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Sample records for thermal rocket lpntr

  1. Low Pressure Nuclear Thermal Rocket (LPNTR) concept

    International Nuclear Information System (INIS)

    Ramsthaler, J.H.

    1991-01-01

    A background and a description of the low pressure nuclear thermal system are presented. Performance, mission analysis, development, critical issues, and some conclusions are discussed. The following subject areas are covered: LPNTR's inherent advantages in critical NTR requirement; reactor trade studies; reference LPNTR; internal configuration and flow of preliminary LPNTR; particle bed fuel assembly; preliminary LPNTR neutronic study results; multiple LPNTR engine concept; tank and engine configuration for mission analysis; LPNTR reliability potential; LPNTR development program; and LPNTR program costs

  2. Alternate Propellant Thermal Rocket, Phase I

    Data.gov (United States)

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

  3. Preliminary design study for a carbide LEU-nuclear thermal rocket

    International Nuclear Information System (INIS)

    Venneri, P.F.; Kim, Y.

    2014-01-01

    Nuclear space propulsion is a requirement for the successful exploration of the solar system. It offers the possibility of having both a high specific impulse and a relatively high thrust, allowing rapid transit times with a minimum usage of fuel. This paper proposes a nuclear thermal rocket design based on heritage NERVA rockets that makes use of Low Enriched Uranium (LEU) fuel. The Carbide LEU Nuclear Thermal Rocket (C-LEU-NTR) is designed to fulfill the rocket requirements as set forth in the NASA 2009 Mars Mission Design Reference Architecture 5.0, that is provide 25,000 lbf of thrust, operate at full power condition for at least two hours, and have a specific impulse close to 900 s. The neutronics analysis was done using MCNP5 with the ENDF/B-VII.1 neutron library. The thermal hydraulic calculations and size optimization were completed with a finite difference code being developed at the Center for Space Nuclear Research. (authors)

  4. Unique nuclear thermal rocket engine

    International Nuclear Information System (INIS)

    Culver, D.W.; Rochow, R.

    1993-06-01

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

  5. Nuclear thermal rocket nozzle testing and evaluation program

    International Nuclear Information System (INIS)

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

    1993-01-01

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

  6. 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

  7. Feasibility and Performance of the Microwave Thermal Rocket Launcher

    Science.gov (United States)

    Parkin, Kevin L. G.; Culick, Fred E. C.

    2004-03-01

    Beamed-energy launch concepts employing a microwave thermal thruster are feasible in principle, and microwave sources of sufficient power to launch tons into LEO already exist. Microwave thermal thrusters operate on an analogous principle to nuclear thermal thrusters, which have experimentally demonstrated specific impulses exceeding 850 seconds. Assuming such performance, simple application of the rocket equation suggests that payload fractions of 10% are possible for a single stage to orbit (SSTO) microwave thermal rocket. We present an SSTO concept employing a scaled X-33 aeroshell. The flat aeroshell underside is covered by a thin-layer microwave absorbent heat-exchanger that forms part of the thruster. During ascent, the heat-exchanger faces the microwave beam. A simple ascent trajectory analysis incorporating X-33 aerodynamic data predicts a 10% payload fraction for a 1 ton craft of this type. In contrast, the Saturn V had 3 non-reusable stages and achieved a payload fraction of 4%.

  8. 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

  9. 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.

  10. Feasibility and Performance of the Microwave Thermal Rocket Launcher

    OpenAIRE

    Parkin, Kevin L. G.; Culick, Fred E. C.

    2004-01-01

    Beamed-energy launch concepts employing a microwave thermal thruster are feasible in principle, and microwave sources of sufficient power to launch tons into LEO already exist. Microwave thermal thrusters operate on an analogous principle to nuclear thermal thrusters, which have experimentally demonstrated specific impulses exceeding 850 seconds. Assuming such performance, simple application of the rocket equation suggests that payload fractions of 10% are possible for a single stage to orbit...

  11. A computational model for thermal fluid design analysis of nuclear thermal rockets

    International Nuclear Information System (INIS)

    Given, J.A.; Anghaie, S.

    1997-01-01

    A computational model for simulation and design analysis of nuclear thermal propulsion systems has been developed. The model simulates a full-topping expander cycle engine system and the thermofluid dynamics of the core coolant flow, accounting for the real gas properties of the hydrogen propellant/coolant throughout the system. Core thermofluid studies reveal that near-wall heat transfer models currently available may not be applicable to conditions encountered within some nuclear rocket cores. Additionally, the possibility of a core thermal fluid instability at low mass fluxes and the effects of the core power distribution are investigated. Results indicate that for tubular core coolant channels, thermal fluid instability is not an issue within the possible range of operating conditions in these systems. Findings also show the advantages of having a nonflat centrally peaking axial core power profile from a fluid dynamic standpoint. The effects of rocket operating conditions on system performance are also investigated. Results show that high temperature and low pressure operation is limited by core structural considerations, while low temperature and high pressure operation is limited by system performance constraints. The utility of these programs for finding these operational limits, optimum operating conditions, and thermal fluid effects is demonstrated

  12. 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.

  13. On the use of a pulsed nuclear thermal rocket for interplanetary travel

    OpenAIRE

    Arias Montenegro, Francisco Javier

    2016-01-01

    The object of this work is a first assessment of the use of a pulsed nuclear thermal rocket for thrust and specific impulse (Isp) augmentation with particular reference to interplanetary travel. The basis of the novel space propulsion idea is the possibility of working in a bimodal fashion where the classical stationary nuclear thermal rocket (NTR) could be switch on or switch off as a pulsed reactor as desired by the mission planners. It was found that the key factor for Isp augmentation ...

  14. 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

  15. 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.

  16. 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

  17. Study of organic ablative thermal-protection coating for solid rocket motor

    Science.gov (United States)

    Hua, Zenggong

    1992-06-01

    A study is conducted to find a new interior thermal-protection material that possesses good thermal-protection performance and simple manufacturing possibilities. Quartz powder and Cr2O3 are investigated using epoxy resin as a binder and Al2O3 as the burning inhibitor. Results indicate that the developed thermal-protection coating is suitable as ablative insulation material for solid rocket motors.

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

    Science.gov (United States)

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

    2015-01-01

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

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

    Science.gov (United States)

    Emrich, William J. Jr.; Moran, Robert P.; Pearson, J. Boise

    2012-01-01

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

  20. 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.

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

    Science.gov (United States)

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

    2013-01-01

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

  2. 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

  3. 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

  4. 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

  5. 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

  6. 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

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

    International Nuclear Information System (INIS)

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

    1992-01-01

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

  8. 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)

  9. 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

  10. 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

  11. 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.

  12. 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

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

    International Nuclear Information System (INIS)

    Walton, J.T.

    1992-11-01

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

  14. 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.

  15. Nuclear Thermal Rocket Design Using LEU Tungsten Fuel

    International Nuclear Information System (INIS)

    Venneri, Paolo; Kim, Yonghee; Husemeyer, Peter and others

    2013-01-01

    This would then open the possibility for the commercial development and implementation of an NTR. The result was a design for a 114.66 kN thrust rocket engine, with an optimized specific impulse of 801 second, and a thrust-to-weight ratio 5.08. The development and analysis of the reactor was done using an integrated neutronics and thermal hydraulics code that combines MCNP5 using ENDF-B/VI cross sections with a purpose-built thermal hydraulics code. A proof of concept has been proposed for W LEU-NTR design. The current design is built upon traditional NTR design work and implements many of the proven design characteristics and materials from previous designs. Despite the current reactor design being preliminary, it already shows promise in being able to have similar, if not better performance characteristics than current and previous NTR designs. Future work will involve the flattening of radial power profile, optimization of the axial power profile, researching methods to address the full water immersion accident scenario, and further studies regarding the breeding potential in the reactor

  16. Nuclear Thermal Rocket Design Using LEU Tungsten Fuel

    Energy Technology Data Exchange (ETDEWEB)

    Venneri, Paolo; Kim, Yonghee; Husemeyer, Peter and others

    2013-10-15

    This would then open the possibility for the commercial development and implementation of an NTR. The result was a design for a 114.66 kN thrust rocket engine, with an optimized specific impulse of 801 second, and a thrust-to-weight ratio 5.08. The development and analysis of the reactor was done using an integrated neutronics and thermal hydraulics code that combines MCNP5 using ENDF-B/VI cross sections with a purpose-built thermal hydraulics code. A proof of concept has been proposed for W LEU-NTR design. The current design is built upon traditional NTR design work and implements many of the proven design characteristics and materials from previous designs. Despite the current reactor design being preliminary, it already shows promise in being able to have similar, if not better performance characteristics than current and previous NTR designs. Future work will involve the flattening of radial power profile, optimization of the axial power profile, researching methods to address the full water immersion accident scenario, and further studies regarding the breeding potential in the reactor.

  17. Initial risk assessment for a single stage to orbit nuclear thermal rocket

    Energy Technology Data Exchange (ETDEWEB)

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

    2015-06-15

    Highlights: • The risks posed by the surface launch of a nuclear thermal rocket are considered. • Radiation exposure at the public viewing distance is insignificant. • Production of fission products and actinides during launch is limited. • The production of activated argon around the rocket may be a significant concern. - Abstract: In order to consider the possibility of a nuclear thermal rocket (NTR) ground launch, it is necessary to evaluate the risks from such a launch. This includes analysis of the radiation dose rate around the rocket, determining the rate of activation of the materials near the launch, and considering the radionuclides present in the core after the launch. This paper evaluates the potential risk of the NTR ground launch for a range of payloads from 1 to 15 metric tons (MT) using three NTR reactor cores (40, 80, and 120 cm in length) designed in a previous study, based on data produced by MCNP5 and MCNPX models. At the same power level, the 40 cm core length reactor results in the lowest radiation dose rate of the three reactors. Radiation dose rates decrease to background levels 3.5 km from the launch site. After a 1-year decay time, all of the activated materials produced by an NTR launch would be classified as Class A low-level waste. The activation of air produces significant amounts of argon-41 and nitrogen-16 within 100 m of the launch. The derived air concentration (DAC) ratio of the activation products decays to less than unity within 2 days, with only argon-41 remaining. After 10 min of full power operation, the 120 cm core for a 15 MT payload contains 2.5 × 10{sup 13}, 1.4 × 10{sup 12} and 1.5 × 10{sup 12} Bq of {sup 131}I, {sup 137}Cs, and {sup 90}Sr, respectively. The decay heat after shutdown increases with increasing reactor power with a maximum decay heat of 108 kW immediately after shutdown for the 15 MT payload.

  18. Modification of Bonding Strength Test of WC HVOF Thermal Spray Coating on Rocket Nozzle

    Directory of Open Access Journals (Sweden)

    Bondan Sofyan

    2010-10-01

    Full Text Available One way to reduce structural weight of RX-100 rocket is by modifying the nozzle material and processing. Nozzle is the main target in weight reduction due to the fact that it contributes 30 % to the total weight of the structur. An alternative for this is by substitution of massive graphite, which is currently used as thermal protector in the nozzle, with thin layer of HVOF (High Velocity Oxy-Fuel thermal spray layer. This paper presents the characterization of nozzle base material as well as the modification of bonding strength test, by designing additional jig to facilitate testing processes while maintaining level of test accuracy. The results showed that the material used for  RX-100 rocket nozzle is confirmed to be S45C steel. Modification of the bonding strength test was conducted by utilizing chains, which improve test flexibility and maintains level of accuracy of the test.

  19. Nuclear thermal rockets using indigenous extraterrestrial propellants

    International Nuclear Information System (INIS)

    Zubrin, R.M.

    1990-01-01

    A preliminary examination of a concept for a Mars and outer solar system exploratory vehicle is presented. Propulsion is provided by utilizing a nuclear thermal reactor to heat a propellant volatile indigenous to the destination world to form a high thrust rocket exhaust. Candidate propellants, whose performance, materials compatibility, and ease of acquisition are examined and include carbon dioxide, water, methane, nitrogen, carbon monoxide, and argon. Ballistics and winged supersonic configurations are discussed. It is shown that the use of this method of propulsion potentially offers high payoff to a manned Mars mission. This is accomplished by sharply reducing the initial mission mass required in low earth orbit, and by providing Mars explorers with greatly enhanced mobility in traveling about the planet through the use of a vehicle that can refuel itself each time it lands. Thus, the nuclear landing craft is utilized in combination with a hydrogen-fueled nuclear-thermal interplanetary launch. By utilizing such a system in the outer solar system, a low level aerial reconnaissance of Titan combined with a multiple sample return from nearly every satellite of Saturn can be accomplished in a single launch of a Titan 4 or the Space Transportation System (STS). Similarly a multiple sample return from Callisto, Ganymede, and Europa can also be accomplished in one launch of a Titan 4 or the STS

  20. Investigation of Post-Flight Solid Rocket Booster Thermal Protection System

    Science.gov (United States)

    Nelson, Linda A.

    2006-01-01

    After every Shuttle mission, the Solid Rocket Boosters (SRBs) are recovered and observed for missing material. Most of the SRB is covered with a cork-based thermal protection material (MCC-l). After the most recent shuttle mission, STS-114, the forward section of the booster appeared to have been impacted during flight. The darkened fracture surfaces indicated that this might have occurred early in flight. The scope of the analysis included microscopic observations to assess the degree of heat effects and locate evidence of the impact source as well as chemical analysis of the fracture surfaces and recovered foreign material using Fourier Transform Infrared Spectroscopy and Scanning Electron Microscopy/Energy Dispersive Spectroscopy. The amount of heat effects and presence of soot products on the fracture surface indicated that the material was impacted prior to SRB re-entry into the atmosphere. Fragments of graphite fibers found on these fracture surfaces were traced to slag inside the Solid Rocket Motor (SRM) that forms during flight as the propellant is spent and is ejected throughout the descent of the SRB after separation. The direction of the impact mark matches with the likely trajectory of SRBs tumbling prior to re-entry.

  1. 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

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

    International Nuclear Information System (INIS)

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

    1991-01-01

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

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

    Science.gov (United States)

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

    2011-01-01

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

  4. 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

  5. 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

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

    Science.gov (United States)

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

    2012-12-01

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

  7. 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

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

    Energy Technology Data Exchange (ETDEWEB)

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

    2015-05-15

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

  9. 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

  10. 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.

  11. Nuclear thermal rocket plume interactions with spacecraft. Final report

    International Nuclear Information System (INIS)

    Mauk, B.H.; Gatsonis, N.A.; Buzby, J.; Yin, X.

    1997-01-01

    This is the first study that has treated the Nuclear Thermal Rocket (NTR) effluent problem in its entirety, beginning with the reactor core, through the nozzle flow, to the plume backflow. The summary of major accomplishments is given below: (1) Determined the NTR effluents that include neutral, ionized and radioactive species, under typical NTR chamber conditions. Applied an NTR chamber chemistry model that includes conditions and used nozzle geometries and chamber conditions typical of NTR configurations. (2) Performed NTR nozzle flow simulations using a Navier-Stokes solver. We assumed frozen chemistry at the chamber conditions and used nozzle geometries and chamber conditions typical of NTR configurations. (3) Performed plume simulations using a Direct Simulation Monte Carlo (DSMC) code with chemistry. In order to account for radioactive trace species that may be important for contamination purposes we developed a multi-weighted DSMC methodology. The domain in our simulations included large regions downstream and upstream of the exit. Inputs were taken from the Navier-Stokes solutions

  12. Design methods in solid rocket motors

    Energy Technology Data Exchange (ETDEWEB)

    1987-03-01

    A compilation of lectures summarizing the current state-of-the-art in designing solid rocket motors and and their components is presented. The experience of several countries in the use of new technologies and methods is represented. Specific sessions address propellant grains, cases, nozzles, internal thermal insulation, and the general optimization of solid rocket motor designs.

  13. 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.

  14. 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

  15. Analysis of In Situ Thermal Ion Measurements from the MICA Sounding Rocket

    Science.gov (United States)

    Fernandes, P. A.; Lynch, K. A.; Zettergren, M. D.; Hampton, D. L.; Fisher, L. E.; Powell, S. P.

    2014-12-01

    The MICA sounding rocket launched on 19 Feb. 2012 into several discrete, localized arcs in the wake of a westward traveling surge. In situ and ground-based observations provide a measured response of the ionosphere to preflight and localized auroral drivers. Initial analysis of the in situ thermal ion data indicate possible measurement of an ion conic at low altitude (< 325 km). In the low-energy regime, the response of the instrument varies from the ideal because the measured thermal ion population is sensitive to the presence of the instrument. The plasma is accelerated in the frame of the instrument due to flows, ram, and acceleration through the sheath which forms around the spacecraft. The energies associated with these processes are large compared to the thermal energy. Correct interpretation of thermal plasma measurements requires accounting for all of these plasma processes and the non-ideal response of the instrument in the low-energy regime. This is an experimental and modeling project which involves thorough analysis of ionospheric thermal ion data from the MICA campaign. Analysis includes modeling and measuring the instrument response in the low-energy regime as well as accounting for the complex sheath formed around the instrument. This results in a forward model in which plasma parameters of the thermal plasma are propagated through the sheath and instrument models, resulting in an output which matches the in situ measurement. In the case of MICA, we are working toward answering the question of the initiating source processes that result, at higher altitudes, in well-developed conics and outflow on auroral field lines.

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

    Science.gov (United States)

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

    2015-01-01

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

  17. 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

  18. 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.

  19. 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.

  20. Convection and dendrite crystallization. [during coasting phase of sounding rocket flight

    Science.gov (United States)

    Grodzka, P. G.; Johnston, M. H.; Griner, C. S.

    1977-01-01

    The convection and thermal conditions in aqueous and metallic liquid systems under conditions of the Dendrite Remelting Rocket Experiment were assessed to help establish the relevance of the rocket experiment to the metals casting phenomena. The results of the study indicate that aqueous or metallic convection velocities in the cell are of insignificant magnitudes at the 0.0001 to 0.00001 g levels of the experiment. The crystallization phenomena observed in the rocket experiment, therefore, may be indicative of how metals will solidify in low-g. The influence of possibly differing thermal fields, however, remains to be assessed. The rocket experiment may also be relevant to how metals solidify on the ground at temperature differences and in cell configurations such that the flow velocities are not high enough to break or bend delicate dendrite arms. Again, however, the influence of the thermal fields must be assessed.

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

    Science.gov (United States)

    Emrich, William J., Jr.

    2014-01-01

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

  2. 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.

  3. 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.

  4. 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.

  5. 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

  6. 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...

  7. An Analysis of Ionospheric Thermal Ions Using a SIMION-based Forward Instrument Model: In Situ Observations of Vertical Thermal Ion Flows as Measured by the MICA Sounding Rocket

    Science.gov (United States)

    Fernandes, P. A.; Lynch, K. A.; Zettergren, M. D.; Hampton, D. L.; Fisher, L. E.; Powell, S. P.

    2013-12-01

    The MICA sounding rocket launched on 19 Feb. 2012 into several discrete, localized arcs in the wake of a westward traveling surge. In situ and ground-based observations provide a measured response of the ionosphere to preflight and localized auroral drivers. In this presentation we focus on in situ measurements of the thermal ion distribution. We observe thermal ions flowing both up and down the auroral field line, with upflows concentrated in Alfvénic and downward current regions. The in situ data are compared with recent ionospheric modeling efforts (Zettergren et al., this session) which show structured patterns of ion upflow and downflow consistent with these observations. In the low-energy thermal plasma regime, instrument response to the measured thermal ion population is very sensitive to the presence of the instrument. The plasma is shifted and accelerated in the frame of the instrument due to flows, ram, and acceleration through the payload sheath. The energies associated with these processes are large compared to the thermal energy. Rigorous quantitative analysis of the instrument response is necessary to extract the plasma properties which describe the full 3D distribution function at the instrument aperture. We introduce an instrument model, developed in the commercial software package SIMION, to characterize instrument response at low energies. The instrument model provides important insight into how we would modify our instrument for future missions, including fine-tuning parameters such as the analyzer sweep curve, the geometry factor, and the aperture size. We use the results from the instrument model to develop a forward model, from which we can extract anisotropic ion temperatures, flows, and density of the thermal plasma at the aperture. Because this plasma has transited a sheath to reach the aperture, we must account for the acceleration due to the sheath. Modeling of this complex sheath is being conducted by co-author Fisher, using a PIC code

  8. 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.

  9. 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.

  10. An Internal Thermal Environment Model of an Aluminized Solid Rocket Motor with Experimental Validation

    Science.gov (United States)

    Martin, Heath T.

    2015-01-01

    Due to the severity of the internal solid rocket motor (SRM) environment, very few direct measurements of that environment exist; therefore, the appearance of such data provides a unique opportunity to assess current thermal/fluid modeling capabilities. As part of a previous study of SRM internal insulation performance, the internal thermal environment of a laboratory-scale SRM featuring aluminized propellant was characterized with two types of custom heat-flux calorimeters: one that measured the total heat flux to a graphite slab within the SRM chamber and another that measured the thermal radiation flux. Therefore, in the current study, a thermal/fluid model of this lab-scale SRM was constructed using ANSYS Fluent to predict not only the flow field structure within the SRM and the convective heat transfer to the interior walls, but also the resulting dispersion of alumina droplets and the radiative heat transfer to the interior walls. The dispersion of alumina droplets within the SRM chamber was determined by employing the Lagrangian discrete phase model that was fully coupled to the Eulerian gas-phase flow. The P1-approximation was engaged to model the radiative heat transfer through the SRM chamber where the radiative contributions of the gas phase were ignored and the aggregate radiative properties of the alumina dispersion were computed from the radiative properties of its individual constituent droplets, which were sourced from literature. The convective and radiative heat fluxes computed from the thermal/fluid model were then compared with those measured in the lab-scale SRM test firings and the modeling approach evaluated.

  11. Elastomeric Thermal Insulation Design Considerations in Long, Aluminized Solid Rocket Motors

    Science.gov (United States)

    Martin, Heath T.

    2017-01-01

    An all-new sounding rocket was designed at NASA's Marshall Space Flight Center that featured an aft finocyl, aluminized solid propellant grain and silica-filled ethylene-propylene-diene monomer (SFEPDM) internal insulation. Upon the initial static firing of the first of this new design, the solid rocket motor (SRM) case failed thermally just upstream of the aft closure early in the burn time. Subsequent fluid modeling indicated that the high-velocity combustion-product jets emanating from the fin-slots in the propellant grain were likely inducing a strongly swirling flow, thus substantially increasing the severity of the convective environment on the exposed portion of the SFEPDM insulation in this region. The aft portion of the fin-slots in another of the motors were filled with propellant to eliminate the possibility of both direct jet impingement on the exposed SFEPDM and the appearance of strongly swirling flow in the aft region of the motor. When static-fired, this motor's case still failed in the same axial location, and, though somewhat later than for the first static firing, still in less than 1/3rd of the desired burn duration. These results indicate that the extreme material decomposition rates of the SFEPDM in this application are not due to gas-phase convection or shear but rather to interactions with burning aluminum or alumina slag. Further comparisons with between SFEPDM performance in this design and that in other hot-fire tests provide insight into the mechanisms of SFEPDM decomposition in SRM aft domes that can guide the upcoming redesign effort, as well as other future SRM designs. These data also highlight the current limitations of modeling elastomeric insulators solely with diffusion-controlled, gas-phase thermochemistry in SRM regions with significant viscous shear and/or condense-phase impingement or flow.

  12. 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.

  13. 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

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

    Science.gov (United States)

    Borowski, Stanley K.

    1994-09-01

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

  15. 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.

  16. Microwave Thermal Propulsion

    Science.gov (United States)

    Parkin, Kevin L. G.; Lambot, Thomas

    2017-01-01

    We have conducted research in microwave thermal propulsion as part of the space exploration access technologies (SEAT) research program, a cooperative agreement (NNX09AF52A) between NASA and Carnegie Mellon University. The SEAT program commenced on the 19th of February 2009 and concluded on the 30th of September 2015. The DARPA/NASA Millimeter-wave Thermal Launch System (MTLS) project subsumed the SEAT program from May 2012 to March 2014 and one of us (Parkin) served as its principal investigator and chief engineer. The MTLS project had no final report of its own, so we have included the MTLS work in this report and incorporate its conclusions here. In the six years from 2009 until 2015 there has been significant progress in millimeter-wave thermal rocketry (a subset of microwave thermal rocketry), most of which has been made under the auspices of the SEAT and MTLS programs. This final report is intended for multiple audiences. For researchers, we present techniques that we have developed to simplify and quantify the performance of thermal rockets and their constituent technologies. For program managers, we detail the facilities that we have built and the outcomes of experiments that were conducted using them. We also include incomplete and unfruitful lines of research. For decision-makers, we introduce the millimeter-wave thermal rocket in historical context. Considering the economic significance of space launch, we present a brief but significant cost-benefit analysis, for the first time showing that there is a compelling economic case for replacing conventional rockets with millimeter-wave thermal rockets.

  17. 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.

  18. 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.

