solar-thermal engine is made of 100 percent chemical 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 I(sub sp). 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

Ionospheric shock waves triggered by rockets

Directory of Open Access Journals (Sweden)

C. H. Lin

2014-09-01

Full Text Available This paper presents a two-dimensional structure of the shock wave signatures in ionospheric electron density resulting from a rocket transit using the rate of change of the total electron content (TEC derived from ground-based GPS receivers around Japan and Taiwan for the first time. From the TEC maps constructed for the 2009 North Korea (NK Taepodong-2 and 2013 South Korea (SK Korea Space Launch Vehicle-II (KSLV-II rocket launches, features of the V-shaped shock wave fronts in TEC perturbations are prominently seen. These fronts, with periods of 100–600 s, produced by the propulsive blasts of the rockets appear immediately and then propagate perpendicularly outward from the rocket trajectory with supersonic velocities between 800–1200 m s−1 for both events. Additionally, clear rocket exhaust depletions of TECs are seen along the trajectory and are deflected by the background thermospheric neutral wind. Twenty minutes after the rocket transits, delayed electron density perturbation waves propagating along the bow wave direction appear with phase velocities of 800–1200 m s−1. According to their propagation character, these delayed waves may be generated by rocket exhaust plumes at earlier rocket locations at lower altitudes.