  19. A multilayered thick cylindrical shell under internal pressure and thermal loads applicable to solid propellant rocket motors

    Energy Technology Data Exchange (ETDEWEB)

    Renganathan, K.; Nageswara Rao, B.; Jana, M.K. [Vikram Sarabhai Space Centre, Trivandrum (India). Structural Engineering Group

    2000-09-01

    A solid propellant rocket motor can be considered to be made of various circumferential layers of different properties. A simple procedure is described here to obtain an analytical solution for the general case of multilayered thick cyclindrical shell for internal pressure and thermal loads. This analytical procedure is useful in the preliminary design analysis of solid propellant rocket motors. Since solid propellant material is of viscoelastic behaviour an approximate viscoelastic solution methodology for the multilayered shell is described for estimation of time dependent solutions of propellant grain in a rocket motor. The analytical solution for a two layer reinforced thick cylindrical shell available in the literature is shown to be a special case of the present analytical solution. The results from the present analytical solution for multilayers is found to be in good agreement with FEA results. (orig.) [German] Der grundlegende Aufbau von Feststoffraketenmotoren kann auf einen Zylinder aus mehreren Schichten mit unterschiedlichen Eigenschaften zurueckgefuehrt werden. Eine einfache Berechnungsprozedur fuer die analytische Loesung des allgemeinen Falles eines mehrschichtigen Zylinders unter innerem Druck und thermischer Belastung wird hier vorgestellt. Diese analytische Methodik ist fuer den Auslegungsprozess von Feststoffraketenmotoren von grundlegender Bedeutung. Das viskoelastische Fliessverhalten des festen Brennstoffes, das den zeitlichen Ablauf des Verbrennungsprozesses wesentlich bestimmt, wird durch ein Naeherungsverfahren gut erfasst. Ein in der Literatur enthaltenes spezielles Ergebnis fuer einen zweischaligen verstaerkten Zylinder ergibt sich als Sonderfall der hier vorgestellten Methodik. Die analytisch erhaltenen Loesungen fuer mehrschichtige Aufbauten sind in guter Uebereinstimmung mit mittels der FEM ermittelten Ergebnisse. (orig.)

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

    Energy Technology Data Exchange (ETDEWEB)

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

    2012-07-01

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

  1. 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

  2. Non-destructive testing of rocket fuse by thermal neutron radiography

    International Nuclear Information System (INIS)

    An Fulin; Li Furong

    1999-01-01

    A neutron radiography system in reactor horizontal hole of Tsinghua University was introduced, and its capability of neutron radiography was evaluated by theory and experiment, the non-destructive testing for rocket fuse is successful

  3. On the coordination of EISCAT measurements with rocket and satellite observations

    International Nuclear Information System (INIS)

    Hultqvist, B.

    1977-01-01

    The scientific interest of combining EISCAT measurements of the thermal ionospheric plasma with sounding rocket and/or satellite measurements of the hot plasma distribution function and other variables is discussed briefly. Some examples are presented where such coordinated measurements are of great interest. The importance of being able to launch rockets through, or at least quite close to, the radar beam is emphasized. (Auth.)

  4. 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

  5. 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

  6. 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

  7. 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.

  8. 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.

  9. Antithermal shield for rockets with heat evacuation by infrared radiation reflection

    Directory of Open Access Journals (Sweden)

    Ioan RUSU

    2010-12-01

    Full Text Available At high speed, the friction between the air mass and the rocket surface causes a localheating of over 1000 Celsius degrees. For the heat protection of the rocket, on its outside surfacethermal shields are installed.Studying the Coanda effect, the fluid flow on solids surface, respectively, the author Ioan Rusuhas discovered by simply researches that the Coanda effect could be /extended also to the fluid flowon discontinuous solids, namely, on solids provided with orifices. This phenomenon was named by theauthor, the expanded Coanda effect. Starting with this discovery, the author has invented a thermalshield, registered at The State Office for inventions and Trademarks OSIM, deposit F 2010 0153This thermal shield:- is built as a covering rocket sheet with many orifices installed with a minimum space fromthe rocket body- takes over the heat fluid generated by the frontal part of the rocket and avoids the directcontact between the heat fluid and the rocket body- ensures the evacuation of the infrared radiation, generated by the heat fluid flowing overthe shield because of the extended Coanda effect by reflection from the rocket bodysurface.

  10. 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.

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

    Science.gov (United States)

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

    2016-01-01

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

  12. 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.

  13. 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.

  14. 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.

  15. 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)

  16. A review of the Los Alamos effort in the development of nuclear rocket propulsion

    International Nuclear Information System (INIS)

    Durham, F.P.; Kirk, W.L.; Bohl, R.J.

    1991-01-01

    This paper reviews the achievements of the Los Alamos nuclear rocket propulsion program and describes some specific reactor design and testing problems encountered during the development program along with the progress made in solving these problems. The relevance of these problems to a renewed nuclear thermal rocket development program for the Space Exploration Initiative (SEI) is discussed. 11 figs

  17. 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.

  18. Examination of Cross-Scale Coupling During Auroral Events using RENU2 and ISINGLASS Sounding Rocket Data.

    Science.gov (United States)

    Kenward, D. R.; Lessard, M.; Lynch, K. A.; Hysell, D. L.; Hampton, D. L.; Michell, R.; Samara, M.; Varney, R. H.; Oksavik, K.; Clausen, L. B. N.; Hecht, J. H.; Clemmons, J. H.; Fritz, B.

    2017-12-01

    The RENU2 sounding rocket (launched from Andoya rocket range on December 13th, 2015) observed Poleward Moving Auroral Forms within the dayside cusp. The ISINGLASS rockets (launched from Poker Flat rocket range on February 22, 2017 and March 2, 2017) both observed aurora during a substorm event. Despite observing very different events, both campaigns witnessed a high degree of small scale structuring within the larger auroral boundary, including Alfvenic signatures. These observations suggest a method of coupling large-scale energy input to fine scale structures within aurorae. During RENU2, small (sub-km) scale drivers persist for long (10s of minutes) time scales and result in large scale ionospheric (thermal electron) and thermospheric response (neutral upwelling). ISINGLASS observations show small scale drivers, but with short (minute) time scales, with ionospheric response characterized by the flight's thermal electron instrument (ERPA). The comparison of the two flights provides an excellent opportunity to examine ionospheric and thermospheric response to small scale drivers over different integration times.

  19. 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.

  20. Cryogenic rocket engine development at Delft aerospace rocket engineering

    NARCIS (Netherlands)

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

    2016-01-01

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

  1. 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.

  2. Terrier Black Brant VC design characteristics and program status. [rocket development

    Science.gov (United States)

    Payne, B. R.; Mayo, E. E.

    1979-01-01

    In the present paper, the design analysis of the Terrier-Black Brant VC, representing the latest addition to the Black Brant rocket family, is discussed, including the aerodynamic, structural, thermal, and operational aspects. An appreciable increase in apogee, as compared to the BBVC and Nike/BBVC, is achieved without any modifications to the well-proven BBV motor or degradation of the thermal or dynamic flight environment.

  3. 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.

  4. 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.

  5. Design Analysis of a High Temperature Radiator for the Variable Specific Impulse Magnetoplasma Rocket (VASIMR)

    Science.gov (United States)

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

    2011-01-01

    The Variable Specific Impulse Magnetoplasma Rocket (VASIMR), currently under development by Ad Astra Rocket Company, is a unique propulsion system that can potentially change the way space propulsion is performed. VASIMR's efficiency, when compared to that of a conventional chemical rocket, reduce 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. The VF-200 will consist of two 100 kW thruster units packaged together in one engine bus. Each thruster unit has a unique heat rejection requirement of about 27 kW over a firing time of 15 minutes. In order to control rocket core temperatures, peak operating temperatures of about 300 C are expected within the thermal control loop. 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 radiator design options for the VASIMR thermal control system for use on ISS as well as future exploration vehicles.

  6. 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.

  7. 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

  8. 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.

  9. Polymer degradation rate control of hybrid rocket combustion

    Science.gov (United States)

    Stickler, D. B.; Ramohalli, K. N. R.

    1970-01-01

    Polymer degradation to small fragments is treated as a rate controlling step in hybrid rocket combustion. Both numerical and approximate analytical solutions of the complete energy and polymer chain bond conservation equations for the condensed phase are obtained. Comparison with inert atmosphere data is very good. It is found that the intersect of curves of pyrolysis rate versus interface temperature for hybrid combustors, with the thermal degradation theory, falls at a pyrolysis rate very close to that for which a pressure dependence begins to be observable. Since simple thermal degradation cannot give sufficient depolymerization at higher pyrolysis rates, it is suggested that oxidative catalysis of the process occurs at the surface, giving a first order dependence on reactive species concentration at the wall. Estimates of the ratio of this activation energy and interface temperature are in agreement with best fit procedures for hybrid combustion data. Requisite active species concentrations and flux are shown to be compatible with turbulent transport. Pressure dependence of hybrid rocket fuel regression rate is thus shown to be describable in a consistent manner in terms of reactive species catalysis of polymer degradation.

  10. 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

  11. 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.

  12. 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

  13. Bimodal Nuclear Thermal Rocket Analysis Developments

    Science.gov (United States)

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

    2014-01-01

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

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

    Science.gov (United States)

    Jones, Hansen; Jeansonne, Christopher; Menon, Shyam

    2017-11-01

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

  15. 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.

  16. 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)

  17. Origin of how steam rockets can reduce space transport cost by orders of magnitude

    International Nuclear Information System (INIS)

    Zuppero, A.; Larson, T.K.; Schnitzler, B.G.; Rice, J.W.; Hill, T.J.; Richins, W.D.; Parlier, L.; Werner, J.E.

    1999-01-01

    A brief sketch shows the origin of why and how thermal rocket propulsion has the unique potential to dramatically reduce the cost of space transportation for most inner solar system missions of interest. Orders of magnitude reduction in cost are apparently possible when compared to all processes requiring electrolysis for the production of rocket fuels or propellants and to all electric propulsion systems. An order of magnitude advantage can be attributed to rocket propellant tank factors associated with storing water propellant, compared to cryogenic liquids. An order of magnitude can also be attributed to the simplicity of the extraction and processing of ice on the lunar surface, into an easily stored, non-cryogenic rocket propellant (water). A nuclear heated thermal rocket can deliver thousands of times its mass to Low Earth Orbit from the Lunar surface, providing the equivalent to orders of magnitude drop in launch cost for mass in Earth orbit. Mass includes water ice. These cost reductions depend (exponentially) on the mission delta-v requirements being less than about 6 km/s, or about 3 times the specific velocity of steam rockets (2 km/s, from Isp 200 sec). Such missions include: from the lunar surface to Low Lunar Orbit, (LLO), from LLO to lunar escape, from Low Earth Orbit (LEO) to Geosynchronous Orbit (GEO), from LEO to Earth Escape, from LEO to Mars Transfer Orbit, from LLO to GEO, missions returning payloads from about 10% of the periodic comets using propulsive capture to orbits around Earth itself, and fast, 100 day missions from Lunar Escape to Mars. All the assertions depend entirely and completely on the existence of abundant, nearly pure ice at the permanently dark North and South Poles of the Moon. copyright 1999 American Institute of Physics

  18. 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.

  19. 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

  20. 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

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

    National Research Council Canada - National Science Library

    Shelley, J

    2001-01-01

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

  2. 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…

  3. Evaluation of thermal shock strengths for graphite materials using a laser irradiation method

    International Nuclear Information System (INIS)

    Kim, Jae Hoon; Lee, Young Shin; Kim, Duck Hoi; Park, No Seok; Suh, Jeong; Kim, Jeng O.; Il Moon, Soon

    2004-01-01

    Thermal shock is a physical phenomenon that occurs during the exposure to rapidly high temperature and pressure changes or during quenching of a material. The rocket nozzle throat is exposed to combustion gas of high temperature. Therefore, it is important to select suitable materials having the appropriate thermal shock resistance and to evaluate these materials for rocket nozzle design. The material of this study is ATJ graphite, which is the candidate material for rocket nozzle throat. This study presents an experimental method to evaluate the thermal shock resistance and thermal shock fracture toughness of ATJ graphite using laser irradiation. In particular, thermal shock resistance tests are conducted with changes of specimen thickness, with laser source irradiated at the center of the specimen. Temperature distributions on the specimen surface are detected using type K and C thermocouples. Scanning electron microscope (SEM) is used to observe the thermal cracks on specimen surface

  4. 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.

  5. Laser Shearography Inspection of TPS (Thermal Protection System) Cork on RSRM (Reusable Solid Rocket Motors)

    Science.gov (United States)

    Lingbloom, Mike; Plaia, Jim; Newman, John

    2006-01-01

    Laser Shearography is a viable inspection method for detection of de-bonds and voids within the external TPS (thermal protection system) on to the Space Shuttle RSRM (reusable solid rocket motors). Cork samples with thicknesses up to 1 inch were tested at the LTI (Laser Technology Incorporated) laboratory using vacuum-applied stress in a vacuum chamber. The testing proved that the technology could detect cork to steel un-bonds using vacuum stress techniques in the laboratory environment. The next logical step was to inspect the TPS on a RSRM. Although detailed post flight inspection has confirmed that ATK Thiokol's cork bonding technique provides a reliable cork to case bond, due to the Space Shuttle Columbia incident there is a great interest in verifying bond-lines on the external TPS. This interest provided and opportunity to inspect a RSRM motor with Laser Shearography. This paper will describe the laboratory testing and RSRM testing that has been performed to date. Descriptions of the test equipment setup and techniques for data collection and detailed results will be given. The data from the test show that Laser Shearography is an effective technology and readily adaptable to inspect a RSRM.

  6. 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...

  7. 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.

  8. 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.

  9. 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

  10. 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…

  11. 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.)

  12. Small Rocket/Spacecraft Technology (SMART) Platform

    Science.gov (United States)

    Esper, Jaime; Flatley, Thomas P.; Bull, James B.; Buckley, Steven J.

    2011-01-01

    The NASA Goddard Space Flight Center (GSFC) and the Department of Defense Operationally Responsive Space (ORS) Office are exercising a multi-year collaborative agreement focused on a redefinition of the way space missions are designed and implemented. A much faster, leaner and effective approach to space flight requires the concerted effort of a multi-agency team tasked with developing the building blocks, both programmatically and technologically, to ultimately achieve flights within 7-days from mission call-up. For NASA, rapid mission implementations represent an opportunity to find creative ways for reducing mission life-cycle times with the resulting savings in cost. This in tum enables a class of missions catering to a broader audience of science participants, from universities to private and national laboratory researchers. To that end, the SMART (Small Rocket/Spacecraft Technology) micro-spacecraft prototype demonstrates an advanced avionics system with integrated GPS capability, high-speed plug-and-playable interfaces, legacy interfaces, inertial navigation, a modular reconfigurable structure, tunable thermal technology, and a number of instruments for environmental and optical sensing. Although SMART was first launched inside a sounding rocket, it is designed as a free-flyer.

  13. 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)

  14. 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

  15. Vibrational-rotational temperature measurement of N2 in the lower thermosphere by the rocket experiment

    Science.gov (United States)

    Kurihara, J.; Oyama, K.; Suzuki, K.; Iwagami, N.

    The vibrational temperature (Tv), the rotational temperature (Tr) and the density of atmospheric N2 between 100 - 150 km were measured in situ by a sounding rocket S310-30, over Kagoshima, Japan at 10:30 UT on February 6, 2002. The main purpose of this rocket experiment is to study the dynamics and the thermal energy budget in the lower thermosphere. N2 was ionized using an electron gun and the emission of the 1st negative bands of N2+ was measured by a sensitive spectrometer. Tv and Tr were determined by fitting the observed spectrum for the simulated spectrum, and the number density was deduced from the intensities of the spectrum. We will report preliminary results of our measurement and discuss the observed thermal structure that indicates the effect of tides and gravity waves.

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

    Science.gov (United States)

    Walton, J. T.

    1994-01-01

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

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

    Science.gov (United States)

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

    1995-01-01

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

  18. Combustion diagnosis for analysis of solid propellant rocket abort hazards: Role of spectroscopy

    Science.gov (United States)

    Gill, W.; Cruz-Cabrera, A. A.; Donaldson, A. B.; Lim, J.; Sivathanu, Y.; Bystrom, E.; Haug, A.; Sharp, L.; Surmick, D. M.

    2014-11-01

    Solid rocket propellant plume temperatures have been measured using spectroscopic methods as part of an ongoing effort to specify the thermal-chemical-physical environment in and around a burning fragment of an exploded solid rocket at atmospheric pressures. Such specification is needed for launch safety studies where hazardous payloads become involved with large fragments of burning propellant. The propellant burns in an off-design condition producing a hot gas flame loaded with burning metal droplets. Each component of the flame (soot, droplets and gas) has a characteristic temperature, and it is only through the use of spectroscopy that their temperature can be independently identified.

  19. Combustion diagnosis for analysis of solid propellant rocket abort hazards: Role of spectroscopy

    International Nuclear Information System (INIS)

    Gill, W; Cruz-Cabrera, A A; Bystrom, E; Donaldson, A B; Haug, A; Sharp, L; Lim, J; Sivathanu, Y; Surmick, D M

    2014-01-01

    Solid rocket propellant plume temperatures have been measured using spectroscopic methods as part of an ongoing effort to specify the thermal-chemical-physical environment in and around a burning fragment of an exploded solid rocket at atmospheric pressures. Such specification is needed for launch safety studies where hazardous payloads become involved with large fragments of burning propellant. The propellant burns in an off-design condition producing a hot gas flame loaded with burning metal droplets. Each component of the flame (soot, droplets and gas) has a characteristic temperature, and it is only through the use of spectroscopy that their temperature can be independently identified

  20. Reusable Solid Rocket Motor - Accomplishments, Lessons, and a Culture of Success

    Science.gov (United States)

    Moore, Dennis R.; Phelps, Willie J.

    2011-01-01

    The Reusable Solid Rocket Motor represents the largest solid rocket motor ever flown and the only human rated solid motor. Each Reusable Solid Rocket Motor (RSRM) provides approximately 3-million lb of thrust to lift the integrated Space Shuttle vehicle from the launch pad. The motors burn out approximately 2 minutes later, separate from the vehicle and are recovered and refurbished. The size of the motor and the need for high reliability were challenges. Thrust shaping, via shaping of the propellant grain, was needed to limit structural loads during ascent. The motor design evolved through several block upgrades to increase performance and to increase safety and reliability. A major redesign occurred after STS-51L with the Redesigned Solid Rocket Motor. Significant improvements in the joint sealing systems were added. Design improvements continued throughout the Program via block changes with a number of innovations including development of low temperature o-ring materials and incorporation of a unique carbon fiber rope thermal barrier material. Recovery of the motors and post flight inspection improved understanding of hardware performance, and led to key design improvements. Because of the multidecade program duration material obsolescence was addressed, and requalification of materials and vendors was sometimes needed. Thermal protection systems and ablatives were used to protect the motor cases and nozzle structures. Significant understanding of design and manufacturing features of the ablatives was developed during the program resulting in optimization of design features and processing parameters. The project advanced technology in eliminating ozone-depleting materials in manufacturing processes and the development of an asbestos-free case insulation. Manufacturing processes for the large motor components were unique and safety in the manufacturing environment was a special concern. Transportation and handling approaches were also needed for the large

  1. 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...

  2. Thermal Analysis of the Fastrac Chamber/Nozzle

    Science.gov (United States)

    Davis, Darrell

    2001-01-01

    This paper will describe the thermal analysis techniques used to predict temperatures in the film-cooled ablative rocket nozzle used on the Fastrac 60K rocket engine. A model was developed that predicts char and pyrolysis depths, liner thermal gradients, and temperatures of the bondline between the overwrap and liner. Correlation of the model was accomplished by thermal analog tests performed at Southern Research, and specially instrumented hot fire tests at the Marshall Space Flight Center. Infrared thermography was instrumental in defining nozzle hot wall surface temperatures. In-depth and outboard thermocouple data was used to correlate the kinetic decomposition routine used to predict char and pyrolysis depths. These depths were anchored with measured char and pyrolysis depths from cross-sectioned hot-fire nozzles. For the X-34 flight analysis, the model includes the ablative Thermal Protection System (TPS) material that protects the overwrap from the recirculating plume. Results from model correlation, hot-fire testing, and flight predictions will be discussed.

  3. Liquid Rocket Engine Testing

    Science.gov (United States)

    2016-10-21

    Briefing Charts 3. DATES COVERED (From - To) 17 October 2016 – 26 October 2016 4. TITLE AND SUBTITLE Liquid Rocket Engine Testing 5a. CONTRACT NUMBER...298 (Rev. 8-98) Prescribed by ANSI Std. 239.18 Liquid Rocket Engine Testing SFTE Symposium 21 October 2016 Jake Robertson, Capt USAF AFRL...Distribution Unlimited. PA Clearance 16493 Liquid Rocket Engine Testing • Engines and their components are extensively static-tested in development • This

  4. REIMR - A Process for Utilizing Liquid Rocket Propulsion-Oriented 'Lessons Learned' to Mitigate Development Risk in Nuclear Thermal Propulsion

    International Nuclear Information System (INIS)

    Ballard, Richard O.

    2006-01-01

    This paper is a summary overview of a study conducted at the NASA Marshall Space Flight Center (NASA-MSFC) during the initial phases of the Space Launch Initiative (SLI) program to evaluate a large number of technical problems associated with the design, development, test, evaluation and operation of several major liquid propellant rocket engine systems (i.e., SSME, Fastrac, J-2, F-1). One of the primary results of this study was the identification of the 'Fundamental Root Causes' that enabled the technical problems to manifest, and practices that can be implemented to prevent them from recurring in future propulsion system development efforts, such as that which is currently envisioned in the field of nuclear thermal propulsion (NTP). This paper will discus the Fundamental Root Causes, cite some examples of how the technical problems arose from them, and provide a discussion of how they can be mitigated or avoided in the development of an NTP system

  5. REIMR - A Process for Utilizing Liquid Rocket Propulsion-Oriented 'Lessons Learned' to Mitigate Development Risk in Nuclear Thermal Propulsion

    Science.gov (United States)

    Ballard, RIchard O.

    2006-01-01

    This paper is a summary overview of a study conducted at the NASA Marshall Space Flight Center (NASA MSFC) during the initial phases of the Space Launch Initiative (SLI) program to evaluate a large number of technical problems associated with the design, development, test, evaluation and operation of several major liquid propellant rocket engine systems (i.e., SSME, Fastrac, J-2, F-1). One of the primary results of this study was the identification of the Fundamental Root Causes that enabled the technical problems to manifest, and practices that can be implemented to prevent them from recurring in future propulsion system development efforts, such as that which is currently envisioned in the field of nuclear thermal propulsion (NTF). This paper will discuss the Fundamental Root Causes, cite some examples of how the technical problems arose from them, and provide a discussion of how they can be mitigated or avoided in the development of an NTP system

  6. 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.

  7. 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…

  8. 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.

  9. 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.

  10. Transient thermal analysis of Vega launcher structures

    Energy Technology Data Exchange (ETDEWEB)

    Gori, F. [University of Rome ' Tor Vergata' , Rome (Italy); De Stefanis, M. [Thales Alenia Space Italia, Rome (Italy); Worek, W.M. [University of Illinois at Chicago, Chicago (United States)], E-mail: wworek@uic.edu; Minkowycz, W.J. [University of Illinois at Chicago, Chicago (United States)

    2008-12-15

    A transient thermal analysis is carried out to verify the base cover thermal protection system of Vega 2nd stage Solid Rocket Motor (SRM) and the flange coupling of the inter-stage 2/3. The analysis is performed with a finite element code. The work has developed suitable numerical Fortran subroutines to assign radiation and convection boundary conditions. The thermal behaviour of the structures is presented.

  11. 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.

  12. New method for evaluating the kinetic constant of thermal protection materials

    International Nuclear Information System (INIS)

    Bae, Ji Yeul; Yi, Jong Ju; Park, Sul Ki; Cho, Hyung Hee; Bae, Ju Chan; Ham, Hee Cheol

    2013-01-01

    Thermal protection material (TPM) is used to protect rocket structures from extreme conditions created by the hot exhaust of the rocket. Designing TPM is an important step in the rocket design process. Considering that an increase in the system weight decreases the overall performance of a rocket, the amount of TPM is carefully determined during the design process. Therefore, the precise properties of TPM guarantee an accurate thermal analysis and the successful design of the rocket. Among the many properties of TPM, the kinetic constant and activation energy, which govern the thermochemical reaction of the TPM, are the most important. Thus, an experiment to measure the kinetic constant and activation energy is conducted as part of this research. A theoretical approach to deduce the properties from measured data is discussed, and a method to apply the theory to experimental data, termed the R 2 method, is developed. Compared to a previous method which was difficult to apply, the R 2 method reduces unclear selections of the reaction time and does not require intervention by an interpreter. The properties deduced by the R 2 method show good agreement with the other method despite the limited number of experimental results.

  13. Thermal Analysis of the MC-1 Chamber/Nozzle

    Science.gov (United States)

    Davis, Darrell W.; Phelps, Lisa H. (Technical Monitor)

    2001-01-01

    This paper will describe the thermal analysis techniques used to predict temperatures in the film-cooled ablative rocket nozzle used on the MC-1 60K rocket engine. A model was developed that predicts char and pyrolysis depths, liner thermal gradients, and temperatures of the bondline between the overwrap and liner. Correlation of the model was accomplished by thermal analog tests performed at Southern Research, and specially instrumented hot fire tests at the Marshall Space Flight Center. Infrared thermography was instrumental in defining nozzle hot wall surface temperatures. In-depth and outboard thermocouple data was used to correlate the kinetic decomposition routine used to predict char and pyrolysis depths. These depths were anchored with measured char and pyrolysis depths from cross-sectioned hot-fire nozzles. For the X-34 flight analysis, the model includes the ablative Thermal Protection System (TPS) material that protects the overwrap from the recirculating plume. Results from model correlation, hot-fire testing, and flight predictions will be discussed.

  14. 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)

  15. Thermal-Flow Code for Modeling Gas Dynamics and Heat Transfer in Space Shuttle Solid Rocket Motor Joints

    Science.gov (United States)

    Wang, Qunzhen; Mathias, Edward C.; Heman, Joe R.; Smith, Cory W.

    2000-01-01

    A new, thermal-flow simulation code, called SFLOW. has been developed to model the gas dynamics, heat transfer, as well as O-ring and flow path erosion inside the space shuttle solid rocket motor joints by combining SINDA/Glo, a commercial thermal analyzer. and SHARPO, a general-purpose CFD code developed at Thiokol Propulsion. SHARP was modified so that friction, heat transfer, mass addition, as well as minor losses in one-dimensional flow can be taken into account. The pressure, temperature and velocity of the combustion gas in the leak paths are calculated in SHARP by solving the time-dependent Navier-Stokes equations while the heat conduction in the solid is modeled by SINDA/G. The two codes are coupled by the heat flux at the solid-gas interface. A few test cases are presented and the results from SFLOW agree very well with the exact solutions or experimental data. These cases include Fanno flow where friction is important, Rayleigh flow where heat transfer between gas and solid is important, flow with mass addition due to the erosion of the solid wall, a transient volume venting process, as well as some transient one-dimensional flows with analytical solutions. In addition, SFLOW is applied to model the RSRM nozzle joint 4 subscale hot-flow tests and the predicted pressures, temperatures (both gas and solid), and O-ring erosions agree well with the experimental data. It was also found that the heat transfer between gas and solid has a major effect on the pressures and temperatures of the fill bottles in the RSRM nozzle joint 4 configuration No. 8 test.

  16. 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.

  17. A numerical simulation package for analysis of neutronics and thermal fluids of space nuclear power and propulsion systems

    International Nuclear Information System (INIS)

    Anghaie, S.; Feller, G.J.; Peery, S.D.; Parsley, R.C.

    1993-01-01

    A system of computer codes for engineering simulation and in-depth analysis of nuclear and thermal fluid design of nuclear thermal rockets is developed. The computational system includes a neutronic solver package, a thermal fluid solver package and a propellant and materials property package. The Rocket Engine Transient Simulation (ROCETS) system code is incorporated with computational modules specific to nuclear powered engines. ROCETS features a component based performance architecture that interfaces component modules into the user designed configuration, interprets user commands, creates an executable FORTRAN computer program, and executes the program to provide output to the user. Basic design features of the Pratt ampersand Whitney XNR2000 nuclear rocket concept and its operational performance are analyzed and simulated

  18. Liquid Rocket Engine Testing Overview

    Science.gov (United States)

    Rahman, Shamim

    2005-01-01

    Contents include the following: Objectives and motivation for testing. Technology, Research and Development Test and Evaluation (RDT&E), evolutionary. Representative Liquid Rocket Engine (LRE) test compaigns. Apollo, shuttle, Expandable Launch Vehicles (ELV) propulsion. Overview of test facilities for liquid rocket engines. Boost, upper stage (sea-level and altitude). Statistics (historical) of Liquid Rocket Engine Testing. LOX/LH, LOX/RP, other development. Test project enablers: engineering tools, operations, processes, infrastructure.

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

    International Nuclear Information System (INIS)

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

    1993-01-01

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

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

    Science.gov (United States)

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

    1993-01-01

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

  1. 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.

  2. Real-Time Inhibitor Recession Measurements in Two Space Shuttle Reusable Solid Rocket Motors

    Science.gov (United States)

    McWhorter, B. B.; Ewing, M. E.; Bolton, D. E.; Albrechtsen, K. U.; Earnest, T. E.; Noble, T. C.; Longaker, M.

    2003-01-01

    Real-time internal motor insulation char line recession measurements have been evaluated for two full-scale static tests of the Space Shuttle Reusable Solid Rocket Motor (RSRM). These char line recession measurements were recorded on the forward facing propellant grain inhibitors to better understand the thermal performance of these inhibitors. The RSRM propellant grain inhibitors are designed to erode away during motor operation, thus making it difficult to use post-fire observations to determine inhibitor thermal performance. Therefore, this new internal motor instrumentation is invaluable in establishing an accurate understanding of inhibitor recession versus motor operation time. The data for the first test was presented at the 37th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit (AIAA 2001-3280) in July 2001. Since that time, a second full scale static test has delivered additional real-time data on inhibitor thermal performance. The evaluation of this data is presented in this paper. The second static test, in contrast to the first test, used a slightly different arrangement of instrumentation in the inhibitors. This instrumentation has yielded a better understanding of the inhibitor time dependent inboard tip recession. Graphs of inhibitor recession profiles with time are presented. Inhibitor thermal ablation models have been created from theoretical principals. The model predictions compare favorably with data from both tests. This verified modeling effort is important to support new inhibitor designs for a five segment Space Shuttle solid rocket motor. The internal instrumentation project on RSRM static tests is providing unique opportunities for other real-time internal motor measurements that could not otherwise be directly quantified.

  3. 16 CFR 1507.10 - Rockets with sticks.

    Science.gov (United States)

    2010-01-01

    ... 16 Commercial Practices 2 2010-01-01 2010-01-01 false Rockets with sticks. 1507.10 Section 1507.10... FIREWORKS DEVICES § 1507.10 Rockets with sticks. Rockets with sticks (including skyrockets and bottle rockets) shall utilize a straight and rigid stick to provide a direct and stable flight. Such sticks shall...

  4. 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.

  5. New method for evaluating the kinetic constant of thermal protection materials

    Energy Technology Data Exchange (ETDEWEB)

    Bae, Ji Yeul; Yi, Jong Ju; Park, Sul Ki; Cho, Hyung Hee [Yonsei University, Seoul (Korea, Republic of); Bae, Ju Chan; Ham, Hee Cheol [Agency for Defense Development, Daegu (Korea, Republic of)

    2013-06-15

    Thermal protection material (TPM) is used to protect rocket structures from extreme conditions created by the hot exhaust of the rocket. Designing TPM is an important step in the rocket design process. Considering that an increase in the system weight decreases the overall performance of a rocket, the amount of TPM is carefully determined during the design process. Therefore, the precise properties of TPM guarantee an accurate thermal analysis and the successful design of the rocket. Among the many properties of TPM, the kinetic constant and activation energy, which govern the thermochemical reaction of the TPM, are the most important. Thus, an experiment to measure the kinetic constant and activation energy is conducted as part of this research. A theoretical approach to deduce the properties from measured data is discussed, and a method to apply the theory to experimental data, termed the R{sup 2} method, is developed. Compared to a previous method which was difficult to apply, the R{sup 2} method reduces unclear selections of the reaction time and does not require intervention by an interpreter. The properties deduced by the R{sup 2} method show good agreement with the other method despite the limited number of experimental results.

  6. 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.

  7. Pressure-Equalizing Cradle for Booster Rocket Mounting

    Science.gov (United States)

    Rutan, Elbert L. (Inventor)

    2015-01-01

    A launch system and method improve the launch efficiency of a booster rocket and payload. A launch aircraft atop which the booster rocket is mounted in a cradle, is flown or towed to an elevation at which the booster rocket is released. The cradle provides for reduced structural requirements for the booster rocket by including a compressible layer, that may be provided by a plurality of gas or liquid-filled flexible chambers. The compressible layer contacts the booster rocket along most of the length of the booster rocket to distribute applied pressure, nearly eliminating bending loads. Distributing the pressure eliminates point loading conditions and bending moments that would otherwise be generated in the booster rocket structure during carrying. The chambers may be balloons distributed in rows and columns within the cradle or cylindrical chambers extending along a length of the cradle. The cradle may include a manifold communicating gas between chambers.

  8. Aerodynamics and flow characterisation of multistage rockets

    Science.gov (United States)

    Srinivas, G.; Prakash, M. V. S.

    2017-05-01

    The main objective of this paper is to conduct a systematic flow analysis on single, double and multistage rockets using ANSYS software. Today non-air breathing propulsion is increasing dramatically for the enhancement of space exploration. The rocket propulsion is playing vital role in carrying the payload to the destination. Day to day rocket aerodynamic performance and flow characterization analysis has becoming challenging task to the researchers. Taking this task as motivation a systematic literature is conducted to achieve better aerodynamic and flow characterization on various rocket models. The analyses on rocket models are very little especially in numerical side and experimental area. Each rocket stage analysis conducted for different Mach numbers and having different flow varying angle of attacks for finding the critical efficiency performance parameters like pressure, density and velocity. After successful completion of the analysis the research reveals that flow around the rocket body for Mach number 4 and 5 best suitable for designed payload. Another major objective of this paper is to bring best aerodynamics flow characterizations in both aero and mechanical features. This paper also brings feature prospectus of rocket stage technology in the field of aerodynamic design.

  9. 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

  10. Production of high-energy substances in the process of thorough treatment of hard rocket fuel

    International Nuclear Information System (INIS)

    Shiman, L.N.; Sobolev, V.V.

    2010-01-01

    The high-energy products taken from hard rocket fuel are explored on the sensitivity to a shock, friction, electrostatic discharge, detonation impulse, vibroloads, and capsule-detonator activity. Their chemical stability, thermal stability, and other physicochemical parameters are studied. The field of a recurring effective utilization of gained octogene, ammonium, and potassium perchlorate is shown.

  11. Radiometric probe design for the measurement of heat flux within a solid rocket motor nozzle

    Science.gov (United States)

    Goldey, Charles L.; Laughlin, William T.; Popper, Leslie A.

    1996-11-01

    Improvements to solid rocket motor (SRM) nozzle designs and material performance is based on the ability to instrument motors during test firings to understand the internal combustion processes and the response of nozzle components to the severe heating environment. Measuring the desired parameters is very difficult because the environment inside of an SRM is extremely severe. Instrumentation can be quickly destroyed if exposed to the internal rocket motor environment. An optical method is under development to quantify the heating of the internal nozzle surface. A radiometric probe designed for measuring the thermal response and material surface recession within a nozzle while simultaneously confining the combustion products has been devised and demonstrated. As part of the probe design, optical fibers lead to calibrated detectors that measure the interior nozzle thermal response. This two color radiometric measurement can be used for a direct determination of the total heat flux impinging on interior nozzle surfaces. This measurement has been demonstrated using a high power CO2 laser to simulate SRM nozzle heating conditions on carbon phenolic and graphite phenolic materials.

  12. The Swedish sounding rocket programme

    International Nuclear Information System (INIS)

    Bostroem, R.

    1980-01-01

    Within the Swedish Sounding Rocket Program the scientific groups perform experimental studies of magnetospheric and ionospheric physics, upper atmosphere physics, astrophysics, and material sciences in zero g. New projects are planned for studies of auroral electrodynamics using high altitude rockets, investigations of noctilucent clouds, and active release experiments. These will require increased technical capabilities with respect to payload design, rocket performance and ground support as compared with the current program. Coordination with EISCAT and the planned Viking satellite is essential for the future projects. (Auth.)

  13. On thermal conductivity of gas mixtures containing hydrogen

    Science.gov (United States)

    Zhukov, Victor P.; Pätz, Markus

    2017-06-01

    A brief review of formulas used for the thermal conductivity of gas mixtures in CFD simulations of rocket combustion chambers is carried out in the present work. In most cases, the transport properties of mixtures are calculated from the properties of individual components using special mixing rules. The analysis of different mixing rules starts from basic equations and ends by very complex semi-empirical expressions. The formulas for the thermal conductivity are taken for the analysis from the works on modelling of rocket combustion chambers. \\hbox {H}_2{-}\\hbox {O}_2 mixtures are chosen for the evaluation of the accuracy of the considered mixing rules. The analysis shows that two of them, of Mathur et al. (Mol Phys 12(6):569-579, 1967), and of Mason and Saxena (Phys Fluids 1(5):361-369, 1958), have better agreement with the experimental data than other equations for the thermal conductivity of multicomponent gas mixtures.

  14. 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…

  15. 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

  16. Sounding rockets explore the ionosphere

    International Nuclear Information System (INIS)

    Mendillo, M.

    1990-01-01

    It is suggested that small, expendable, solid-fuel rockets used to explore ionospheric plasma can offer insight into all the processes and complexities common to space plasma. NASA's sounding rocket program for ionospheric research focuses on the flight of instruments to measure parameters governing the natural state of the ionosphere. Parameters include input functions, such as photons, particles, and composition of the neutral atmosphere; resultant structures, such as electron and ion densities, temperatures and drifts; and emerging signals such as photons and electric and magnetic fields. Systematic study of the aurora is also conducted by these rockets, allowing sampling at relatively high spatial and temporal rates as well as investigation of parameters, such as energetic particle fluxes, not accessible to ground based systems. Recent active experiments in the ionosphere are discussed, and future sounding rocket missions are cited

  17. On use of hybrid rocket propulsion for suborbital vehicles

    Science.gov (United States)

    Okninski, Adam

    2018-04-01

    While the majority of operating suborbital rockets use solid rocket propulsion, recent advancements in the field of hybrid rocket motors lead to renewed interest in their use in sounding rockets. This paper presents results of optimisation of sounding rockets using hybrid propulsion. An overview of vehicles under development during the last decade, as well as heritage systems is provided. Different propellant combinations are discussed and their performance assessment is given. While Liquid Oxygen, Nitrous Oxide and Nitric Acid have been widely tested with various solid fuels in flight, Hydrogen Peroxide remains an oxidiser with very limited sounding rocket applications. The benefits of hybrid propulsion for sounding rockets are given. In case of hybrid rocket motors the thrust curve can be optimised for each flight, using a flow regulator, depending on the payload and mission. Results of studies concerning the optimal burn duration and nozzle selection are given. Specific considerations are provided for the Polish ILR-33 "Amber" sounding rocket. Low regression rates, which up to date were viewed as a drawback of hybrid propulsion may be used to the benefit of maximising rocket performance if small solid rocket boosters are used during the initial flight period. While increased interest in hybrid propulsion is present, no up-to-date reference concerning use of hybrid rocket propulsion for sounding rockets is available. The ultimate goal of the paper is to provide insight into the sensitivity of different design parameters on performance of hybrid sounding rockets and delve into the potential and challenges of using hybrid rocket technology for expendable suborbital applications.

  18. Rocket center Peenemünde — Personal memories

    Science.gov (United States)

    Dannenberg, Konrad; Stuhlinger, Ernst

    Von Braun built his first rockets as a young teenager. At 14, he started making plans for rockets for human travel to the Moon and Mars. The German Army began a rocket program in 1929. Two years later, Colonel (later General) Becker contacted von Braun who experimented with rockets in Berlin, gave him a contract in 1932, and, jointly with the Air Force, in 1936 built the rocket center Peenemünde where von Braun and his team developed the A-4 (V-2) rocket under Army auspices, while the Air Force developed the V-1 (buzz bomb), wire-guided bombs, and rocket planes. Albert Speer, impressed by the work of the rocketeers, allowed a modest growth of the Peenemünde project; this brought Dannenberg to the von Braun team in 1940. Hitler did not believe in rockets; he ignored the A-4 project until 1942 when he began to support it, expecting that it could turn the fortunes of war for him. He drastically increased the Peenemünde work force and allowed the transfer of soldiers from the front to Peenemünde; that was when Stuhlinger, in 1943, came to Peenemünde as a Pfc.-Ph.D. Later that year, Himmler wrenched the authority over A-4 production out of the Army's hands, put it under his command, and forced production of the immature rocket at Mittelwerk, and its military deployment against targets in France, Belgium, and England. Throughout the development of the A-4 rocket, von Braun was the undisputed leader of the project. Although still immature by the end of the war, the A-4 had proceeded to a status which made it the first successful long-range precision rocket, the prototype for a large number of military rockets built by numerous nations after the war, and for space rockets that launched satellites and traveled to the Moon and the planets.

  19. 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

  20. 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 ...

  1. This Is Rocket Science!

    Science.gov (United States)

    Keith, Wayne; Martin, Cynthia; Veltkamp, Pamela

    2013-09-01

    Using model rockets to teach physics can be an effective way to engage students in learning. In this paper, we present a curriculum developed in response to an expressed need for helping high school students review physics equations in preparation for a state-mandated exam. This required a mode of teaching that was more advanced and analytical than that offered by Estes Industries, but more basic than the analysis of Nelson et al. In particular, drag is neglected until the very end of the exercise, which allows the concept of conservation of energy to be shown when predicting the rocket's flight. Also, the variable mass of the rocket motor is assumed to decrease linearly during the flight (while the propulsion charge and recovery delay charge are burning) and handled simplistically by using an average mass value. These changes greatly simplify the equations needed to predict the times and heights at various stages of flight, making it more useful as a review of basic physics. Details about model rocket motors, range safety, and other supplemental information may be found online at Apogee Components4 and the National Association of Rocketry.5

  2. Photometric observations of local rocket-atmosphere interactions

    Science.gov (United States)

    Greer, R. G. H.; Murtagh, D. P.; Witt, G.; Stegman, J.

    1983-06-01

    Photometric measurements from rocket flights which recorded a strong foreign luminance in the altitude region between 90 and 130 km are reported. From one Nike-Orion rocket the luminance appeared on both up-leg and down-leg; from a series of Petrel rockets the luminance was apparent only on the down-leg. The data suggest that the luminance may be distributed mainly in the wake region along the rocket trajectory. The luminance is believed to be due to a local interaction between the rocket and the atmosphere although the precise nature of the interaction is unknown. It was measured at wavelengths ranging from 275 nm to 1.61 microns and may be caused by a combination of reactions.

  3. 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

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

    Science.gov (United States)

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

    2017-12-01

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

  5. Rocket Flight Path

    Directory of Open Access Journals (Sweden)

    Jamie Waters

    2014-09-01

    Full Text Available This project uses Newton’s Second Law of Motion, Euler’s method, basic physics, and basic calculus to model the flight path of a rocket. From this, one can find the height and velocity at any point from launch to the maximum altitude, or apogee. This can then be compared to the actual values to see if the method of estimation is a plausible. The rocket used for this project is modeled after Bullistic-1 which was launched by the Society of Aeronautics and Rocketry at the University of South Florida.

  6. 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

  7. The Isinglass Auroral Sounding Rocket Campaign: data synthesis incorporating remote sensing, in situ observations, and modelling

    Science.gov (United States)

    Lynch, K. A.; Clayton, R.; Roberts, T. M.; Hampton, D. L.; Conde, M.; Zettergren, M. D.; Burleigh, M.; Samara, M.; Michell, R.; Grubbs, G. A., II; Lessard, M.; Hysell, D. L.; Varney, R. H.; Reimer, A.

    2017-12-01

    The NASA auroral sounding rocket mission Isinglass was launched from Poker Flat Alaska in winter 2017. This mission consists of two separate multi-payload sounding rockets, over an array of groundbased observations, including radars and filtered cameras. The science goal is to collect two case studies, in two different auroral events, of the gradient scale sizes of auroral disturbances in the ionosphere. Data from the in situ payloads and the groundbased observations will be synthesized and fed into an ionospheric model, and the results will be studied to learn about which scale sizes of ionospheric structuring have significance for magnetosphere-ionosphere auroral coupling. The in situ instrumentation includes thermal ion sensors (at 5 points on the second flight), thermal electron sensors (at 2 points), DC magnetic fields (2 point), DC electric fields (one point, plus the 4 low-resource thermal ion RPA observations of drift on the second flight), and an auroral precipitation sensor (one point). The groundbased array includes filtered auroral imagers, the PFISR and SuperDarn radars, a coherent scatter radar, and a Fabry-Perot interferometer array. The ionospheric model to be used is a 3d electrostatic model including the effects of ionospheric chemistry. One observational and modelling goal for the mission is to move both observations and models of auroral arc systems into the third (along-arc) dimension. Modern assimilative tools combined with multipoint but low-resource observations allow a new view of the auroral ionosphere, that should allow us to learn more about the auroral zone as a coupled system. Conjugate case studies such as the Isinglass rocket flights allow for a test of the models' intepretation by comparing to in situ data. We aim to develop and improve ionospheric models to the point where they can be used to interpret remote sensing data with confidence without the checkpoint of in situ comparison.

  8. Nuclear Thermal Rocket (NTR) Development Risk Communication

    Science.gov (United States)

    Kim, Tony

    2014-01-01

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

  9. Wake effect in rocket observation

    International Nuclear Information System (INIS)

    Matsumoto, Haruya; Kaya, Nobuyuki; Yamanaka, Akira; Hayashi, Tomomasa

    1975-01-01

    The mechanism of the wake phenomena due to a probe and in rocket observation is discussed on the basis of experimental data. In the low energy electron measurement performed with the L-3H-5 rocket, the electron count rate changed synchronously with the rocket spin. This seems to be a wake effect. It is also conceivable that the probe itself generates the wake of ion beam. The latter problem is considered in the first part. Experiment was performed with laboratory plasma, in which a portion of the electron component of the probe current was counted with a CEM (a channel type multiplier). The change of probe voltage-count rate charactersitics due to the change of relative position of the ion source was observed. From the measured angular distributions of electron density and electron temperature around the probe, it is concluded that anisotropy exists around the probe, which seems to be a kinds of wake structure. In the second part, the wake effect due to a rocket is discussed on the basis of the measurement of leaking electrons with L-3H-5 rocket. Comparison between the theory of wake formation and the measured results is also shortly made in the final part. (Aoki, K.)

  10. Nitrous Oxide/Paraffin Hybrid Rocket Engines

    Science.gov (United States)

    Zubrin, Robert; Snyder, Gary

    2010-01-01

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

  11. Oxygen Containment System Options for Nuclear Thermal Propulsion Testing

    Data.gov (United States)

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

  12. A Flight Demonstration of Plasma Rocket Propulsion

    Science.gov (United States)

    Petro, Andrew

    1999-01-01

    The Advanced Space Propulsion Laboratory at the Johnson Space Center has been engaged in the development of a magneto-plasma rocket for several years. This type of rocket could be used in the future to propel interplanetary spacecraft. One advantageous feature of this rocket concept is the ability to vary its specific impulse so that it can be operated in a mode which maximizes propellant efficiency or a mode which maximizes thrust. This presentation will describe a proposed flight experiment in which a simple version of the rocket will be tested in space. In addition to the plasma rocket, the flight experiment will also demonstrate the use of a superconducting electromagnet, extensive use of heat pipes, and possibly the transfer of cryogenic propellant in space.

  13. 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.

  14. 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

  15. Performances Study of a Hybrid Rocket Engine

    Directory of Open Access Journals (Sweden)

    Adrian-Nicolae BUTURACHE

    2018-06-01

    Full Text Available This paper presents a study which analyses the functioning and performances optimization of a hybrid rocket engine based on gaseous oxygen and polybutadiene polymer (HTPB. Calculations were performed with NASA CEA software in order to obtain the parameters resulted following the combustion process. Using these parameters, the main parameters of the hybrid rocket engine were optimized. Using the calculus previously stated, an experimental rocket engine producing 100 N of thrust was pre-dimensioned, followed by an optimization of the rocket engine as a function of several parameters. Having the geometry and the main parameters of the hybrid rocket engine combustion process, numerical simulations were performed in the CFX – ANSYS commercial software, which allowed visualizing the flow field and the jet expansion. Finally, the analytical calculus was validated through numerical simulations.

  16. Rocket Ozone Data Recovery for Digital Archival

    Science.gov (United States)

    Hwang, S. H.; Krueger, A. J.; Hilsenrath, E.; Haffner, D. P.; Bhartia, P. K.

    2014-12-01

    Ozone distributions in the photochemically-controlled upper stratosphere and mesosphere were first measured using spectrometers on V-2 rockets after WWII. The IGY(1957-1958) spurred development of new optical and chemical instruments for flight on meteorological and sounding rockets. In the early 1960's, the US Navy developed an Arcas rocket-borne optical ozonesonde and NASA GSFC developed chemiluminescent ozonesonde onboard Nike_Cajun and Arcas rocket. The Navy optical ozone program was moved in 1969 to GSFC where rocket ozone research was expanded and continued until 1994 using Super Loki-Dart rocket at 11 sites in the range of 0-65N and 35W-160W. Over 300 optical ozone soundings and 40 chemiluminescent soundings were made. The data have been used to produce the US Standard Ozone Atmosphere, determine seasonal and diurnal variations, and validate early photochemical models. The current effort includes soundings conducted by Australia, Japan, and Korea using optical techniques. New satellite ozone sounding techniques were initially calibrated and later validated using the rocket ozone data. As satellite techniques superseded the rocket methods, the sponsoring agencies lost interest in the data and many of those records have been discarded. The current task intends to recover as much of the data as possible from the private records of the experimenters and their publications, and to archive those records in the WOUDC (World Ozone and Ultraviolet Data Centre). The original data records are handwritten tabulations, computer printouts that are scanned with OCR techniques, and plots digitized from publications. This newly recovered digital rocket ozone profile data from 1965 to 2002 could make significant contributions to the Earth science community in atmospheric research including long-term trend analysis.

  17. 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.

  18. 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

  19. Numerical simulations of a sounding rocket in ionospheric plasma: Effects of magnetic field on the wake formation and rocket potential

    Science.gov (United States)

    Darian, D.; Marholm, S.; Paulsson, J. J. P.; Miyake, Y.; Usui, H.; Mortensen, M.; Miloch, W. J.

    2017-09-01

    The charging of a sounding rocket in subsonic and supersonic plasma flows with external magnetic field is studied with numerical particle-in-cell (PIC) simulations. A weakly magnetized plasma regime is considered that corresponds to the ionospheric F2 layer, with electrons being strongly magnetized, while the magnetization of ions is weak. It is demonstrated that the magnetic field orientation influences the floating potential of the rocket and that with increasing angle between the rocket axis and the magnetic field direction the rocket potential becomes less negative. External magnetic field gives rise to asymmetric wake downstream of the rocket. The simulated wake in the potential and density may extend as far as 30 electron Debye lengths; thus, it is important to account for these plasma perturbations when analyzing in situ measurements. A qualitative agreement between simulation results and the actual measurements with a sounding rocket is also shown.

  20. Plume Particle Collection and Sizing from Static Firing of Solid Rocket Motors

    Science.gov (United States)

    Sambamurthi, Jay K.

    1995-01-01

    Thermal radiation from the plume of any solid rocket motor, containing aluminum as one of the propellant ingredients, is mainly from the microscopic, hot aluminum oxide particles in the plume. The plume radiation to the base components of the flight vehicle is primarily determined by the plume flowfield properties, the size distribution of the plume particles, and their optical properties. The optimum design of a vehicle base thermal protection system is dependent on the ability to accurately predict this intense thermal radiation using validated theoretical models. This article describes a successful effort to collect reasonably clean plume particle samples from the static firing of the flight simulation motor (FSM-4) on March 10, 1994 at the T-24 test bed at the Thiokol space operations facility as well as three 18.3% scaled MNASA motors tested at NASA/MSFC. Prior attempts to collect plume particles from the full-scale motor firings have been unsuccessful due to the extremely hostile thermal and acoustic environment in the vicinity of the motor nozzle.

  1. Plasma waves observed by sounding rockets

    International Nuclear Information System (INIS)

    Kimura, I.

    1977-01-01

    Observations of plasma wave phenomena have been conducted with several rockets launched at Kagoshima Space Center, Kyushu, Japan, and at Showa Base, Antarctica. This report presents some results of the observations in anticipation of having valuable comments from other plasma physicists, especially from those who are concerned with laboratory plasma. In the K-9M-41 rocket experiment, VLF plasma waves were observed. In this experiment, the electron beam of several tens of uA was emitted from a hot cathode when a positive dc bias changing from 0 to 10V at 1V interval each second was applied to a receiving dipole antenna. The discrete emissions with 'U' shaped frequency spectrum were observed for the dc bias over 3 volts. The U emissions appeared twice per spin period of the rocket. Similar rocket experiment was performed at Showa Base using a loop and dipole antenna and without hot cathode. Emissions were observed with varying conditions. At present, the authors postulate that such emissions may be produced just in the vicinity of a rocket due to a kind of wake effect. (Aoki, K.)

  2. THE ADIABATIC DEMAGNETIZATION REFRIGERATOR FOR THE MICRO-X SOUNDING ROCKET TELESCOPE

    International Nuclear Information System (INIS)

    Wikus, P.; Bagdasarova, Y.; Figueroa-Feliciano, E.; Leman, S. W.; Rutherford, J. M.; Trowbridge, S. N.; Adams, J. S.; Bandler, S. R.; Eckart, M. E.; Kelley, R. L.; Kilbourne, C. A.; Porter, F. S.; Doriese, W. B.; McCammon, D.

    2010-01-01

    The Micro-X Imaging X-ray Spectrometer is a sounding rocket payload slated for launch in 2011. An array of Transition Edge Sensors, which is operated at a bath temperature of 50 mK, will be used to obtain a high resolution spectrum of the Puppis-A supernova remnant. An Adiabatic Demagnetization Refrigerator (ADR) with a 75 gram Ferric Ammonium Alum (FAA) salt pill in the bore of a 4 T superconducting magnet provides a stable heat sink for the detector array only a few seconds after burnout of the rocket motors. This requires a cold stage design with very short thermal time constants. A suspension made from Kevlar strings holds the 255 gram cold stage in place. It is capable of withstanding loads in excess of 200 g. Stable operation of the TES array in proximity to the ADR magnet is ensured by a three-stage magnetic shielding system which consists of a superconducting can, a high-permeability shield and a bucking coil. The development and testing of the Micro-X payload is well underway.

  3. The Spanish national programme of balloons and sounding rockets

    International Nuclear Information System (INIS)

    Casas, J.; Pueyo, L.

    1978-01-01

    The main points of the Spanish scientific programme are briefly described: CONIE/NASA cooperative project on meteorological sounding rocket launchings; ozonospheric programme; CONIE/NASA/CNES cooperative ionospheric sounding rocket project; D-layer research; rocket infrared dayglow measurements; ultraviolet astronomy research; cosmic ray research. The schedule of sounding rocket launchings at El Arenosillo station during 1977 is given

  4. Space Shuttle Redesigned Solid Rocket Motor nozzle natural frequency variations with burn time

    Science.gov (United States)

    Lui, C. Y.; Mason, D. R.

    1991-01-01

    The effects of erosion and thermal degradation on the Space Shuttle Redesigned Solid Rocket Motor (RSRM) nozzle's structural dynamic characteristics were analytically evaluated. Also considered was stiffening of the structure due to internal pressurization. A detailed NASTRAN finite element model of the nozzle was developed and used to evaluate the influence of these effects at several discrete times during motor burn. Methods were developed for treating erosion and thermal degradation, and a procedure was developed to account for internal pressure stiffening using differential stiffness matrix techniques. Results were verified using static firing test accelerometer data. Fast Fourier Transform and Maximum Entropy Method techniques were applied to the data to generate waterfall plots which track modal frequencies with burn time. Results indicate that the lower frequency nozzle 'vectoring' modes are only slightly affected by erosion, thermal effects and internal pressurization. The higher frequency shell modes of the nozzle are, however, significantly reduced.

  5. 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.

  6. RX LAPAN Rocket data Program With Dbase III Plus

    International Nuclear Information System (INIS)

    Sauman

    2001-01-01

    The components data rocket RX LAPAN are taken from workshop product and assembling rocket RX. In this application software, the test data are organized into two data files, i.e. test file and rocket file. Besides [providing facilities to add, edit and delete data, this software provides also data manipulation facility to support analysis and identification of rocket RX failures and success

  7. Fracture Characteristics of C/SiC Composites for Rocket Nozzle at Elevated Temperature

    Energy Technology Data Exchange (ETDEWEB)

    Yoon, Dong Hyun; Lee, Jeong Won; Kim, Jae Hoon [Chungnam Nat’l Univ., Daejeon (Korea, Republic of); Sihn, Ihn Cheol; Lim, Byung Joo [Dai-Yang Industries Co., Daejeon (Korea, Republic of)

    2016-11-15

    In a solid propulsion system, the rocket nozzle is exposed to high temperature combustion gas. Hence, choosing an appropriate material that could demonstrate adequate performance at high temperature is important. As advanced materials, carbon/silicon carbide composites (C/SiC) have been studied with the aim of using them for the rocket nozzle throat. However, when compared with typical structural materials, C/SiC composites are relatively weak in terms of both strength and toughness, owing to their quasi-brittle behavior and oxidation at high temperatures. Therefore, it is important to evaluate the thermal and mechanical properties of this material before using it in this application. This study presents an experimental method to investigate the fracture behavior of C/SiC composite material manufactured using liquid silicon infiltration (LSI) method at elevated temperatures. In particular, the effects of major parameters, such as temperature, loading, oxidation conditions, and fiber direction on strength and fracture characteristics were investigated. Fractography analysis of the fractured specimens was performed using an SEM.

  8. The Norwegian sounding rocket programme 1978-81

    International Nuclear Information System (INIS)

    Landmark, B.

    1978-01-01

    The Norwegian sounding rocket programme is reasonably well defined up to and including the winter of 1981/82. All the projects have been planned and will be carried out in international cooperation. Norwegian scientists so far plan to participate in a number of 24 rocket payloads over the period. Out of these 18 will be launched from the Andoya rocket range, 3 from Esrange and 3 from the siple station in the antarctic. (author)

  9. 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.

  10. Assessment of exposure-response functions for rocket-emission toxicants

    National Research Council Canada - National Science Library

    Subcommittee on Rocket-Emission Toxicants, National Research Council

    ... aborted launch that results in a rocket being destroyed near the ground. Assessment of Exposure-Response Functions for Rocket-Emmission Toxicants evaluates the model and the data used for three rocket emission toxicants...

  11. '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

  12. Integrated Composite Rocket Nozzle Extension, Phase I

    Data.gov (United States)

    National Aeronautics and Space Administration — ORBITEC proposes to develop and demonstrate an Integrated Composite Rocket Nozzle Extension (ICRNE) for use in rocket thrust chambers. The ICRNE will utilize an...

  13. 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.

  14. Thermal-hydraulics Analysis of a Radioisotope-powered Mars Hopper Propulsion System

    International Nuclear Information System (INIS)

    O'Brien, Robert C.; Klein, Andrew C.; Taitano, William T.; Gibson, Justice; Myers, Brian; Howe, Steven D.

    2011-01-01

    Thermal-hydraulics analyses results produced using a combined suite of computational design and analysis codes are presented for the preliminary design of a concept Radioisotope Thermal Rocket (RTR) propulsion system. Modeling of the transient heating and steady state temperatures of the system is presented. Simulation results for propellant blow down during impulsive operation are also presented. The results from this study validate the feasibility of a practical thermally capacitive RTR propulsion system.

  15. 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

  16. Flow-Structural Interaction in Solid Rocket Motors

    National Research Council Canada - National Science Library

    Murdock, John

    2004-01-01

    .... The static test failure of the Titan solid rocket motor upgrade (SRMU) that occurred on 1 April, 1991, demonstrated the importance of flow-structural modeling in the design of large, solid rocket motors...

  17. 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.

  18. Rocket measurements of electron density irregularities during MAC/SINE

    Science.gov (United States)

    Ulwick, J. C.

    1989-01-01

    Four Super Arcas rockets were launched at the Andoya Rocket Range, Norway, as part of the MAC/SINE campaign to measure electron density irregularities with high spatial resolution in the cold summer polar mesosphere. They were launched as part of two salvos: the turbulent/gravity wave salvo (3 rockets) and the EISCAT/SOUSY radar salvo (one rocket). In both salvos meteorological rockets, measuring temperature and winds, were also launched and the SOUSY radar, located near the launch site, measured mesospheric turbulence. Electron density irregularities and strong gradients were measured by the rocket probes in the region of most intense backscatter observed by the radar. The electron density profiles (8 to 4 on ascent and 4 on descent) show very different characteristics in the peak scattering region and show marked spatial and temporal variability. These data are intercompared and discussed.

  19. Design criteria of launching rockets for burst aerial shells

    Energy Technology Data Exchange (ETDEWEB)

    Kuwahara, T.; Takishita, Y.; Onda, T.; Shibamoto, H.; Hosaya, F. [Hosaya Kako Co. Ltd (Japan); Kubota, N. [Mitsubishi Electric Corporation (Japan)

    2000-04-01

    Rocket motors attached to large-sized aerial shells are proposed to compensate for the increase in the lifting charge in the mortar and the thickness of the shell wall. The proposal is the result of an evaluation of the performance of solid propellants to provide information useful in designing launch rockets for large-size shells. The propellants composed of ammonium perchlorate and hydroxy-terminated polybutadiene were used to evaluate the ballistic characteristics such as the relationship between propellant mass and trajectories of shells and launch rockets. In order to obtain an optimum rocket design, the evaluation also included a study of the velocity and height of the rocket motor and shell separation. A launch rocket with a large-sized shell (84.5 cm in diameter) was designed to verify the effectiveness of this class of launch system. 2 refs., 6 figs.

  20. Rocket Science at the Nanoscale.

    Science.gov (United States)

    Li, Jinxing; Rozen, Isaac; Wang, Joseph

    2016-06-28

    Autonomous propulsion at the nanoscale represents one of the most challenging and demanding goals in nanotechnology. Over the past decade, numerous important advances in nanotechnology and material science have contributed to the creation of powerful self-propelled micro/nanomotors. In particular, micro- and nanoscale rockets (MNRs) offer impressive capabilities, including remarkable speeds, large cargo-towing forces, precise motion controls, and dynamic self-assembly, which have paved the way for designing multifunctional and intelligent nanoscale machines. These multipurpose nanoscale shuttles can propel and function in complex real-life media, actively transporting and releasing therapeutic payloads and remediation agents for diverse biomedical and environmental applications. This review discusses the challenges of designing efficient MNRs and presents an overview of their propulsion behavior, fabrication methods, potential rocket fuels, navigation strategies, practical applications, and the future prospects of rocket science and technology at the nanoscale.

  1. Infrared Imagery of Solid Rocket Exhaust Plumes

    Science.gov (United States)

    Moran, Robert P.; Houston, Janice D.

    2011-01-01

    The Ares I Scale Model Acoustic Test program consisted of a series of 18 solid rocket motor static firings, simulating the liftoff conditions of the Ares I five-segment Reusable Solid Rocket Motor Vehicle. Primary test objectives included acquiring acoustic and pressure data which will be used to validate analytical models for the prediction of Ares 1 liftoff acoustics and ignition overpressure environments. The test article consisted of a 5% scale Ares I vehicle and launch tower mounted on the Mobile Launch Pad. The testing also incorporated several Water Sound Suppression Systems. Infrared imagery was employed during the solid rocket testing to support the validation or improvement of analytical models, and identify corollaries between rocket plume size or shape and the accompanying measured level of noise suppression obtained by water sound suppression systems.

  2. The microwave thermal thruster and its application to the launch problem

    Science.gov (United States)

    Parkin, Kevin L. G.

    Nuclear thermal thrusters long ago bypassed the 50-year-old specific impulse (Isp) limitation of conventional thrusters, using nuclear powered heat exchangers in place of conventional combustion to heat a hydrogen propellant. These heat exchanger thrusters experimentally achieved an Isp of 825 seconds, but with a thrust-to-weight ratio (T/W) of less than ten they have thus far been too heavy to propel rockets into orbit. This thesis proposes a new idea to achieve both high Isp and high T/W The Microwave Thermal Thruster. This thruster covers the underside of a rocket aeroshell with a lightweight microwave absorbent heat exchange layer that may double as a re-entry heat shield. By illuminating the layer with microwaves directed from a ground-based phased array, an Isp of 700--900 seconds and T/W of 50--150 is possible using a hydrogen propellant. The single propellant simplifies vehicle design, and the high Isp increases payload fraction and structural margins. These factors combined could have a profound effect on the economics of building and reusing rockets. A laboratory-scale microwave thermal heat exchanger is constructed using a single channel in a cylindrical microwave resonant cavity, and new type of coupled electromagnetic-conduction-convection model is developed to simulate it. The resonant cavity approach to small-scale testing reveals several drawbacks, including an unexpected oscillatory behavior. Stable operation of the laboratory-scale thruster is nevertheless successful, and the simulations are consistent with the experimental results. In addition to proposing a new type of propulsion and demonstrating it, this thesis provides three other principal contributions: The first is a new perspective on the launch problem, placing it in a wider economic context. The second is a new type of ascent trajectory that significantly reduces the diameter, and hence cost, of the ground-based phased array. The third is an eclectic collection of data, techniques, and

  3. Analysis of rocket flight stability based on optical image measurement

    Science.gov (United States)

    Cui, Shuhua; Liu, Junhu; Shen, Si; Wang, Min; Liu, Jun

    2018-02-01

    Based on the abundant optical image measurement data from the optical measurement information, this paper puts forward the method of evaluating the rocket flight stability performance by using the measurement data of the characteristics of the carrier rocket in imaging. On the basis of the method of measuring the characteristics of the carrier rocket, the attitude parameters of the rocket body in the coordinate system are calculated by using the measurements data of multiple high-speed television sets, and then the parameters are transferred to the rocket body attack angle and it is assessed whether the rocket has a good flight stability flying with a small attack angle. The measurement method and the mathematical algorithm steps through the data processing test, where you can intuitively observe the rocket flight stability state, and also can visually identify the guidance system or failure analysis.

  4. Rocket experiment METS - Microwave Energy Transmission in Space

    Science.gov (United States)

    Kaya, N.; Matsumoto, H.; Akiba, R.

    A Microwave Energy Transmission in Space (METS) rocket experiment is being planned by the Solar Power Satellite Working Group at the Institute of Space and Astronautical Science in Japan for the forthcoming International Space Year, 1992. The METS experiment is an advanced version of the previous MINIX rocket experiment (Matsumoto et al., 1990). This paper describes a conceptual design of the METS rocket experiment. It aims at verifying a newly developed microwave energy transmission system for space use and to study nonlinear effects of the microwave energy beam in the space plasma environment. A high power microwave of 936 W will be transmitted by the new phased-array antenna from a mother rocket to a separated target (daughter rocket) through the ionospheric plasma. The active phased-array system has a capability of focusing the microwave energy around any spatial point by controlling the digital phase shifters individually.

  5. Rocket experiment METS Microwave Energy Transmission in Space

    Science.gov (United States)

    Kaya, N.; Matsumoto, H.; Akiba, R.

    A METS (Microwave Energy Transmission in Space) rocket experiment is being planned by the SPS (Solar Power Satellite) Working Group at the Institute of Space and Astronautical Science (ISAS) in Japan for the forthcoming International Space Year (ISY), 1992. The METS experiment is an advanced version of our MINIX rocket experiment. This paper describes the conceptual design for the METS rocket experiment. Aims are to verify the feasibility of a newly developed microwave energy transmission system designed for use in space and to study nonlinear effects of the microwave energy beam on space plasma. A high power microwave (936 W) will be transmitted by a new phase-array antenna from a mother rocket to a separate target (daughter rocket) through the Earth's ionospheric plasma. The active phased-array system has the capability of being able to focus the microwave energy at any spatial point by individually controlling the digital phase shifters.

  6. Adaptive Time Stepping for Transient Network Flow Simulation in Rocket Propulsion Systems

    Science.gov (United States)

    Majumdar, Alok K.; Ravindran, S. S.

    2017-01-01

    Fluid and thermal transients found in rocket propulsion systems such as propellant feedline system is a complex process involving fast phases followed by slow phases. Therefore their time accurate computation requires use of short time step initially followed by the use of much larger time step. Yet there are instances that involve fast-slow-fast phases. In this paper, we present a feedback control based adaptive time stepping algorithm, and discuss its use in network flow simulation of fluid and thermal transients. The time step is automatically controlled during the simulation by monitoring changes in certain key variables and by feedback. In order to demonstrate the viability of time adaptivity for engineering problems, we applied it to simulate water hammer and cryogenic chill down in pipelines. Our comparison and validation demonstrate the accuracy and efficiency of this adaptive strategy.

  7. Wave-particle interaction phenomena observed by antarctic rockets

    International Nuclear Information System (INIS)

    Kimura, I.; Hirasawa, T.

    1979-01-01

    Rocket measurements of wave and particles activities made at Syowa Station in Antarctica during IMS period are reviewed. Nine rockets were used for such observations, out of which 6 rockets were launched in the auroral sky. In the VLF frequency range, 0 - 10 KHz, wideband spectra of wave electric and magnetic fields, Poynting flux and the direction of propagation vector were measured for chorus, ELF and VLF hiss, and for electrostatic noises. In the MF and HF range, the dynamic frequency spectra of 0.1 - 10 MHz were measured. The relationship of these wave phenomena with energetic particle activities measured by the same rockets are discussed. (author)

  8. The electromagnetic rocket gun impact fusion driver

    International Nuclear Information System (INIS)

    Winterberg, F.

    1984-01-01

    A macroparticle accelerator to be used as an impact fusion driver is discussed and which can accelerate a small projectile to --200 km/sec over a distance of a few 100 meters. The driver which we have named electromagnetic rocket gun, accelerates a small rocket-like projectile by a travelling magnetic wave. The rocket propellant not only serves as a sink to absorb the heat produced in the projectile by resistive energy losses, but at the same time is also the source of additional thrust through the heating of the propellant to high temperatures by the travelling magnetic wave. The total thrust on the projectile is the sum of the magnetic and recoil forces. In comparison to a rocket, the efficiency is here much larger, with the momentum transferred to the gun barrel of the gun rather than to a tenuous jet. (author)

  9. Hybrid rocket motor testing at Nammo Raufoss A/S

    Science.gov (United States)

    Rønningen, Jan-Erik; Kubberud, Nils

    2005-08-01

    Hybrid rocket motor technology and the use of hybrid rockets have gained increased interest in recent years in many countries. A typical hybrid rocket consists of a tank containing the oxidizer in either liquid or gaseous state connected to the combustion chamber containing an injector, inert solid fuel grain and nozzle. Nammo Raufoss A/S has for almost 40 years designed and produced high-performance solid propellant rocket motors for many military missile systems as well as solid propellant rocket motors for civil space use. In 2003 an in-house technology program was initiated to investigate and study hybrid rocket technology. On 23 September 2004 the first in-house designed hybrid test rocket motor was static test fired at Nammo Raufoss Test Center. The oxidizer was gaseous oxygen contained in a tank pressurized to 10MPa, flow controlled through a sonic orifice into the combustion chamber containing a multi port radial injector and six bore cartridge-loaded fuel grain containing a modified HTPB fuel composition. The motor was ignited using a non-explosive heated wire. This paper will present what has been achieved at Nammo Raufoss since the start of the program.

  10. Maneuver of Spinning Rocket in Flight

    OpenAIRE

    HAYAKAWA, Satio; ITO, Koji; MATSUI, Yutaka; NOGUCHI, Kunio; UESUGI, Kuninori; YAMASHITA, Kojun

    1980-01-01

    A Yo-despin device successfully functioned to change in flight the precession axis of a sounding rocket for astronomical observation. The rocket attitudes before and after yodespin were measured with a UV star sensor, an infrared horizon sensor and an infrared telescope. Instrumentation and performance of these devices as well as the attitude data during flight are described.

  11. The Alabama Space and Rocket Center: The Second Decade.

    Science.gov (United States)

    Buckbee, Edward O.

    1983-01-01

    The Alabama Space and Rocket Center in Huntsville, the world's largest rocket and space museum, includes displays illustrating American rocket history, exhibits and demonstrations on rocketry principles and experiences, and simulations of space travel. A new project includes an integrated recreational-educational complex, described in the three…

  12. An historical collection of papers on nuclear thermal propulsion

    Science.gov (United States)

    The present volume of historical papers on nuclear thermal propulsion (NTP) encompasses NTP technology development regarding solid-core NTP technology, advanced concepts from the early years of NTP research, and recent activities in the field. Specific issues addressed include NERVA rocket-engine technology, the development of nuclear rocket propulsion at Los Alamos, fuel-element development, reactor testing for the Rover program, and an overview of NTP concepts and research emphasizing two decades of NASA research. Also addressed are the development of the 'nuclear light bulb' closed-cycle gas core and a demonstration of a fissioning UF6 gas in an argon vortex. The recent developments reviewed include the application of NTP to NASA's Lunar Space Transportation System, the use of NTP for the Space Exploration Initiative, and the development of nuclear rocket engines in the former Soviet Union.

  13. Measurements of temperature profiles at the exit of small rockets.

    Science.gov (United States)

    Griggs, M; Harshbarger, F C

    1966-02-01

    The sodium line reversal technique was used to determine the reversal temperature profile across the exit of small rockets. Measurements were made on one 73-kg thrust rocket, and two 23-kg thrust rockets with different injectors. The large rocket showed little variation of reversal temperature across the plume. However, the 23-kg rockets both showed a large decrease of reversal temperature from the axis to the edge of the plume. In addition, the sodium line reversal technique of temperature measurement was compared with an infrared technique developed in these laboratories.

  14. 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.

  15. 14 CFR 437.67 - Tracking a reusable suborbital rocket.

    Science.gov (United States)

    2010-01-01

    ... 14 Aeronautics and Space 4 2010-01-01 2010-01-01 false Tracking a reusable suborbital rocket. 437... a reusable suborbital rocket. A permittee must— (a) During permitted flight, measure in real time the position and velocity of its reusable suborbital rocket; and (b) Provide position and velocity...

  16. Stress and reliability analyses of multilayered composite cylinder under thermal and mechanical loads

    Science.gov (United States)

    Wang, Xiaohua

    The coupling resulting from the mutual influence of material thermal and mechanical parameters is examined in the thermal stress analysis of a multilayered isotropic composite cylinder subjected to sudden axisymmetric external and internal temperature. The method of complex frequency response functions together with the Fourier transform technique is utilized. Because the coupling parameters for some composite materials, such as carbon-carbon, are very small, the effect of coupling is neglected in the orthotropic thermal stress analysis. The stress distributions in multilayered orthotropic cylinders subjected to sudden axisymmetric temperature loading combined with dynamic pressure as well as asymmetric temperature loading are also obtained. The method of Fourier series together with the Laplace transform is utilized in solving the heat conduction equation and thermal stress analysis. For brittle materials, like carbon-carbon composites, the strength variability is represented by two or three parameter Weibull distributions. The 'weakest link' principle which takes into account both the carbon-carbon composite cylinders. The complex frequency response analysis is performed on a multilayered orthotropic cylinder under asymmetrical thermal load. Both deterministic and random thermal stress and reliability analyses can be based on the results of this frequency response analysis. The stress and displacement distributions and reliability of rocket motors under static or dynamic line loads are analyzed by an elasticity approach. Rocket motors are modeled as long hollow multilayered cylinders with an air core, a thick isotropic propellant inner layer and a thin orthotropic kevlar-epoxy case. The case is treated as a single orthotropic layer or a ten layered orthotropic structure. Five material properties and the load are treated as random variable with normal distributions when the reliability of the rocket motor is analyzed by the first-order, second-moment method (FOSM).

  17. Ablative thermal protection systems

    International Nuclear Information System (INIS)

    Vaniman, J.; Fisher, R.; Wojciechowski, C.; Dean, W.

    1983-01-01

    The procedures used to establish the TPS (thermal protection system) design of the SRB (solid rocket booster) element of the Space Shuttle vehicle are discussed. A final evaluation of the adequacy of this design will be made from data obtained from the first five Shuttle flights. Temperature sensors installed at selected locations on the SRB structure covered by the TPS give information as a function of time throughout the flight. Anomalies are to be investigated and computer design thermal models adjusted if required. In addition, the actual TPS ablator material loss is to be measured after each flight and compared with analytically determined losses. The analytical methods of predicting ablator performance are surveyed. 5 references

  18. Ceremony celebrates 50 years of rocket launches

    Science.gov (United States)

    2000-01-01

    Ceremony celebrates 50 years of rocket launches PL00C-10364.12 At the 50th anniversary ceremony celebrating the first rocket launch from pad 3 on what is now Cape Canaveral Air Force Station, Norris Gray waves to the audience. Gray was part of the team who successfully launched the first rocket, known as Bumper 8. The ceremony was hosted by the Air Force Space & Missile Museum Foundation, Inc. , and included launch of a Bumper 8 model rocket, presentation of a Bumper Award to Florida Sen. George Kirkpatrick by the National Space Club; plus remarks by Sen. Kirkpatrick, KSC's Center Director Roy Bridges, and the Commander of the 45th Space Wing, Brig. Gen. Donald Pettit. Also attending the ceremony were other members of the original Bumper 8 team. A reception followed at Hangar C. Since 1950 there have been a total of 3,245 launches from Cape Canaveral.

  19. 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.

  20. The NASA Sounding Rocket Program and space sciences

    Science.gov (United States)

    Gurkin, L. W.

    1992-01-01

    High altitude suborbital rockets (sounding rockets) have been extensively used for space science research in the post-World War II period; the NASA Sounding Rocket Program has been on-going since the inception of the Agency and supports all space science disciplines. In recent years, sounding rockets have been utilized to provide a low gravity environment for materials processing research, particularly in the commercial sector. Sounding rockets offer unique features as a low gravity flight platform. Quick response and low cost combine to provide more frequent spaceflight opportunities. Suborbital spacecraft design practice has achieved a high level of sophistication which optimizes the limited available flight times. High data-rate telemetry, real-time ground up-link command and down-link video data are routinely used in sounding rocket payloads. Standard, off-the-shelf, active control systems are available which limit payload body rates such that the gravitational environment remains less than 10(-4) g during the control period. Operational launch vehicles are available which can provide up to 7 minutes of experiment time for experiment weights up to 270 kg. Standard payload recovery systems allow soft impact retrieval of payloads. When launched from White Sands Missile Range, New Mexico, payloads can be retrieved and returned to the launch site within hours.

  1. Development of Displacement Gages Exposed to Solid Rocket Motor Internal Environments

    Science.gov (United States)

    Bolton, D. E.; Cook, D. J.

    2003-01-01

    The Space Shuttle Reusable Solid Rocket Motor (RSRM) has three non-vented segment-to-segment case field joints. These joints use an interference fit J-joint that is bonded at assembly with a Pressure Sensitive Adhesive (PSA) inboard of redundant O-ring seals. Full-scale motor and sub-scale test article experience has shown that the ability to preclude gas leakage past the J-joint is a function of PSA type, joint moisture from pre-assembly humidity exposure, and the magnitude of joint displacement during motor operation. To more accurately determine the axial displacements at the J-joints, two thermally durable displacement gages (one mechanical and one electrical) were designed and developed. The mechanical displacement gage concept was generated first as a non-electrical, self-contained gage to capture the maximum magnitude of the J-joint motion. When it became feasible, the electrical displacement gage concept was generated second as a real-time linear displacement gage. Both of these gages were refined in development testing that included hot internal solid rocket motor environments and simulated vibration environments. As a result of this gage development effort, joint motions have been measured in static fired RSRM J-joints where intentional venting was produced (Flight Support Motor #8, FSM-8) and nominal non-vented behavior occurred (FSM-9 and FSM-10). This data gives new insight into the nominal characteristics of the three case J-joint positions (forward, center and aft) and characteristics of some case J-joints that became vented during motor operation. The data supports previous structural model predictions. These gages will also be useful in evaluating J-joint motion differences in a five-segment Space Shuttle solid rocket motor.

  2. 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

  3. Trade-off analysis of high-aspect-ratio-cooling-channels for rocket engines

    International Nuclear Information System (INIS)

    Pizzarelli, Marco; Nasuti, Francesco; Onofri, Marcello

    2013-01-01

    Highlights: • Aspect ratio has a significant effect on cooling efficiency and hydraulic losses. • Minimizing power loss is of paramount importance in liquid rocket engine cooling. • A suitable quasi-2D model is used to get fast cooling system analysis. • Trade-off with assigned weight, temperature, and channel height or wall thickness. • Aspect ratio is found that minimizes power loss in the cooling circuit. -- Abstract: High performance liquid rocket engines are often characterized by rectangular cooling channels with high aspect ratio (channel height-to-width ratio) because of their proven superior cooling efficiency with respect to a conventional design. However, the identification of the optimum aspect ratio is not a trivial task. In the present study a trade-off analysis is performed on a cooling channel system that can be of interest for rocket engines. This analysis requires multiple cooling channel flow calculations and thus cannot be efficiently performed by CFD solvers. Therefore, a proper numerical approach, referred to as quasi-2D model, is used to have fast and accurate predictions of cooling system properties. This approach relies on its capability of describing the thermal stratification that occurs in the coolant and in the wall structure, as well as the coolant warming and pressure drop along the channel length. Validation of the model is carried out by comparison with solutions obtained with a validated CFD solver. Results of the analysis show the existence of an optimum channel aspect ratio that minimizes the requested pump power needed to overcome losses in the cooling circuit

  4. Testing of electroformed deposited iridium/powder metallurgy rhenium rockets

    Science.gov (United States)

    Reed, Brian D.; Dickerson, Robert

    1996-01-01

    High-temperature, oxidation-resistant chamber materials offer the thermal margin for high performance and extended lifetimes for radiation-cooled rockets. Rhenium (Re) coated with iridium (Ir) allow hours of operation at 2200 C on Earth-storable propellants. One process for manufacturing Ir/Re rocket chambers is the fabrication of Re substrates by powder metallurgy (PM) and the application of Ir coatings by using electroformed deposition (ED). ED Ir coatings, however, have been found to be porous and poorly adherent. The integrity of ED Ir coatings could be improved by densification after the electroforming process. This report summarizes the testing of two 22-N, ED Ir/PM Re rocket chambers that were subjected to post-deposition treatments in an effort to densify the Ir coating. One chamber was vacuum annealed, while the other chamber was subjected to hot isostatic pressure (HIP). The chambers were tested on gaseous oxygen/gaseous hydrogen propellants, at mixture ratios that simulated the oxidizing environments of Earth-storable propellants. ne annealed ED Ir/PM Re chamber was tested for a total of 24 firings and 4.58 hr at a mixture ratio of 4.2. After only 9 firings, the annealed ED Ir coating began to blister and spall upstream of the throat. The blistering and spalling were similar to what had been experienced with unannealed, as-deposited ED Ir coatings. The HIP ED Ir/PM Re chamber was tested for a total of 91 firings and 11.45 hr at mixture ratios of 3.2 and 4.2. The HIP ED Ir coating remained adherent to the Re substrate throughout testing; there were no visible signs of coating degradation. Metallography revealed, however, thinning of the HIP Ir coating and occasional pores in the Re layer upstream of the throat. Pinholes in the Ir coating may have provided a path for oxidation of the Re substrate at these locations. The HIP ED Ir coating proved to be more effective than vacuum annealed and as-deposited ED Ir. Further densification is still required to

  5. NASA Sounding Rocket Program Educational Outreach

    Science.gov (United States)

    Rosanova, G.

    2013-01-01

    Educational and public outreach is a major focus area for the National Aeronautics and Space Administration (NASA). The NASA Sounding Rocket Program (NSRP) shares in the belief that NASA plays a unique and vital role in inspiring future generations to pursue careers in science, mathematics, and technology. To fulfill this vision, the NSRP engages in a variety of educator training workshops and student flight projects that provide unique and exciting hands-on rocketry and space flight experiences. Specifically, the Wallops Rocket Academy for Teachers and Students (WRATS) is a one-week tutorial laboratory experience for high school teachers to learn the basics of rocketry, as well as build an instrumented model rocket for launch and data processing. The teachers are thus armed with the knowledge and experience to subsequently inspire the students at their home institution. Additionally, the NSRP has partnered with the Colorado Space Grant Consortium (COSGC) to provide a "pipeline" of space flight opportunities to university students and professors. Participants begin by enrolling in the RockOn! Workshop, which guides fledgling rocketeers through the construction and functional testing of an instrumentation kit. This is then integrated into a sealed canister and flown on a sounding rocket payload, which is recovered for the students to retrieve and process their data post flight. The next step in the "pipeline" involves unique, user-defined RockSat-C experiments in a sealed canister that allow participants more independence in developing, constructing, and testing spaceflight hardware. These experiments are flown and recovered on the same payload as the RockOn! Workshop kits. Ultimately, the "pipeline" culminates in the development of an advanced, user-defined RockSat-X experiment that is flown on a payload which provides full exposure to the space environment (not in a sealed canister), and includes telemetry and attitude control capability. The RockOn! and Rock

  6. Remote control video cameras on a suborbital rocket

    International Nuclear Information System (INIS)

    Wessling, Francis C.

    1997-01-01

    Three video cameras were controlled in real time from the ground to a sub-orbital rocket during a fifteen minute flight from White Sands Missile Range in New Mexico. Telemetry communications with the rocket allowed the control of the cameras. The pan, tilt, zoom, focus, and iris of two of the camera lenses, the power and record functions of the three cameras, and also the analog video signal that would be sent to the ground was controlled by separate microprocessors. A microprocessor was used to record data from three miniature accelerometers, temperature sensors and a differential pressure sensor. In addition to the selected video signal sent to the ground and recorded there, the video signals from the three cameras also were recorded on board the rocket. These recorders were mounted inside the pressurized segment of the rocket payload. The lenses, lens control mechanisms, and the three small television cameras were located in a portion of the rocket payload that was exposed to the vacuum of space. The accelerometers were also exposed to the vacuum of space

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

    Science.gov (United States)

    Davidian, Kenneth O.

    1990-01-01

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

  8. Lymphocytes on sounding rocket flights.

    Science.gov (United States)

    Cogoli-Greuter, M; Pippia, P; Sciola, L; Cogoli, A

    1994-05-01

    Cell-cell interactions and the formation of cell aggregates are important events in the mitogen-induced lymphocyte activation. The fact that the formation of cell aggregates is only slightly reduced in microgravity suggests that cells are moving and interacting also in space, but direct evidence was still lacking. Here we report on two experiments carried out on a flight of the sounding rocket MAXUS 1B, launched in November 1992 from the base of Esrange in Sweden. The rocket reached the altitude of 716 km and provided 12.5 min of microgravity conditions.

  9. 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

  10. Analysis of plume backflow around a nozzle lip in a nuclear rocket

    International Nuclear Information System (INIS)

    Chung, C.H.; Kim, S.C.; Stubbs, R.M.; De Witt, K.J.

    1993-06-01

    The structure of the flow around a nuclear thermal rocket nozzle lip has been investigated using the direct simulation Monte Carlo method. Special attention has been paid to the behavior of a small amount of harmful particles that may be present in the rocket exhaust gas. The harmful fission product particles are modeled by four inert gases whose molecular weights are in a range of 4 131. Atomic hydrogen, which exists in the flow due to the extremely high nuclear fuel temperature in the reactor, is also included. It is shown that the plume backflow is primarily determined by the thin subsonic fluid layer adjacent to the surface of the nozzle lip, and that the inflow boundary in the plume region has negligible effect on the backflow. It is also shown that a relatively large amount of the lighter species is scattered into the backflow region while the amount of the heavier species becomes negligible in this region due to extreme separation between the species. Results indicate that the backscattered molecules are very energetic and are fast-moving along the surface in the backflow region near the nozzle lip. 22 refs

  11. 14 CFR 437.95 - Inspection of additional reusable suborbital rockets.

    Science.gov (United States)

    2010-01-01

    ... suborbital rockets. 437.95 Section 437.95 Aeronautics and Space COMMERCIAL SPACE TRANSPORTATION, FEDERAL... of an Experimental Permit § 437.95 Inspection of additional reusable suborbital rockets. A permittee may launch or reenter additional reusable suborbital rockets of the same design under the permit after...

  12. Directional solidification of Al2-Cu-Al and Al3-Ni-Al eutectics during TEXUS rocket flight

    Science.gov (United States)

    Favier, J. J.; Degoer, J.

    1984-01-01

    One lamellar eutectic sample and one fiber-like eutectic sample were solidified directionally during the TEXUS-6 rocket flight. The microstructures and the results of the thermal analysis, obtained from the temperatures recorded on the cartridge skin, are compared. No appreciable modifications of the regularity of the eutectic structures were observed by passing from 1 g to 0.0001 g in these experiments. No steady state growth conditions were achieved in these experiments.

  13. Evaluation of the Effect of Exhausts from Liquid and Solid Rockets on Ozone Layer

    Science.gov (United States)

    Yamagiwa, Yoshiki; Ishimaki, Tetsuya

    This paper reports the analytical results of the influences of solid rocket and liquid rocket exhausts on ozone layer. It is worried about that the exhausts from solid propellant rockets cause the ozone depletion in the ozone layer. Some researchers try to develop the analytical model of ozone depletion by rocket exhausts to understand its physical phenomena and to find the effective design of rocket to minimize its effect. However, these models do not include the exhausts from liquid rocket although there are many cases to use solid rocket boosters with a liquid rocket at the same time in practical situations. We constructed combined analytical model include the solid rocket exhausts and liquid rocket exhausts to analyze their effects. From the analytical results, we find that the exhausts from liquid rocket suppress the ozone depletion by solid rocket exhausts.

  14. Metallic Hydrogen: A Game Changing Rocket Propellant

    Science.gov (United States)

    Silvera, Isaac F.

    2016-01-01

    The objective of this research is to produce metallic hydrogen in the laboratory using an innovative approach, and to study its metastability properties. Current theoretical and experimental considerations expect that extremely high pressures of order 4-6 megabar are required to transform molecular hydrogen to the metallic phase. When metallic hydrogen is produced in the laboratory it will be extremely important to determine if it is metastable at modest temperatures, i.e. remains metallic when the pressure is released. Then it could be used as the most powerful chemical rocket fuel that exists and revolutionize rocketry, allowing single-stage rockets to enter orbit and chemically fueled rockets to explore our solar system.

  15. Consort 1 sounding rocket flight

    Science.gov (United States)

    Wessling, Francis C.; Maybee, George W.

    1989-01-01

    This paper describes a payload of six experiments developed for a 7-min microgravity flight aboard a sounding rocket Consort 1, in order to investigate the effects of low gravity on certain material processes. The experiments in question were designed to test the effect of microgravity on the demixing of aqueous polymer two-phase systems, the electrodeposition process, the production of elastomer-modified epoxy resins, the foam formation process and the characteristics of foam, the material dispersion, and metal sintering. The apparatuses designed for these experiments are examined, and the rocket-payload integration and operations are discussed.

  16. Robust Exploration and Commercial Missions to the Moon Using Nuclear Thermal Rocket Propulsion and Lunar Liquid Oxygen Derived from FeO-Rich Pyroclasitc Deposits

    Science.gov (United States)

    Borowski, Stanley K.; Ryan, Stephen W.; Burke, Laura M.; McCurdy, David R.; Fittje, James E.; Joyner, Claude R.

    2018-01-01

    The nuclear thermal rocket (NTR) has frequently been identified as a key space asset required for the human exploration of Mars. This proven technology can also provide the affordable access through cislunar space necessary for commercial development and sustained human presence on the Moon. It is a demonstrated technology capable of generating both high thrust and high specific impulse (I(sub sp) approx. 900 s) twice that of today's best chemical rockets. Nuclear lunar transfer vehicles-consisting of a propulsion stage using three approx. 16.5-klb(sub f) small nuclear rocket engines (SNREs), an in-line propellant tank, plus the payload-are reusable, enabling a variety of lunar missions. These include cargo delivery and crewed lunar landing missions. Even weeklong ''tourism'' missions carrying passengers into lunar orbit for a day of sightseeing and picture taking are possible. The NTR can play an important role in the next phase of lunar exploration and development by providing a robust in-space lunar transportation system (LTS) that can allow initial outposts to evolve into settlements supported by a variety of commercial activities such as in-situ propellant production used to supply strategically located propellant depots and transportation nodes. The use of lunar liquid oxygen (LLO2) derived from iron oxide (FeO)-rich volcanic glass beads, found in numerous pyroclastic deposits on the Moon, can significantly reduce the launch mass requirements from Earth by enabling reusable, surface-based lunar landing vehicles (LLVs)that use liquid oxygen and hydrogen (LO2/LH2) chemical rocket engines. Afterwards, a LO2/LH2 propellant depot can be established in lunar equatorial orbit to supply the LTS. At this point a modified version of the conventional NTR-called the LO2-augmented NTR, or LANTR-is introduced into the LTS allowing bipropellant operation and leveraging the mission benefits of refueling with lunar-derived propellants for Earth return. The bipropellant LANTR

  17. Russian Meteorological and Geophysical Rockets of New Generation

    Science.gov (United States)

    Yushkov, V.; Gvozdev, Yu.; Lykov, A.; Shershakov, V.; Ivanov, V.; Pozin, A.; Afanasenkov, A.; Savenkov, Yu.; Kuznetsov, V.

    2015-09-01

    To study the process in the middle and upper atmosphere, ionosphere and near-Earth space, as well as to monitor the geophysical environment in Russian Federal Service for Hydrology and Environmental Monitoring (ROSHYDROMET) the development of new generation of meteorological and geophysical rockets has been completed. The modern geophysical research rocket system MR-30 was created in Research and Production Association RPA "Typhoon". The basis of the complex MR-30 is a new geophysical sounding rocket MN-300 with solid propellant, Rocket launch takes place at an angle of 70º to 90º from the launcher, which is a farm with a guide rail type required for imparting initial rotation rocket. The Rocket is spin stabilized with a spin rate between 5 and 7 Hz. Launch weight is 1564 kg, and the mass of the payload of 50 to 150 kg. MR-300 is capable of lifting up to 300 km, while the area of dispersion points for booster falling is an ellipse with parameters 37x 60 km. The payload of the rocket MN-300 consists of two sections: a sealed, located below the instrument compartment, and not sealed, under the fairing. Block of scientific equipment is formed on the platform in a modular layout. This makes it possible to solve a wide range of tasks and conduct research and testing technologies using a unique environment of space, as well as to conduct technological experiments testing and research systems and spacecraft equipment. New Russian rocket system MERA (MEteorological Rocket for Atmospheric Research) belongs to so called "dart" technique that provide lifting of small scientific payload up to altitude 100 km and descending with parachute. It was developed at Central Aerological Observatory jointly with State Unitary Enterprise Instrument Design Bureau. The booster provides a very rapid acceleration to about Mach 5. After the burning phase of the buster the dart is separated and continues ballistic flight for about 2 minutes. The dart carries the instrument payload+ parachute

  18. The National Space Science Data Center guide to international rocket data

    Science.gov (United States)

    Dubach, L. L.

    1972-01-01

    Background information is given which briefly describes the mission of the National Space Science Data Center (NSSDC), including its functions and systems, along with its policies and purposes for collecting rocket data. The operation of a machine-sensible rocket information system, which allows the Data Center to have convenient access to information and data concerning all rocket flights carrying scientific experiments, is also described. The central feature of this system, an index of rocket flights maintained on magnetic tape, is described. Standard outputs for NSSDC and for the World Data Center A (WDC-A) for Rockets and Satellites are described.

  19. 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

  20. ACCESS: Thermal Mechanical Design, Performance, and Status

    Science.gov (United States)

    Kaiser, Mary Elizabeth; Morris, M. J.; McCandliss, S. R.; Rauscher, B. J.; Kimble, R. A.; Kruk, J. W.; Wright, E. L.; Bohlin, R.; Kurucz, R. L.; Riess, A. G.; Pelton, R.; Deustua, S. E.; Dixon, W. V.; Sahnow, D. J.; Benford, D. J.; Gardner, J. P.; Feldman, P. D.; Moos, H. W.; Lampton, M.; Perlmutter, S.; Woodgate, B. E.

    2014-01-01

    Systematic errors associated with astrophysical data used to probe fundamental astrophysical questions, such as SNeIa observations used to constrain dark energy theories, are now rivaling and exceeding the statistical errors associated with these measurements. ACCESS: Absolute Color Calibration Experiment for Standard Stars is a series of rocket-borne sub-orbital missions and ground-based experiments designed to enable improvements in the precision of the astrophysical flux scale through the transfer of absolute laboratory detector standards from the National Institute of Standards and Technology (NIST) to a network of stellar standards with a calibration accuracy of 1% and a spectral resolving power of 500 across the 0.35 - 1.7μm bandpass. Achieving this level of accuracy requires characterization and stability of the instrument and detector including a thermal background that contributes less than 1% to the flux per resolution element in the NIR. We will present the instrument and calibration status with a focus on the thermal mechanical design and associated performance data. The detector control and performance will be presented in a companion poster (Morris, et al). NASA APRA sounding rocket grant NNX08AI65G supports this work.

  1. 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.

  2. 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.

  3. Thermal-hydraulic analysis of an annular fuel element: The Achilles' heel of the particle bed reactor

    International Nuclear Information System (INIS)

    Dibben, M.J.; Tuttle, R.F.

    1993-01-01

    The low pressure nuclear thermal propulsion (LPNTP) concept offers significant improvements in rocket engine specific impulse over rockets employment chemical propulsion. This study investigated a parametric thermal-hydraulic analysis of an annular fueld element, also referred to as a fuel pipe, using the computer code ATHENA (Advanced Thermal Hydraulic Energy Network Analyzer). The fuelpipe is an annular particle bed fuel element of the reactor with radially inward flow of hydrogen through the element. In this study, the outlet temperature of the hydrogen is parametrically related to key effects, including the reactor power at two different pressure drops, the effect of power coupling for in-core testing, and the effect of hydrogen flow rates. Results show that the temperature is linearly related to the reactor power, but not to pressure drop, and that cross flow inside the fuelpipe occurs at approximately 0.3 percent of the radial flow rates

  4. Technology for low cost solid rocket boosters.

    Science.gov (United States)

    Ciepluch, C.

    1971-01-01

    A review of low cost large solid rocket motors developed at the Lewis Research Center is given. An estimate is made of the total cost reduction obtainable by incorporating this new technology package into the rocket motor design. The propellant, case material, insulation, nozzle ablatives, and thrust vector control are discussed. The effect of the new technology on motor cost is calculated for a typical expandable 260-in. booster application. Included in the cost analysis is the influence of motor performance variations due to specific impulse and weight changes. It is found for this application that motor costs may be reduced by up to 30% and that the economic attractiveness of future large solid rocket motors will be improved when the new technology is implemented.

  5. Optimization of Construction of the rocket-assisted projectile

    Directory of Open Access Journals (Sweden)

    Arkhipov Vladimir

    2017-01-01

    Full Text Available New scheme of the rocket motor of rocket-assisted projectile providing the increase in distance of flight due to controlled and optimal delay time of ignition of the solid-propellant charge of the SRM and increase in reliability of initiation of the SRM by means of the autonomous system of ignition excluding the influence of high pressure gases of the propellant charge in the gun barrel has been considered. Results of the analysis of effectiveness of using of the ignition delay device on motion characteristics of the rocket-assisted projectile has been presented.

  6. The History of Rockets.

    Science.gov (United States)

    Newby, J. C.

    1988-01-01

    Discusses the origins and development of rockets mainly from the perspective of warfare. Includes some early enthusiasts, such as Congreve, Tsiolkovosky, Goddard, and Oberth. Describes developments from World War II, and during satellite development. (YP)

  7. Impact and mitigation of stratospheric ozone depletion by chemical rockets

    International Nuclear Information System (INIS)

    Mcdonald, A.J.

    1992-03-01

    The American Institute of Aeronautics and Astronautics (AIAA) conducted a workshop in conjunction with the 1991 AIAA Joint Propulsion Conference in Sacramento, California, to assess the impact of chemical rocket propulsion on the environment. The workshop included recognized experts from the fields of atmospheric physics and chemistry, solid rocket propulsion, liquid rocket propulsion, government, and environmental agencies, and representatives from several responsible environmental organizations. The conclusion from this workshop relative to stratospheric ozone depletion was that neither solid nor liquid rocket launchers have a significant impact on stratospheric ozone depletion, and that there is no real significant difference between the two

  8. Laser-fusion rocket for interplanetary propulsion

    International Nuclear Information System (INIS)

    Hyde, R.A.

    1983-01-01

    A rocket powered by fusion microexplosions is well suited for quick interplanetary travel. Fusion pellets are sequentially injected into a magnetic thrust chamber. There, focused energy from a fusion Driver is used to implode and ignite them. Upon exploding, the plasma debris expands into the surrounding magnetic field and is redirected by it, producing thrust. This paper discusses the desired features and operation of the fusion pellet, its Driver, and magnetic thrust chamber. A rocket design is presented which uses slightly tritium-enriched deuterium as the fusion fuel, a high temperature KrF laser as the Driver, and a thrust chamber consisting of a single superconducting current loop protected from the pellet by a radiation shield. This rocket can be operated with a power-to-mass ratio of 110 W gm -1 , which permits missions ranging from occasional 9 day VIP service to Mars, to routine 1 year, 1500 ton, Plutonian cargo runs

  9. Von Braun Rocket Team at Fort Bliss, Texas

    Science.gov (United States)

    1940-01-01

    The German Rocket Team, also known as the Von Braun Rocket Team, poses for a group photograph at Fort Bliss, Texas. After World War II ended in 1945, Dr. Wernher von Braun led some 120 of his Peenemuende Colleagues, who developed the V-2 rocket for the German military during the War, to the United Sttes under a contract to the U.S. Army Corps as part of Operation Paperclip. During the following five years the team worked on high altitude firings of the captured V-2 rockets at the White Sands Missile Range in New Mexico, and a guided missile development unit at Fort Bliss, Texas. In April 1950, the group was transferred to the Army Ballistic Missile Agency (ABMA) at Redstone Arsenal in Huntsville, Alabama, and continued to work on the development of the guided missiles for the U.S. Army until transferring to a newly established field center of the National Aeronautic and Space Administration (NASA), George C. Marshall Space Flight Center (MSFC).

  10. The Relativistic Rocket

    Science.gov (United States)

    Antippa, Adel F.

    2009-01-01

    We solve the problem of the relativistic rocket by making use of the relation between Lorentzian and Galilean velocities, as well as the laws of superposition of successive collinear Lorentz boosts in the limit of infinitesimal boosts. The solution is conceptually simple, and technically straightforward, and provides an example of a powerful…

  11. 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

  12. Designing on-Board Data Handling for EDF (Electric Ducted Fan) Rocket

    Science.gov (United States)

    Mulyana, A.; Faiz, L. A. A.

    2018-02-01

    The EDF (Electric Ducted Fan) rocket to launch requires a system of monitoring, tracking and controlling to allow the rocket to glide properly. One of the important components in the rocket is OBDH (On-Board Data Handling) which serves as a medium to perform commands and data processing. However, TTC (Telemetry, Tracking, and Command) are required to communicate between GCS (Ground Control Station) and OBDH on EDF rockets. So the design control system of EDF rockets and GCS for telemetry and telecommand needs to be made. In the design of integrated OBDH controller uses a lot of electronics modules, to know the behavior of rocket used IMU sensor (Inertial Measurement Unit) in which consist of 3-axis gyroscope sensor and Accelerometer 3-axis. To do tracking using GPS, compass sensor as a determinant of the direction of the rocket as well as a reference point on the z-axis of gyroscope sensor processing and used barometer sensors to measure the height of the rocket at the time of glide. The data can be known in real-time by sending data through radio modules at 2.4 GHz frequency using XBee-Pro S2B to GCS. By using windows filter, noises can be reduced, and it used to guarantee monitoring and controlling system can work properly.

  13. 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.

  14. Numerical Calculation of Effect of Elastic Deformation on Aerodynamic Characteristics of a Rocket

    Directory of Open Access Journals (Sweden)

    Laith K. Abbas

    2014-01-01

    Full Text Available The application and workflow of Computational Fluid Dynamics (CFD/Computational Structure Dynamics (CSD on solving the static aeroelastic problem of a slender rocket are introduced. To predict static aeroelastic behavior accurately, two-way coupling and inertia relief methods are used to calculate the static deformations and aerodynamic characteristics of the deformed rocket. The aerodynamic coefficients of rigid rocket are computed firstly and compared with the experimental data, which verified the accuracy of CFD output. The results of the analysis for elastic rocket in the nonspinning and spinning states are compared with the rigid ones. The results highlight that the rocket deformation aspects are decided by the normal force distribution along the rocket length. Rocket deformation becomes larger with increasing the flight angle of attack. Drag and lift force coefficients decrease and pitching moment coefficients increase due to rocket deformations, center of pressure location forwards, and stability of the rockets decreases. Accordingly, the flight trajectory may be affected by the change of these aerodynamic coefficients and stability.

  15. Rocket Assembly and Checkout Facility

    Data.gov (United States)

    Federal Laboratory Consortium — FUNCTION: Integrates, tests, and calibrates scientific instruments flown on sounding rocket payloads. The scientific instruments are assembled on an optical bench;...

  16. New class of thermosetting plastics has improved strength, thermal and chemical stability

    Science.gov (United States)

    Burns, E. A.; Dubrow, B.; Lubowitz, H. R.

    1967-01-01

    New class of thermosetting plastics has high hydrocarbon content, high stiffness, thermal stability, humidity resistance, and workability in the precured state. It is designated cyclized polydiene urethane, and is applicable as matrices to prepare chemically stable ablative materials for rocket nose cones of nozzles.

  17. Hypothetical Dark Matter/axion Rockets:. Dark Matter in Terms of Space Physics Propulsion

    Science.gov (United States)

    Beckwith, A.

    2010-12-01

    Current proposed photon rocket designs include the Nuclear Photonic Rocket and the Antimatter Photonic Rocket (proposed by Eugen Sanger in the 1950s, as reported by Ref. 1). This paper examines the feasibility of improving the thrust of photon-driven ramjet propulsion by using DM rocket propulsion. The open question is: would a heavy WIMP, if converted to photons, upgrade the power (thrust) of a photon rocket drive, to make interstellar travel a feasible proposition?

  18. 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...

  19. Linear stability analysis in a solid-propellant rocket motor

    Energy Technology Data Exchange (ETDEWEB)

    Kim, K.M.; Kang, K.T.; Yoon, J.K. [Agency for Defense Development, Taejon (Korea, Republic of)

    1995-10-01

    Combustion instability in solid-propellant rocket motors depends on the balance between acoustic energy gains and losses of the system. The objective of this paper is to demonstrate the capability of the program which predicts the standard longitudinal stability using acoustic modes based on linear stability analysis and T-burner test results of propellants. Commercial ANSYS 5.0A program can be used to calculate the acoustic characteristic of a rocket motor. The linear stability prediction was compared with the static firing test results of rocket motors. (author). 11 refs., 17 figs.

  20. The Norwegian sounding rocket programme 1980-83

    International Nuclear Information System (INIS)

    Egeland, A.; Gundersen, A.

    1980-01-01

    As illustrated by the rocket program presented and discussed in this paper, exploration of the polar ionosphere still plays a central part in the Norwegian research program in science. A cornerstone in the Norwegian space program is the Andoeya Rocket Range. It will be shown that advanced radio installations in northern Scandinavia together with the new optical site at Svalbard will stimulate towards further in situ measurements and theoretical work of the polar ionosphere. (Auth.)

  1. Rocket Tablet,

    Science.gov (United States)

    1984-09-12

    not accustomed to Chinese food, he ran off directly to the home of the Mayor of Beijing and requested two Western cuisine cooks from a hotel. At the...played out by our Chinese sons and daughters of ancient times. The famous Han dynasty general Li Guang was quickly cured of disease and led an army...Union) of China. This place was about to become the birthplace of the Chinese people’s first rocket baby. Section One In this eternal wasteland called

  2. 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.

  3. Hydrodynamic Stability Analysis of Particle-Laden Solid Rocket Motors

    Science.gov (United States)

    Elliott, T. S.; Majdalani, J.

    2014-11-01

    Fluid-wall interactions within solid rocket motors can result in parietal vortex shedding giving rise to hydrodynamic instabilities, or unsteady waves, that translate into pressure oscillations. The oscillations can result in vibrations observed by the rocket, rocket subsystems, or payload, which can lead to changes in flight characteristics, design failure, or other undesirable effects. For many years particles have been embedded in solid rocket propellants with the understanding that their presence increases specific impulse and suppresses fluctuations in the flowfield. This study utilizes a two dimensional framework to understand and quantify the aforementioned two-phase flowfield inside a motor case with a cylindrical grain perforation. This is accomplished through the use of linearized Navier-Stokes equations with the Stokes drag equation and application of the biglobal ansatz. Obtaining the biglobal equations for analysis requires quantification of the mean flowfield within the solid rocket motor. To that end, the extended Taylor-Culick form will be utilized to represent the gaseous phase of the mean flowfield while the self-similar form will be employed for the particle phase. Advancing the mean flowfield by quantifying the particle mass concentration with a semi-analytical solution the finalized mean flowfield is combined with the biglobal equations resulting in a system of eight partial differential equations. This system is solved using an eigensolver within the framework yielding the entire spectrum of eigenvalues, frequency and growth rate components, at once. This work will detail the parametric analysis performed to demonstrate the stabilizing and destabilizing effects of particles within solid rocket combustion.

  4. Hydrodynamic Stability Analysis of Particle-Laden Solid Rocket Motors

    International Nuclear Information System (INIS)

    Elliott, T S; Majdalani, J

    2014-01-01

    Fluid-wall interactions within solid rocket motors can result in parietal vortex shedding giving rise to hydrodynamic instabilities, or unsteady waves, that translate into pressure oscillations. The oscillations can result in vibrations observed by the rocket, rocket subsystems, or payload, which can lead to changes in flight characteristics, design failure, or other undesirable effects. For many years particles have been embedded in solid rocket propellants with the understanding that their presence increases specific impulse and suppresses fluctuations in the flowfield. This study utilizes a two dimensional framework to understand and quantify the aforementioned two-phase flowfield inside a motor case with a cylindrical grain perforation. This is accomplished through the use of linearized Navier-Stokes equations with the Stokes drag equation and application of the biglobal ansatz. Obtaining the biglobal equations for analysis requires quantification of the mean flowfield within the solid rocket motor. To that end, the extended Taylor-Culick form will be utilized to represent the gaseous phase of the mean flowfield while the self-similar form will be employed for the particle phase. Advancing the mean flowfield by quantifying the particle mass concentration with a semi-analytical solution the finalized mean flowfield is combined with the biglobal equations resulting in a system of eight partial differential equations. This system is solved using an eigensolver within the framework yielding the entire spectrum of eigenvalues, frequency and growth rate components, at once. This work will detail the parametric analysis performed to demonstrate the stabilizing and destabilizing effects of particles within solid rocket combustion

  5. Energy production using fission fragment rockets

    International Nuclear Information System (INIS)

    Chapline, G.; Matsuda, Y.

    1991-08-01

    Fission fragment rockets are nuclear reactors with a core consisting of thin fibers in a vacuum, and which use magnetic fields to extract the fission fragments from the reactor core. As an alternative to ordinary nuclear reactors, fission fragment rockets would have the following advantages: Approximately twice as efficient if one can directly convert the fission fragment energy into electricity; by reducing the buildup of a fission fragment inventory in the reactor one could avoid a Chernobyl type disaster; and collecting the fission fragments outside the reactor could simplify the waste disposal problem. 6 refs., 4 figs., 2 tabs

  6. 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

  7. 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

  8. 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.

  9. Vibration Disturbance Damping System Design to Protect Payload of the Rocket

    Directory of Open Access Journals (Sweden)

    Sutisno Sutisno

    2012-12-01

    Full Text Available Rocket motor generates vibrations acting on whole rocket body including its contents. Part of the body which is sensitive to disturbance is the rocket payload. The payload consists of various electronic instruments including: transmitter, various sensors, accelerometer, gyro, the embedded controller system, and others. This paper presents research on rocket vibration influence to the payload and the method to avoid disturbance. Avoiding influence of vibration disturbance can be done using silicone gel material whose typical damping factors are relatively high. The rocket vibration was simulated using electromagnetic motor, and the vibrations were measured using an accelerometer sensor. The measurement results were displayed in the form of curve, indicating the vibration level on some parts of the tested material. Some measurement results can be applied to determine the good material to attenuate vibration disturbance on the instruments of the payload.

  10. Study of Liquid Breakup Process in Solid Rocket Motor Nozzle

    Science.gov (United States)

    2016-02-16

    Laboratory, Edwards, CA Abstract In a solid rocket motor (SRM), when the aluminum based propellant combusts, the fuel is oxidized into alumina (Al2O3...34Chemical Erosion of Refractory-Metal Nozzle Inserts in Solid - Propellant Rocket Motors," J. Propulsion and Power, Vol. 25, no.1,, 2009. [4] E. Y. Wong...34 Solid Rocket Nozzle Design Summary," in 4th AIAA Propulsion Joint Specialist Conference, Cleveland, OH, 1968. [5] Nayfeh, A. H.; Saric, W. S

  11. Rocket observation of electron density irregularities in the lower E region

    International Nuclear Information System (INIS)

    Watanabe, Yuzo; Nakamura, Yoshiharu; Amemiya, Hiroshi.

    1990-01-01

    Local ionospheric electron density irregularities in the scale size of 3 m to 300 m have been measured on the ascending path from 74 km to 93 km by a fix biased Langmuir probe on board the S-310-16 sounding rocket. The rocket was launched at 22:40:00 on February 1, 1986 from Kagoshima Space Center in Japan. It is found from frequency analysis of the data that the spectral index of the irregularities is 0.9 to 1.8 and the irregularity amplitude is 1 to 15 %. The altitude where the amplitude reaches its maximum is 88 km. The generation mechanism of these irregularities is explained by the neutral turbulence theory, which indicates that the spectral index is 5/3 and has been confirmed by a chemical release experiment using rockets over India to be valid up to about 110 km. From frequency analysis of the data observed during the descent in the lower E region, we have found that the rocket-wake effect becomes larger when the probe is situated near the edge of the rocket-wake, and that this is also the case even when the rocket-wake effect does not clearly appear in the DC current signal which approximately changes in proportion to the electron density, where the probe is completely situated inside the rocket-wake region. (author)

  12. Metallized solid rocket propellants based on AN/AP and PSAN/AP for access to space

    Science.gov (United States)

    Levi, S.; Signoriello, D.; Gabardi, A.; Molinari, M.; Galfetti, L.; Deluca, L. T.; Cianfanelli, S.; Klyakin, G. F.

    2009-09-01

    Solid rocket propellants based on dual mixes of inorganic crystalline oxidizers (ammonium nitrate (AN) and ammonium perchlorate (AP)) with binder and a mixture of micrometric-nanometric aluminum were investigated. Ammonium nitrate is a low-cost oxidizer, producing environment friendly combustion products but with lower specific impulse compared to AP. The better performance obtained with AP and the low quantity of toxic emissions obtained by using AN have suggested an interesting compromise based on a dual mixture of the two oxidizers. To improve the thermal response of raw AN, different types of phase stabilized AN (PSAN) and AN/AP co-crystals were investigated.

  13. 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.

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

    Directory of Open Access Journals (Sweden)

    Luís Antonio Silva

    2011-05-01

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

  15. 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.

  16. 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.

  17. Design of a Nickel-Based Bond-Coat Alloy for Thermal Barrier Coatings on Copper Substrates

    Directory of Open Access Journals (Sweden)

    Torben Fiedler

    2014-11-01

    Full Text Available To increase the lifetime of rocket combustion chambers, thermal barrier coatings (TBC may be applied on the copper chamber wall. Since standard TBC systems used in gas turbines are not suitable for rocket-engine application and fail at the interface between the substrate and bond coat, a new bond-coat material has to be designed. This bond-coat material has to be chemically compatible to the copper substrate to improve the adhesion and needs a coefficient of thermal expansion close to that of copper to reduce thermal stresses. One approach to achieve this is to modify the standard NiCrAlY alloy used in gas turbines by adding copper. In this work, the influence of copper on the microstructure of NiCrAlY-alloys is investigated with thermodynamical calculations, optical microscopy, SEM, EDX and calorimetry. Adding copper leads to the formation of a significant amount of \\(\\beta\\ and \\(\\alpha\\ Reducing the aluminum and chromium content leads furthermore to a two-phase fcc microstructure.

  18. Infrasound from the 2009 and 2017 DPRK rocket launches

    Science.gov (United States)

    Evers, L. G.; Assink, J. D.; Smets, P. SM

    2018-06-01

    Supersonic rockets generate low-frequency acoustic waves, that is, infrasound, during the launch and re-entry. Infrasound is routinely observed at infrasound arrays from the International Monitoring System, in place for the verification of the Comprehensive Nuclear-Test-Ban Treaty. Association and source identification are key elements of the verification system. The moving nature of a rocket is a defining criterion in order to distinguish it from an isolated explosion. Here, it is shown how infrasound recordings can be associated, which leads to identification of the rocket. Propagation modelling is included to further constrain the source identification. Four rocket launches by the Democratic People's Republic of Korea in 2009 and 2017 are analysed in which multiple arrays detected the infrasound. Source identification in this region is important for verification purposes. It is concluded that with a passive monitoring technique such as infrasound, characteristics can be remotely obtained on sources of interest, that is, infrasonic intelligence, over 4500+ km.

  19. The UK sounding rocket and balloon programme

    International Nuclear Information System (INIS)

    Delury, J.T.

    1980-01-01

    The UK civil science balloon and rocket programmes for 1979/80/81 are summarised and the areas of scientific interest for the period 1981/85 mentioned. In the main the facilities available are 10 in number balloons up to 40 m cu ft launched from USA or Australia and up to 10 in number 7 1/2'' diameter Petrel rockets. This paper outlines the 1979 and 1980 programmes and explains the longer term plans covering the next 5 years. (Auth.)

  20. Reactive Inkjet Printing of Biocompatible Enzyme Powered Silk Micro-Rockets.

    Science.gov (United States)

    Gregory, David A; Zhang, Yu; Smith, Patrick J; Zhao, Xiubo; Ebbens, Stephen J

    2016-08-01

    Inkjet-printed enzyme-powered silk-based micro-rockets are able to undergo autonomous motion in a vast variety of fluidic environments including complex media such as human serum. By means of digital inkjet printing it is possible to alter the catalyst distribution simply and generate varying trajectory behavior of these micro-rockets. Made of silk scaffolds containing enzymes these micro-rockets are highly biocompatible and non-biofouling. © 2016 The Authors. Published by WILEY-VCH Verlag GmbH & Co. KGaA, Weinheim.

  1. National Report on the NASA Sounding Rocket and Balloon Programs

    Science.gov (United States)

    Eberspeaker, Philip; Fairbrother, Debora

    2013-01-01

    The U. S. National Aeronautics and Space Administration (NASA) Sounding Rockets and Balloon Programs conduct a total of 30 to 40 missions per year in support of the NASA scientific community and other users. The NASA Sounding Rockets Program supports the science community by integrating their experiments into the sounding rocket payloads, and providing both the rocket vehicle and launch operations services. Activities since 2011 have included two flights from Andoya Rocket Range, more than eight flights from White Sands Missile Range, approximately sixteen flights from Wallops Flight Facility, two flights from Poker Flat Research Range, and four flights from Kwajalein Atoll. Other activities included the final developmental flight of the Terrier-Improved Malemute launch vehicle, a test flight of the Talos-Terrier-Oriole launch vehicle, and a host of smaller activities to improve program support capabilities. Several operational missions have utilized the new Terrier-Malemute vehicle. The NASA Sounding Rockets Program is currently engaged in the development of a new sustainer motor known as the Peregrine. The Peregrine development effort will involve one static firing and three flight tests with a target completion data of August 2014. The NASA Balloon Program supported numerous scientific and developmental missions since its last report. The program conducted flights from the U.S., Sweden, Australia, and Antarctica utilizing standard and experimental vehicles. Of particular note are the successful test flights of the Wallops Arc Second Pointer (WASP), the successful demonstration of a medium-size Super Pressure Balloon (SPB), and most recently, three simultaneous missions aloft over Antarctica. NASA continues its successful incremental design qualification program and will support a science mission aboard WASP in late 2013 and a science mission aboard the SPB in early 2015. NASA has also embarked on an intra-agency collaboration to launch a rocket from a balloon to

  2. Intense auroral field-aligned currents and electrojets detected by rocket-borne fluxgate magnetometer

    International Nuclear Information System (INIS)

    Tohyama, Fumio; Fukunishi, Hiroshi; Takahashi, Takao; Kokubun, Susumu; Fujii, Ryoichi; Yamagishi, Hisao.

    1988-01-01

    The S-310JA-11 and S-310JA-12 rockets, having a vector magnetometer with high sensitivity (1.8 nT) and high sampling frequency (100 Hz), were launched into the aurora on May 29 and July 12, 1985, from Syowa Station, Antarctica. The S-310JA-11 rocket penetrated twice quiet arcs, while the S-310JA-12 rocket traversed across intense and active auroral arcs during a large magnetic substorm. In the S-310JA-12 rocket experiment, intense field-aligned currents of 400 - 600 nT were observed when the rocket penetrated an active arc during the descending flight. The magnetometer on board the S-310JA-12 rocket also detected intense electrojet currents with a center at 110 km on the upward leg and at 108 km on the downward leg. The magnetometer data of the S-310JA-11 rocket showed no distinguished magnetic field variation due to field-aligned current and electrojet. (author)

  3. Theoretical Rocket Performance of Liquid Methane with Several Fluorine-Oxygen Mixtures Assuming Frozen Composition

    Science.gov (United States)

    Gordon, Sanford; Kastner, Michael E

    1958-01-01

    Theoretical rocket performance for frozen composition during expansion was calculated for liquid methane with several fluorine-oxygen mixtures for a range of pressure ratios and oxidant-fuel ratios. The parameters included are specific impulse, combustion-chamber temperature, nozzle-exit temperature molecular weight, characteristic velocity, coefficient of thrust, ratio of nozzle-exit area to throat area, specific heat at constant pressure, isentropic exponent, viscosity, and thermal conductivity. The maximum calculated value of specific impulse for a chamber pressure of 600 pounds per square inch absolute (40.827atm) and an exit pressure of 1 atmosphere is 315.3 for 79.67 percent fluorine in the oxidant.

  4. Propulsion and launching analysis of variable-mass rockets by analytical methods

    Directory of Open Access Journals (Sweden)

    D.D. Ganji

    2013-09-01

    Full Text Available In this study, applications of some analytical methods on nonlinear equation of the launching of a rocket with variable mass are investigated. Differential transformation method (DTM, homotopy perturbation method (HPM and least square method (LSM were applied and their results are compared with numerical solution. An excellent agreement with analytical methods and numerical ones is observed in the results and this reveals that analytical methods are effective and convenient. Also a parametric study is performed here which includes the effect of exhaust velocity (Ce, burn rate (BR of fuel and diameter of cylindrical rocket (d on the motion of a sample rocket, and contours for showing the sensitivity of these parameters are plotted. The main results indicate that the rocket velocity and altitude are increased with increasing the Ce and BR and decreased with increasing the rocket diameter and drag coefficient.

  5. Simulation and experimental research on line throwing rocket with flight

    Directory of Open Access Journals (Sweden)

    Wen-bin Gu

    2014-06-01

    Full Text Available The finite segment method is used to model the line throwing rocket system. A dynamic model of line throwing rocket with flight motion based on Kane's method is presented by the kinematics description of the system and the consideration of the forces acting on the system. The experiment designed according to the parameters of the dynamic model is made. The simulation and experiment results, such as range, velocity and flight time, are compared and analyzed. The simulation results are basically agreed with the test data, which shows that the flight motion of the line throwing rocket can be predicted by the dynamic model. A theoretical model and guide for the further research on the disturbance of rope and the guidance, flight control of line throwing rocket are provided by the dynamic modeling.

  6. MEMS-Based Solid Propellant Rocket Array Thruster

    Science.gov (United States)

    Tanaka, Shuji; Hosokawa, Ryuichiro; Tokudome, Shin-Ichiro; Hori, Keiichi; Saito, Hirobumi; Watanabe, Masashi; Esashi, Masayoshi

    The prototype of a solid propellant rocket array thruster for simple attitude control of a 10 kg class micro-spacecraft was completed and tested. The prototype has 10×10 φ0.8 mm solid propellant micro-rockets arrayed at a pitch of 1.2 mm on a 20×22 mm substrate. To realize such a dense array of micro-rockets, each ignition heater is powered from the backside of the thruster through an electrical feedthrough which passes along a propellant cylinder wall. Boron/potassium nitrate propellant (NAB) is used with/without lead rhodanide/potassium chlorate/nitrocellulose ignition aid (RK). Impulse thrust was measured by a pendulum method in air. Ignition required electric power of at least 3 4 W with RK and 4 6 W without RK. Measured impulse thrusts were from 2×10-5 Ns to 3×10-4 Ns after the calculation of compensation for air dumping.

  7. Analysis of plasma behavior in a magnetic nozzle of laser fusion rocket

    International Nuclear Information System (INIS)

    Nagamine, Yoshihiko; Yoshimi, Naofumi; Nakama, Yuji; Muranaka, Takanobu; Mayumi, Takao; Nakashima, Hideki

    1997-01-01

    A magnetic nozzle concept in a laser fusion rocket is suitable for controlling the fusion plasma flow and it has an advantage that thermalization with wall structures in a thrust chamber can be avoided. Rayleigh-Taylor instability would occur at the surface of expanding plasma and it would lead to the degradation of thrust efficiency, due to diffusion of the plasma through ambient decelerating magnetic field. A 3D hybrid particle-in-cell code has been developed to analyze the plasma instability in the magnetic nozzle. The resultant linear growth rate γ of the instability is found to be 2.96 x 10 6 and it is in good agreement with the theoretical value from conventional Rayleigh Taylor instability. (author)

  8. Development of high performance hybrid rocket fuels

    Science.gov (United States)

    Zaseck, Christopher R.

    . In order to examine paraffin/additive combustion in a motor environment, I conducted experiments on well characterized aluminum based additives. In particular, I investigate the influence of aluminum, unpassivated aluminum, milled aluminum/polytetrafluoroethylene (PTFE), and aluminum hydride on the performance of paraffin fuels for hybrid rocket propulsion. I use an optically accessible combustor to examine the performance of the fuel mixtures in terms of characteristic velocity efficiency and regression rate. Each combustor test consumes a 12.7 cm long, 1.9 cm diameter fuel strand under 160 kg/m 2s of oxygen at up to 1.4 MPa. The experimental results indicate that the addition of 5 wt.% 30 mum or 80 nm aluminum to paraffin increases the regression rate by approximately 15% compared to neat paraffin grains. At higher aluminum concentrations and nano-scale particles sizes, the increased melt layer viscosity causes slower regression. Alane and Al/PTFE at 12.5 wt.% increase the regression of paraffin by 21% and 32% respectively. Finally, an aging study indicates that paraffin can protect air and moisture sensitive particles from oxidation. The opposed burner and aluminum/paraffin hybrid rocket experiments show that additives can alter bulk fuel properties, such as viscosity, that regulate entrainment. The general effect of melt layer properties on the entrainment and regression rate of paraffin is not well understood. Improved understanding of how solid additives affect the properties and regression of paraffin is essential to maximize performance. In this document I investigate the effect of melt layer properties on paraffin regression using inert additives. Tests are performed in the optical cylindrical combustor at ˜1 MPa under a gaseous oxygen mass flux of ˜160 kg/m2s. The experiments indicate that the regression rate is proportional to mu0.08rho 0.38kappa0.82. In addition, I explore how to predict fuel viscosity, thermal conductivity, and density prior to testing

  9. 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

  10. A Multiconstrained Ascent Guidance Method for Solid Rocket-Powered Launch Vehicles

    Directory of Open Access Journals (Sweden)

    Si-Yuan Chen

    2016-01-01

    Full Text Available This study proposes a multiconstrained ascent guidance method for a solid rocket-powered launch vehicle, which uses a hypersonic glide vehicle (HGV as payload and shuts off by fuel exhaustion. First, pseudospectral method is used to analyze the two-stage launch vehicle ascent trajectory with different rocket ignition modes. Then, constraints, such as terminal height, velocity, flight path angle, and angle of attack, are converted into the constraints within height-time profile according to the second-stage rocket flight characteristics. The closed-loop guidance method is inferred by different spline curves given the different terminal constraints. Afterwards, a thrust bias energy management strategy is proposed to waste the excess energy of the solid rocket. Finally, the proposed method is verified through nominal and dispersion simulations. The simulation results show excellent applicability and robustness of this method, which can provide a valuable reference for the ascent guidance of solid rocket-powered launch vehicles.

  11. A two-channel wave analyser for sounding rockets and satellites

    International Nuclear Information System (INIS)

    Brondz, E.

    1989-04-01

    Studies of low frequency electromagnetic waves, produced originally by lightning discharges penetrating the ionosphere, provide an important source of valuable information about the earth's surrounding plasma. Use of rockets and satellites supported by ground-based observations implies, unique opportunity for measuring in situ a number of parameters simultaneously in order to correlate data from various measurements. However, every rocket experiment has to be designed bearing in mind telemetry limitations and/or short flight duration. Typical flight duration for Norwegian rockets launched from Andoeya Rocket Range is 500 to 600 s. Therefore, the most desired way to use a rocket or satellite is to carry out data analyses on board in real time. Recent achievements in Digital Signal Processing (DSP) technology have made it possible to undertake very complex on board data manipulation. As a part of rocket instrumentation, a DSP based unit able to carry out on board analyses of low frequency electromagnetic waves in the ionosphere has been designed. The unit can be seen as a general purpose computer built on the basis of a fixed-point 16 bit signal processor. The unit is supplied with a program code in order to perform wave analyses on two independent channels simultaneously. The analyser is able to perform 256 point complex fast fourier transformations, and it produce a spectral power desity estimate on both channels every 85 ms. The design and construction of the DSP based unit is described and results from the tests are presented

  12. Dynamic Analysis of Sounding Rocket Pneumatic System Revision

    Science.gov (United States)

    Armen, Jerald

    2010-01-01

    The recent fusion of decades of advancements in mathematical models, numerical algorithms and curve fitting techniques marked the beginning of a new era in the science of simulation. It is becoming indispensable to the study of rockets and aerospace analysis. In pneumatic system, which is the main focus of this paper, particular emphasis will be placed on the efforts of compressible flow in Attitude Control System of sounding rocket.

  13. The relativistic rocket

    Energy Technology Data Exchange (ETDEWEB)

    Antippa, Adel F [Departement de Physique, Universite du Quebec a Trois-Rivieres, Trois-Rivieres, Quebec G9A 5H7 (Canada)

    2009-05-15

    We solve the problem of the relativistic rocket by making use of the relation between Lorentzian and Galilean velocities, as well as the laws of superposition of successive collinear Lorentz boosts in the limit of infinitesimal boosts. The solution is conceptually simple, and technically straightforward, and provides an example of a powerful method that can be applied to a wide range of special relativistic problems of linear acceleration.

  14. 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

  15. How High? How Fast? How Long? Modeling Water Rocket Flight with Calculus

    Science.gov (United States)

    Ashline, George; Ellis-Monaghan, Joanna

    2006-01-01

    We describe an easy and fun project using water rockets to demonstrate applications of single variable calculus concepts. We provide procedures and a supplies list for launching and videotaping a water rocket flight to provide the experimental data. Because of factors such as fuel expulsion and wind effects, the water rocket does not follow the…

  16. Project Stratos; reaching space with a student-built rocket

    NARCIS (Netherlands)

    Haneveer, M.

    2013-01-01

    In the spring of 2009 a team of 15 TU Delft students travelled to Kiruna, Sweden with only one goal: to launch the rocket Stratos I they had been working on for 2 years to an altitude of over 12km, thereby claiming the European Amateur Rocket Altitude record. These students were part of Delft

  17. 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.

  18. 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%.

  19. Rockets for Extended Source Soft X-ray Spectroscopy

    Science.gov (United States)

    McEntaffer, Randall

    The soft X-ray background surrounds our local galactic environment yet very little is known about the physical characteristics of this plasma. A high-resolution spectrum could unlock the properties of this million degree gas but the diffuse, low intensity nature of the background have made it difficult to observe, especially with a dispersive spectrograph. Previous observations have relied on X-ray detector energy resolution which produces poorly defined spectra that are poorly fit by complex plasma models. Here we propose a series of suborbital rocket flights that will begin the characterization of this elusive source through high-resolution X-ray grating spectroscopy. The rocket-based spectrograph can resolve individual emission lines over the soft X-ray band and place tight constraints on the temperature, density, abundance, ionization state and age of the plasma. These payloads will draw heavily from the heritage gained from previous rocket missions, while also benefiting from related NASA technology development programs. The Pennsylvania State University (PSU) team has a history of designing and flying spectrometer components onboard rockets while also being scientific leaders in the field of diffuse soft X-ray astronomy. The PSU program will provide hands-on training of young scientists in the techniques of instrumental and observational X-ray astronomy. The proposed rocket program will also expose these researchers to a full experiment cycle: design, fabrication, tolerance analysis, assembly, flight-qualification, calibration, integration, launch, and data analysis; using a combination of technologies suitable for adaptation to NASA's major missions. The PSU program in suborbital X-ray astronomy represents an exciting mix of compelling science, heritage, cutting-edge technology development, and training of future scientists.

  20. In-flight calibration of mesospheric rocket plasma probes

    International Nuclear Information System (INIS)

    Havnes, Ove; Hartquist, Thomas W.; Kassa, Meseret; Morfill, Gregor E.

    2011-01-01

    Many effects and factors can influence the efficiency of a rocket plasma probe. These include payload charging, solar illumination, rocket payload orientation and rotation, and dust impact induced secondary charge production. As a consequence, considerable uncertainties can arise in the determination of the effective cross sections of plasma probes and measured electron and ion densities. We present a new method for calibrating mesospheric rocket plasma probes and obtaining reliable measurements of plasma densities. This method can be used if a payload also carries a probe for measuring the dust charge density. It is based on that a dust probe's effective cross section for measuring the charged component of dust normally is nearly equal to its geometric cross section, and it involves the comparison of variations in the dust charge density measured with the dust detector to the corresponding current variations measured with the electron and/or ion probes. In cases in which the dust charge density is significantly smaller than the electron density, the relation between plasma and dust charge density variations can be simplified and used to infer the effective cross sections of the plasma probes. We illustrate the utility of the method by analysing the data from a specific rocket flight of a payload containing both dust and electron probes.

  1. In-flight calibration of mesospheric rocket plasma probes

    Energy Technology Data Exchange (ETDEWEB)

    Havnes, Ove [Institute for Physics and Technology, University of Tromsoe, N-9037 Tromsoe (Norway); University Studies Svalbard (UNIS), N-9170 Longyearbyen, Svalbard (Norway); Hartquist, Thomas W. [School of Physics and Astronomy, University of Leeds, Leeds LS2 9JT (United Kingdom); Kassa, Meseret [Institute for Physics and Technology, University of Tromsoe, N-9037 Tromsoe (Norway); Morfill, Gregor E. [Max-Planck-Institute fuer extraterrestrische Physik, D-85741Garching (Germany)

    2011-07-15

    Many effects and factors can influence the efficiency of a rocket plasma probe. These include payload charging, solar illumination, rocket payload orientation and rotation, and dust impact induced secondary charge production. As a consequence, considerable uncertainties can arise in the determination of the effective cross sections of plasma probes and measured electron and ion densities. We present a new method for calibrating mesospheric rocket plasma probes and obtaining reliable measurements of plasma densities. This method can be used if a payload also carries a probe for measuring the dust charge density. It is based on that a dust probe's effective cross section for measuring the charged component of dust normally is nearly equal to its geometric cross section, and it involves the comparison of variations in the dust charge density measured with the dust detector to the corresponding current variations measured with the electron and/or ion probes. In cases in which the dust charge density is significantly smaller than the electron density, the relation between plasma and dust charge density variations can be simplified and used to infer the effective cross sections of the plasma probes. We illustrate the utility of the method by analysing the data from a specific rocket flight of a payload containing both dust and electron probes.

  2. In-flight calibration of mesospheric rocket plasma probes.

    Science.gov (United States)

    Havnes, Ove; Hartquist, Thomas W; Kassa, Meseret; Morfill, Gregor E

    2011-07-01

    Many effects and factors can influence the efficiency of a rocket plasma probe. These include payload charging, solar illumination, rocket payload orientation and rotation, and dust impact induced secondary charge production. As a consequence, considerable uncertainties can arise in the determination of the effective cross sections of plasma probes and measured electron and ion densities. We present a new method for calibrating mesospheric rocket plasma probes and obtaining reliable measurements of plasma densities. This method can be used if a payload also carries a probe for measuring the dust charge density. It is based on that a dust probe's effective cross section for measuring the charged component of dust normally is nearly equal to its geometric cross section, and it involves the comparison of variations in the dust charge density measured with the dust detector to the corresponding current variations measured with the electron and/or ion probes. In cases in which the dust charge density is significantly smaller than the electron density, the relation between plasma and dust charge density variations can be simplified and used to infer the effective cross sections of the plasma probes. We illustrate the utility of the method by analysing the data from a specific rocket flight of a payload containing both dust and electron probes.

  3. Review on film cooling of liquid rocket engines

    Directory of Open Access Journals (Sweden)

    S.R. Shine

    2018-03-01

    Full Text Available Film cooling in combination with regenerative cooling is presently considered as an efficient method to guarantee safe operation of liquid rocket engines having higher heat flux densities for long duration. This paper aims to bring all the research carried out in the field of liquid rocket engine film cooling since 1950. The analytical and numerical procedure followed, experimental facilities and measurements made and major inferences drawn are reviewed in detail, and compared where ever possible. Review has been made through a discussion of the analyses methodologies and the factors that influence film cooling performance. An effort has also been made to determine the status of the research, pointing out critical gaps, which are still to be explained and addressed by future generations. Keywords: Heat transfer, Liquid rocket thrust chamber, Film cooling, Cooling effectiveness

  4. The Effect of an Isogrid on Cryogenic Propellant Behavior and Thermal Stratification

    Science.gov (United States)

    Oliveira, Justin; Kirk, Daniel R.; Chintalapati, Sunil; Schallhorn, Paul A.; Piquero, Jorge L.; Campbell, Mike; Chase, Sukhdeep

    2007-01-01

    All models for thermal stratification available in the presentation are derived using smooth, flat plate laminar and turbulent boundary layer models. This study examines the effect of isogrid (roughness elements) on the surface of internal tank walls to mimic the effects of weight-saving isogrid, which is located on the inside of many rocket propellant tanks. Computational Fluid Dynamics (CFD) is used to study the momentum and thermal boundary layer thickness for free convection flows over a wall with generic roughness elements. This presentation makes no mention of actual isogrid sizes or of any specific tank geometry. The magnitude of thermal stratification is compared for smooth and isogrid-lined walls.

  5. Space Shuttle solid rocket booster

    Science.gov (United States)

    Hardy, G. B.

    1979-01-01

    Details of the design, operation, testing and recovery procedures of the reusable solid rocket boosters (SRB) are given. Using a composite PBAN propellant, they will provide the primary thrust (six million pounds maximum at 20 s after ignition) within a 3 g acceleration constraint, as well as thrust vector control for the Space Shuttle. The drogues were tested to a load of 305,000 pounds, and the main parachutes to 205,000. Insulation in the solid rocket motor (SRM) will be provided by asbestos-silica dioxide filled acrylonitrile butadiene rubber ('asbestos filled NBR') except in high erosion areas (principally in the aft dome), where a carbon-filled ethylene propylene diene monomer-neopreme rubber will be utilized. Furthermore, twenty uses for the SRM nozzle will be allowed by its ablative materials, which are principally carbon cloth and silica cloth phenolics.

  6. A technique for rocket-borne detection of electron bunching at megahertz frequencies

    International Nuclear Information System (INIS)

    Gough, M.P.

    1980-01-01

    Energetic electrons precipitating in the auroral ionosphere may be bunched at frequencies up to several megahertz as a result of local wave-particle interactions. A technique is described whereby this megahertz bunching can be observed using conventional rocket-borne energetic electron detectors counting at rates below 10 5 cps. Electron arrival time information is pre-processed on board the rocket and any bunching present can be realized by subsequent computer processing on the ground using only a modest data transmission rate from the rocket. Results of a pilot rocket experiment prove the value of the technique and lead on to formulating the design of a future experiment where the maximum amount of data processing is performed on the rocket. The technique should perform an important diagnostic role, helping us to understand the complex wave-particle interactions occurring in the auroral ionosphere. (orig.)

  7. 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.

  8. Theodore von Karman - Rocket Scientist

    Indian Academy of Sciences (India)

    seminal contributions to several areas of fluid and solid mechanics, as the first head of ... nent position in Aeronautics research, as a pioneer of rocket science in America ... toral work, however, was on the theory of buckling of large structures.

  9. 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

  10. Combustion of metal agglomerates in a solid rocket core flow

    Science.gov (United States)

    Maggi, Filippo; Dossi, Stefano; DeLuca, Luigi T.

    2013-12-01

    The need for access to space may require the use of solid propellants. High thrust and density are appealing features for different applications, spanning from boosting phase to other service applications (separation, de-orbiting, orbit insertion). Aluminum is widely used as a fuel in composite solid rocket motors because metal oxidation increases enthalpy release in combustion chamber and grants higher specific impulse. Combustion process of metal particles is complex and involves aggregation, agglomeration and evolution of reacting particulate inside the core flow of the rocket. It is always stated that residence time should be enough in order to grant complete metal oxidation but agglomerate initial size, rocket grain geometry, burning rate, and other factors have to be reconsidered. New space missions may not require large rocket systems and metal combustion efficiency becomes potentially a key issue to understand whether solid propulsion embodies a viable solution or liquid/hybrid systems are better. A simple model for metal combustion is set up in this paper. Metal particles are represented as single drops trailed by the core flow and reacted according to Beckstead's model. The fluid dynamics is inviscid, incompressible, 1D. The paper presents parametric computations on ideal single-size particles as well as on experimental agglomerate populations as a function of operating rocket conditions and geometries.

  11. Nuclear Thermal Propulsion: Past, Present, and a Look Ahead

    Science.gov (United States)

    Borowski, Stanley K.

    2014-01-01

    NTR: High thrust high specific impulse (2 x LOXLH2 chemical) engine uses high power density fission reactor with enriched uranium fuel as thermal power source. Reactor heat is removed using H2 propellant which is then exhausted to produce thrust. Conventional chemical engine LH2 tanks, turbo pumps, regenerative nozzles and radiation-cooled shirt extensions used -- NTR is next evolutionary step in high performance liquid rocket engines.

  12. 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.

  13. This "Is" Rocket Science!

    Science.gov (United States)

    Keith, Wayne; Martin, Cynthia; Veltkamp, Pamela

    2013-01-01

    Using model rockets to teach physics can be an effective way to engage students in learning. In this paper, we present a curriculum developed in response to an expressed need for helping high school students review physics equations in preparation for a state-mandated exam. This required a mode of teaching that was more advanced and analytical…

  14. ROCKETS: Soar to Success

    Science.gov (United States)

    Brett, Christine E. W.; O'Merle, Mary Jane; White, Gene

    2017-01-01

    This article describes ROCKETS, an after-school program for at-risk youth, and how the university students became involved in this service-learning project. The article discusses the steps that were taken to start the program, what is being done to continue the program, and the challenges that faculty have faced. This program is an authentic…

  15. A summary of results from solar monitoring rocket flights

    Science.gov (United States)

    Duncan, C. H.

    1981-01-01

    Three rocket flights to measure the solar constant and provide calibration data for sensors aboard Nimbus 6, 7, and Solar Maximum Mission (SMM) spacecraft were accomplished. The values obtained by the rocket instruments for the solar constant in SI units are: 1367 w/sq m on 29 June 1976; 1372 w/sq m on 16 November 1978; and 1374 w/sq m on 22 May 1980. The uncertainty of the rocket measurements is + or - 0.5%. The values obtained by the Hickey-Frieden sensor on Nimbus 7 during the second and third flights was 1376 w/sq m. The value obtained by the Active Cavity Radiometer Model IV (ACR IV) on SMM during the flight was 1368 w/sq m.

  16. Parameters Affecting the Erosive Burning of Solid Rocket Motor

    Directory of Open Access Journals (Sweden)

    Abdelaziz Almostafa

    2018-01-01

    Full Text Available Increasing the velocity of gases inside solid rocket motors with low port-to-throat area ratios, leading to increased occurrence and severity of burning rate augmentation due to flow of propellant products across burning propellant surfaces (erosive burning, erosive burning of high energy composite propellant was investigated to supply rocket motor design criteria and to supplement knowledge of combustion phenomena, pressure, burning rate and high velocity of gases all of these are parameters affect on erosive burning. Investigate the phenomena of the erosive burning by using the 2’inch rocket motor and modified one. Different tests applied to fulfil all the parameters that calculated out from the experiments and by studying the pressure time curve and erosive burning phenomena.

  17. Infrared signature modelling of a rocket jet plume - comparison with flight measurements

    International Nuclear Information System (INIS)

    Rialland, V; Perez, P; Roblin, A; Guy, A; Gueyffier, D; Smithson, T

    2016-01-01

    The infrared signature modelling of rocket plumes is a challenging problem involving rocket geometry, propellant composition, combustion modelling, trajectory calculations, fluid mechanics, atmosphere modelling, calculation of gas and particles radiative properties and of radiative transfer through the atmosphere. This paper presents ONERA simulation tools chained together to achieve infrared signature prediction, and the comparison of the estimated and measured signatures of an in-flight rocket plume. We consider the case of a solid rocket motor with aluminized propellant, the Black Brant sounding rocket. The calculation case reproduces the conditions of an experimental rocket launch, performed at White Sands in 1997, for which we obtained high quality infrared signature data sets from DRDC Valcartier. The jet plume is calculated using an in-house CFD software called CEDRE. The plume infrared signature is then computed on the spectral interval 1900-5000 cm -1 with a step of 5 cm -1 . The models and their hypotheses are presented and discussed. Then the resulting plume properties, radiance and spectra are detailed. Finally, the estimated infrared signature is compared with the spectral imaging measurements. The discrepancies are analyzed and discussed. (paper)

  18. Development of a new generation solid rocket motor ignition computer code

    Science.gov (United States)

    Foster, Winfred A., Jr.; Jenkins, Rhonald M.; Ciucci, Alessandro; Johnson, Shelby D.

    1994-01-01

    This report presents the results of experimental and numerical investigations of the flow field in the head-end star grain slots of the Space Shuttle Solid Rocket Motor. This work provided the basis for the development of an improved solid rocket motor ignition transient code which is also described in this report. The correlation between the experimental and numerical results is excellent and provides a firm basis for the development of a fully three-dimensional solid rocket motor ignition transient computer code.

  19. Rocket motors incorporating basalt fiber and nanoclay compositions and methods of insulating a rocket motor with the same

    Science.gov (United States)

    Gajiwala, Himansu M. (Inventor)

    2011-01-01

    An insulation composition that comprises at least one nitrile butadiene rubber, basalt fibers, and nanoclay is disclosed. Further disclosed is an insulation composition that comprises polybenzimidazole fibers, basalt fibers, and nanoclay. The basalt fibers may be present in the insulation compositions in a range of from approximately 1% by weight to approximately 6% by weight of the total weight of the insulation composition. The nanoclay may be present in the insulation compositions in a range of from approximately 5% by weight to approximately 10% by weight of the total weight of the insulation composition. Rocket motors including the insulation compositions and methods of insulating a rocket motor are also disclosed.

  20. Observation and simulation of the ionosphere disturbance waves triggered by rocket exhausts

    Science.gov (United States)

    Lin, Charles C. H.; Chen, Chia-Hung; Matsumura, Mitsuru; Lin, Jia-Ting; Kakinami, Yoshihiro

    2017-08-01

    Observations and theoretical modeling of the ionospheric disturbance waves generated by rocket launches are investigated. During the rocket passage, time rate change of total electron content (rTEC) enhancement with the V-shape shock wave signature is commonly observed, followed by acoustic wave disturbances and region of negative rTEC centered along the trajectory. Ten to fifteen min after the rocket passage, delayed disturbance waves appeared and propagated along direction normal to the V-shape wavefronts. These observation features appeared most prominently in the 2016 North Korea rocket launch showing a very distinct V-shape rTEC enhancement over enormous areas along the southeast flight trajectory despite that it was also appeared in the 2009 North Korea rocket launch with the eastward flight trajectory. Numerical simulations using the physical-based nonlinear and nonhydrostatic coupled model of neutral atmosphere and ionosphere reproduce promised results in qualitative agreement with the characteristics of ionospheric disturbance waves observed in the 2009 event by considering the released energy of the rocket exhaust as the disturbance source. Simulations reproduce the shock wave signature of electron density enhancement, acoustic wave disturbances, the electron density depletion due to the rocket-induced pressure bulge, and the delayed disturbance waves. The pressure bulge results in outward neutral wind flows carrying neutrals and plasma away from it and leading to electron density depletions. Simulations further show, for the first time, that the delayed disturbance waves are produced by the surface reflection of the earlier arrival acoustic wave disturbances.

  1. A spacecraft charging study on the SCEX 3 rocket

    International Nuclear Information System (INIS)

    Mullen, E.G.; Gussenhoven, M.S.; Hardy, D.A.; Murphy, G.P.; Lloyd, J.W.F.; Slutter, W.; Malcolm, P.; Kellogg, P.J.; Monson, S.

    1991-01-01

    Instruments on the SCEX 3 rocket payload flown from the Poker Flats Rocket Range in February 1990 were used to study charging during electron beam emissions. This paper reports that the data show that electrostatic analyzers can be used to measure vehicle charging and direct beam return currents in dense plasma conditions. The data also show return current dependencies on pitch angle, beam current and beam energy

  2. The seven secrets of how to think like a rocket scientist

    CERN Document Server

    Longuski, James

    2007-01-01

    This book explains the methods that rocket scientists use - expressed in a way that could be applied in everyday life. It's short and snappy and written by a rocket scientist. It is intended for general "armchair" scientists.

  3. History of the development of rocket technology and astronautics in Poland

    Science.gov (United States)

    Geisler, W.

    1977-01-01

    The development of rocket technology in Poland is outlined. The history cites 13th century use of war rockets in combating Tartars as well as 20th century studies of the future and reality of space flights.

  4. 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

  5. Results from a tethered rocket experiment (Charge-2)

    Science.gov (United States)

    Kawashima, N.; Sasaki, S.; Oyama, K. I.; Hirao, K.; Obayashi, T.; Raitt, W. J.; White, A. B.; Williamson, P. R.; Banks, P. M.; Sharp, W. F.

    A tethered payload experiment (Charge-2) was carried out as an international program between Japan and the USA using a NASA sounding rocket at White Sands Missile Range. The objective of the experiment was to perform a new type of active experiment in space by injecting an electron beam from a mother-daughter rocket system connected with a long tether wire. The electron beam with voltage and current up to 1 kV and 80 mA (nominal) was injected from the mother payload. An insulated conductive wire of 426 m length connected the two payloads, the longest tether system flown so far. The electron gun system and diagnostic instruments (plasma, optical, particle and wave) functioned correctly throughout the flight. The potential rise of the mother payload during the electron beam emission was measured with respect to the daughter payload. The beam trajectory was detected by a camera onboard the mother rocket. Wave generation and current induction in the wire during the beam emission were also studied.

  6. Thiokol Solid Rocket Motors

    Science.gov (United States)

    Graves, S. R.

    2000-01-01

    This paper presents viewgraphs on thiokol solid rocket motors. The topics include: 1) Communications; 2) Military and government intelligence; 3) Positioning satellites; 4) Remote sensing; 5) Space burial; 6) Science; 7) Space manufacturing; 8) Advertising; 9) Space rescue space debris management; 10) Space tourism; 11) Space settlements; 12) Hazardous waste disposal; 13) Extraterrestrial resources; 14) Fast package delivery; and 15) Space utilities.

  7. Effect of Six Missile-Bay Baffle Configurations and a Rocket End Plate on Ejection Releases of an MB-1 Rocket from a 0.05 Scale Model of the Convair F-106A Airplane

    Science.gov (United States)

    Hinson, William F.; Lee, John B.

    1959-01-01

    As a continuation of an investigation of the release characteristics of an MB-1 rocket carried internally by the Convair F-106A airplane, six missile-bay baffle configurations and a rocket end plate have been investigated in the 27- by 27-inch preflight jet of the NASA Wallops Station. The MB-1 rocket used had retractable fins and was ejected from a missile bay modified by the addition of six different baffle configurations. For some tests a rocket end plate was added to the model. Dynamically scaled models (0.04956 scale) were tested at a simulated altitude of 22,450 feet and Mach numbers of 0.86, 1.59, and 1.98, and at a simulated altitude of 29,450 feet and a Mach number of 1.98. The results of this investigation indicate that the missile-bay baffle configurations and the rocket end plate may be used to reduce the positive pitch amplitude of the MB-1 rocket after release. The initial negative pitching velocity applied to the MB-1 rocket might then be reduced in order to maintain a near-level-flight attitude after release. As the fuselage angle of attack is increased, the negative pitch amplitude of the rocket is decreased.

  8. Mean Flow Augmented Acoustics in Rocket Systems

    Science.gov (United States)

    Fischbach, Sean R.

    2014-01-01

    Oscillatory motion in solid rocket motors and liquid engines has long been a subject of concern. Many rockets display violent fluctuations in pressure, velocity, and temperature originating from the complex interactions between the combustion process and gas dynamics. The customary approach to modeling acoustic waves inside a rocket chamber is to apply the classical inhomogeneous wave equation to the combustion gas. The assumption of a linear, non-dissipative wave in a quiescent fluid remains valid while the acoustic amplitudes are small and local gas velocities stay below Mach 0.2. The converging section of a rocket nozzle, where gradients in pressure, density, and velocity become large, is a notable region where this approach is not applicable. The expulsion of unsteady energy through the nozzle of a rocket is identified as the predominate source of acoustic damping for most rocket systems. An accurate model of the acoustic behavior within this region where acoustic modes are influenced by the presence of a steady mean flow is required for reliable stability predictions. Recently, an approach to address nozzle damping with mean flow effects was implemented by French [1]. This new approach extends the work originated by Sigman and Zinn [2] by solving the acoustic velocity potential equation (AVPE) formulated by perturbing the Euler equations [3]. The acoustic velocity potential (psi) describing the acoustic wave motion in the presence of an inhomogeneous steady high-speed flow is defined by, (del squared)(psi) - (lambda/c)(exp 2)(psi) - M(dot)[M(dot)(del)(del(psi))] - 2(lambda(M/c) + (M(dot)del(M))(dot)del(psi)-2(lambda)(psi)[M(dot)del(1/c)]=0 (1) with M as the Mach vector, c as the speed of sound, and lambda as the complex eigenvalue. French apply the finite volume method to solve the steady flow field within the combustion chamber and nozzle with inviscid walls. The complex eigenvalues and eigenvector are determined with the use of the ARPACK eigensolver. The

  9. 78 FR 40196 - National Environmental Policy Act; Sounding Rockets Program; Poker Flat Research Range

    Science.gov (United States)

    2013-07-03

    ...; Sounding Rockets Program; Poker Flat Research Range AGENCY: National Aeronautics and Space Administration... Sounding Rockets Program (SRP) at Poker Flat Research Range (PFRR), Alaska. SUMMARY: Pursuant to the... government agencies, and educational institutions have conducted suborbital rocket launches from the PFRR...

  10. Applied algorithm in the liner inspection of solid rocket motors

    Science.gov (United States)

    Hoffmann, Luiz Felipe Simões; Bizarria, Francisco Carlos Parquet; Bizarria, José Walter Parquet

    2018-03-01

    In rocket motors, the bonding between the solid propellant and thermal insulation is accomplished by a thin adhesive layer, known as liner. The liner application method involves a complex sequence of tasks, which includes in its final stage, the surface integrity inspection. Nowadays in Brazil, an expert carries out a thorough visual inspection to detect defects on the liner surface that may compromise the propellant interface bonding. Therefore, this paper proposes an algorithm that uses the photometric stereo technique and the K-nearest neighbor (KNN) classifier to assist the expert in the surface inspection. Photometric stereo allows the surface information recovery of the test images, while the KNN method enables image pixels classification into two classes: non-defect and defect. Tests performed on a computer vision based prototype validate the algorithm. The positive results suggest that the algorithm is feasible and when implemented in a real scenario, will be able to help the expert in detecting defective areas on the liner surface.

  11. Water Rockets. Get Funny With Newton's Laws

    Directory of Open Access Journals (Sweden)

    Manuel Roca Vicent

    2017-01-01

    Full Text Available The study of the movement of the rocket has been used for decades to encourage students in the study of physics. This system has an undeniable interest to introduce concepts such as properties of gases, laws of Newton,  exchange  between  different  types  of  energy  and  its  conservation  or fluid  mechanics.  Our  works has  been  to  build  and  launch  these  rockets  in  different  educational  levels  and  in  each  of  these  ones  have introduced  the  part  of  Physics  more  suited  to  the  knowledge  of  our  students.  The  aim  of  the  learning experience  is  to  launch  the  rocket  as  far  as  possible  and  learn  to  predict  the  travelled  distance,  using Newton's  laws  and fluid  mechanics.  After  experimentation  we  demonstrated  to  be  able  to  control  the parameters that improve the performance of our rocket, such as the  fill factor, the volume and mass of the empty  bottle,  liquid  density,  launch  angle,  pressure  prior  air  release.  In addition, it is a fun experience can be attached to all levels of education in primary and high school.

  12. Sounding rocket experiments during the IMS period at Syowa Station, Antarctica

    International Nuclear Information System (INIS)

    Hirasawa, T.; Nagata, T.

    1979-01-01

    During IMS Period, 19 sounding rockets were launched into auroras at various stages of polar substorms from Syowa Station (Geomag. lat. = -69.6 0 , Geomag. log. = 77.1 0 ), Antarctica. Through the successful rocket flights, the significant physical quantities in auroras were obtained: 19 profiles of electron density and temperature, 11 energy spectra of precipitating electrons, 15 frequency spectra of VLF and HF plasma waves and 4 vertical profiles of electric and magnetic fields. These rocket data have been analyzed and compared with the coordinated ground-based observation data for studies of polar substorms. (author)

  13. Night Airglow Observations from Orbiting Spacecraft Compared with Measurements from Rockets.

    Science.gov (United States)

    Koomen, M J; Gulledge, I S; Packer, D M; Tousey, R

    1963-06-07

    A luminous band around the night-time horizon, observed from orbiting capsules by J. H. Glenn and M. S. Carpenter, and identified as the horizon enhancement of the night airglow, is detected regularly in rocket-borne studies of night airglow. Values of luminance and dip angle of this band derived from Carpenter's observations agree remarkably well with values obtained from rocket data. The rocket results, however, do not support Carpenter's observation that the emission which he saw was largely the atomic oxygen line at 5577 A, but assign the principal luminosity to the green continuum.

  14. The efficient future of deep-space travel - electric rockets; Das Zeitalter der Elektrischen Raketen

    Energy Technology Data Exchange (ETDEWEB)

    Choueiri, Edgar Y. [Princeton Univ., NJ (United States). Electric Propulsion and Plasma Dynamics Lab.

    2010-01-15

    Conventional rockets generate thrust by burning chemical fuel. Electric rockets propel space vehicles by applying electric or electromagnetic fields to clouds of charged particles, or plasmas, to accelerate them. Although electric rockets offer much lower thrust levels than their chemical cousins, they can eventually enable spacecraft to reach greater speeds for the same amount of propellant. Electric rockets' high-speed capabilities and their efficient use of propellant make them valuable for deep-space missions. (orig.)

  15. Integration of Flex Nozzle System and Electro Hydraulic Actuators to Solid Rocket Motors

    Science.gov (United States)

    Nayani, Kishore Nath; Bajaj, Dinesh Kumar

    2017-10-01

    A rocket motor assembly comprised of solid rocket motor and flex nozzle system. Integration of flex nozzle system and hydraulic actuators to the solid rocket motors are done after transportation to the required place where integration occurred. The flex nozzle system is integrated to the rocket motor in horizontal condition and the electro hydraulic actuators are assembled to the flex nozzle systems. The electro hydraulic actuators are connected to the hydraulic power pack to operate the actuators. The nozzle-motor critical interface are insulation diametrical compression, inhibition resin-28, insulation facial compression, shaft seal `O' ring compression and face seal `O' ring compression.

  16. The electromagnetic rocket gun - a means to reach ultrahigh velocities

    International Nuclear Information System (INIS)

    Winterberg, F.

    1983-01-01

    A novel kind of electromagnetic launcher for the acceleration of multigram-size macroparticles, up to velocities required for impact fusion, is proposed. The novel launcher concept combines the efficiency of a gun with the much higher velocities attainable by a rocket. In the proposed concept a rocket-like projectile is launched inside a gun barrel, drawing its energy from a travelling magnetic wave. The travelling magnetic wave heats and ionizes the exhaust jet of the rocket. As a result, the projectile i propelled both by the recoil from the jet and the magnetic pressure of the travelling magnetic wave. In comparison to magnetic linear accelerators, accelerating either superconducting or ferromagnetic projectiles, the proposed concept has several important advantages. First, the exhaust jet is much longer than the rocket-like projectile and which permits a much longer switching time to turn on the travelling magnetic wave. Second, the proposed concept does not require superconducting projectiles, or projectiles made from expensive ferromagnetic material. Third, unlike in railgun accelerators, the projectile can be kept away from the wall, and thereby can reach much larger velocities. (orig.)

  17. 76 FR 20715 - National Environmental Policy Act; Sounding Rockets Program; Poker Flat Research Range

    Science.gov (United States)

    2011-04-13

    ...; Sounding Rockets Program; Poker Flat Research Range AGENCY: National Aeronautics and Space Administration... continuing sounding rocket operations at Poker Flat Research Range (PFRR), Alaska. SUMMARY: Pursuant to the... information about NASA's Sounding Rocket Program (SRP) and the University of Alaska-Fairbanks' PFRR may be...

  18. Modeling and validation of Ku-band signal attenuation through rocket plumes

    NARCIS (Netherlands)

    Veek, van der B.J.; Chintalapati, S.; Kirk, D.R.; Gutierrez, H.; Bun, R.F.

    2013-01-01

    Communications to and from a launch vehicle during ascent are of critical importance to the success of rocket-launch operations. During ascent, the rocket's exhaust plume causes significant interference in the radio communications between the vehicle and ground station. This paper presents an

  19. Effect of population density of lettuce intercropped with rocket on productivity and land-use efficiency

    Science.gov (United States)

    2018-01-01

    The objective of this study was to evaluate the influence of the spacing of lettuce rows on the production of a lettuce-rocket intercropping system over two growing seasons (11 August to 25 September 2011 and 12 January to 24 February 2012) in Jaboticabal, São Paulo, Brazil. We evaluated 11 treatments in each season: lettuce-rocket intercrops with five row spacings for the lettuce (0.20, 0.25, 0.30, 0.35 and 0.40 m) and the rocket planted midway between the lettuce rows, sole crops of lettuce at the same five row spacings and a sole crop of rocket. Fresh and dry masses of the lettuce and rocket and number of lettuce leaves per plant were highest with a lettuce row spacing of 0.40 m, but the productivities of the lettuce and rocket were higher with a lettuce row spacing of 0.20 m. The productivities and fresh and dry weights of the lettuce and rocket and the number of lettuce leaves per plant were highest in the sole crops, but the fresh and dry weights of the rocket were higher with intercropping. The land equivalent ratios were >1.0 in both seasons in all intercrops and were highest for the densest crop (1.41). Intercropping was therefore 41% more efficient than sole cropping for the production of lettuce and rocket. PMID:29698401

  20. The German scientific balloon and sounding rocket programme

    International Nuclear Information System (INIS)

    Dahl, A.F.

    1980-01-01

    This report contains information on sounding rocket projects in the scientific field of astronomy, aeronomy, magnetosphere, and material science under microgravity. The scientific balloon projects are performed with emphasis on astronomical research. By means of tables it is attempted to give a survey, as complete as possible, of the projects the time since the last symposium in Ajaccio, Corsica, and of preparations and plans for the future until 1983. The scientific balloon and sounding rocket projects form a small successful part of the German space research programme. (Auth.)

  1. Radiation effects on thermal decomposition of inorganic solids

    International Nuclear Information System (INIS)

    Dedgaonkar, V.G.

    1985-01-01

    Radiation effects on the thermal decomposition characteristics of inorganic oxyanions like permanganates, nitrates, zeolites and particularly ammonium perchlorate (AP) have been highlighted.The last compound finds wide application as an oxidizer in solid rocket propellents and although several hundred papers have been published on it during the last 30-40 years, most of which from the point of view of understanding and controlling the decomposition behaviour, there are only a few reports available in this area following the radiation treatment. (author)

  2. Parameters Affecting the Erosive Burning of Solid Rocket Motor

    OpenAIRE

    Abdelaziz Almostafa; Guozhu Liang; Elsayed Anwer

    2018-01-01

    Increasing the velocity of gases inside solid rocket motors with low port-to-throat area ratios, leading to increased occurrence and severity of burning rate augmentation due to flow of propellant products across burning propellant surfaces (erosive burning), erosive burning of high energy composite propellant was investigated to supply rocket motor design criteria and to supplement knowledge of combustion phenomena, pressure, burning rate and high velocity of gases all of these are parameter...

  3. 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.

  4. Closure Letter Report for Corrective Action Unit 496: Buried Rocket Site - Antelope Lake (TTR)

    International Nuclear Information System (INIS)

    NSTec Environmental Restoration

    2007-01-01

    A Streamlined Approach for Environmental Restoration (SAFER) Plan for investigation and closure of CAU 496, Corrective Action Site (CAS) TA-55-008-TAAL (Buried Rocket), at the Tonopah Test Range (TTR), was approved by the Nevada Department of Environmental Protection (NDEP) on July 21,2004. Approval to transfer CAS TA-55-008-TAAL from CAU 496 to CAU 4000 (No Further Action Sites) was approved by NDEP on December 21, 2005, based on the assumption that the rocket did not present any environmental concern. The approval letter included the following condition: ''NDEP understands, from the NNSA/NSO letter dated November 30,2005, that a search will be conducted for the rocket during the planned characterization of other sites at the Tonopah Test Range and, if found, the rocket will be removed as a housekeeping measure''. NDEP and U.S. Department of Energy, National Nuclear Security Administration Nevada Site Office personnel located the rocket on Mid Lake during a site visit to TTR, and a request to transfer CAS TA-55-008-TAAL from CAU 4000 back to CAU 496 was approved by NDEP on September 11,2006. CAS TA-55-008-TAAL was added to the ''Federal Facility Agreement and Consent Order'' of 1996, based on an interview with a retired TTR worker in 1993. The original interview documented that a rocket was launched from Area 9 to Antelope Lake and was never recovered due to the high frequency of rocket tests being conducted during this timeframe. The interviewee recalled the rocket being an M-55 or N-55 (the M-50 ''Honest John'' rocket was used extensively at TTR from the 1960s to early 1980s). A review of previously conducted interviews with former TTR personnel indicated that the interviewees confused information from several sites. The location of the CAU 496 rocket on Mid Lake is directly south of the TTR rocket launch facility in Area 9 and is consistent with information gathered on the lost rocket during recent interviews. Most pertinently, an interview in 2005 with a

  5. The interaction of bubbles with solidification interfaces. [during coasting phase of sounding rocket flight

    Science.gov (United States)

    Papazian, J. M.; Wilcox, W. R.

    1977-01-01

    The behavior of bubbles at a dendritic solidification interface was studied during the coasting phase of a sounding rocket flight. Sequential photographs of the gradient freeze experiment showed nucleation, growth and coalescence of bubbles at the moving interface during both the low-gravity and one-gravity tests. In the one-gravity test the bubbles were observed to detach from the interface and float to the top of the melt. However, in the low-gravity tests no bubble detachment from the interface or steady state bubble motion occurred and large voids were grown into the crystal. These observations are discussed in terms of the current theory of thermal migration of bubbles and in terms of their implications on the space processing of metals.

  6. Real-Time Inhibitor Recession Measurements in the Space Shuttle Reusable Solid Rocket Motors

    Science.gov (United States)

    McWhorter, Bruce B.; Ewing, Mark E.; McCool, Alex (Technical Monitor)

    2001-01-01

    Real-time char line recession measurements were made on propellant inhibitors of the Space Shuttle Reusable Solid Rocket Motor (RSRM). The RSRM FSM-8 static test motor propellant inhibitors (composed of a rubber insulation material) were successfully instrumented with eroding potentiometers and thermocouples. The data was used to establish inhibitor recession versus time relationships. Normally, pre-fire and post-fire insulation thickness measurements establish the thermal performance of an ablating insulation material. However, post-fire inhibitor decomposition and recession measurements are complicated by the fact that most of the inhibitor is back during motor operation. It is therefore a difficult task to evaluate the thermal protection offered by the inhibitor material. Real-time measurements would help this task. The instrumentation program for this static test motor marks the first time that real-time inhibitors. This report presents that data for the center and aft field joint forward facing inhibitors. The data was primarily used to measure char line recession of the forward face of the inhibitors which provides inhibitor thickness reduction versus time data. The data was also used to estimate the inhibitor height versus time relationship during motor operation.

  7. Multi-Parameter Wireless Monitoring and Telecommand of a Rocket Payload: Design and Implementation

    Science.gov (United States)

    Pamungkas, Arga C.; Putra, Alma A.; Puspitaningayu, Pradini; Fransisca, Yulia; Widodo, Arif

    2018-04-01

    A rocket system generally consists of two parts, the rocket motor and the payload. The payload system is built of several sensors such as accelerometer, gyroscope, magnetometer, and also a surveillance camera. These sensors are used to monitor the rocket in a three-dimensional axis which determine its attitude. Additionally, the payload must be able to perform image capturing in a certain distance using telecommand. This article is intended to describe the design and also the implementation of a rocket payload which has attitude monitoring and telecommand ability from the ground control station using a long-range wireless module Digi XBee Pro 900 HP.

  8. 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.

  9. Review of Nuclear Thermal Propulsion Ground Test Options

    Science.gov (United States)

    Coote, David J.; Power, Kevin P.; Gerrish, Harold P.; Doughty, Glen

    2015-01-01

    High efficiency rocket propulsion systems are essential for humanity to venture beyond the moon. Nuclear Thermal Propulsion (NTP) is a promising alternative to conventional chemical rockets with relatively high thrust and twice the efficiency of highest performing chemical propellant engines. NTP utilizes the coolant of a nuclear reactor to produce propulsive thrust. An NTP engine produces thrust by flowing hydrogen through a nuclear reactor to cool the reactor, heating the hydrogen and expelling it through a rocket nozzle. The hot gaseous hydrogen is nominally expected to be free of radioactive byproducts from the nuclear reactor; however, it has the potential to be contaminated due to off-nominal engine reactor performance. NTP ground testing is more difficult than chemical engine testing since current environmental regulations do not allow/permit open air testing of NTP as was done in the 1960's and 1970's for the Rover/NERVA program. A new and innovative approach to rocket engine ground test is required to mitigate the unique health and safety risks associated with the potential entrainment of radioactive waste from the NTP engine reactor core into the engine exhaust. Several studies have been conducted since the ROVER/NERVA program in the 1970's investigating NTP engine ground test options to understand the technical feasibility, identify technical challenges and associated risks and provide rough order of magnitude cost estimates for facility development and test operations. The options can be divided into two distinct schemes; (1) real-time filtering of the engine exhaust and its release to the environment or (2) capture and storage of engine exhaust for subsequent processing.

  10. Theoretical and Experimental Analysis of the Physics of Water Rockets

    Science.gov (United States)

    Barrio-Perotti, R.; Blanco-Marigorta, E.; Fernandez-Francos, J.; Galdo-Vega, M.

    2010-01-01

    A simple rocket can be made using a plastic bottle filled with a volume of water and pressurized air. When opened, the air pressure pushes the water out of the bottle. This causes an increase in the bottle momentum so that it can be propelled to fairly long distances or heights. Water rockets are widely used as an educational activity, and several…

  11. Integral performance optimum design for multistage solid propellant rocket motors

    Energy Technology Data Exchange (ETDEWEB)

    Zhang, Hongtao (Shaanxi Power Machinery Institute (China))

    1989-04-01

    A mathematical model for integral performance optimization of multistage solid propellant rocket motors is presented. A calculation on a three-stage, volume-fixed, solid propellant rocket motor is used as an example. It is shown that the velocity at burnout of intermediate-range or long-range ballistic missile calculated using this model is four percent greater than that using the usual empirical method.

  12. Data Analysis of the TK-1G Sounding Rocket Installed with a Satellite Navigation System

    Directory of Open Access Journals (Sweden)

    Lesong Zhou

    2017-10-01

    Full Text Available This article gives an in-depth analysis of the experimental data of the TK-1G sounding rocket installed with the satellite navigation system. It turns out that the data acquisition rate of the rocket sonde is high, making the collection of complete trajectory and meteorological data possible. By comparing the rocket sonde measurements with those obtained by virtue of other methods, we find that the rocket sonde can be relatively precise in measuring atmospheric parameters within the scope of 20–60 km above the ground. This establishes the fact that the TK-1G sounding rocket system is effective in detecting near-space atmospheric environment.

  13. Chemical rocket propulsion a comprehensive survey of energetic materials

    CERN Document Server

    Shimada, Toru; Sinditskii, Valery; Calabro, Max

    2017-01-01

    Developed and expanded from the work presented at the New Energetic Materials and Propulsion Techniques for Space Exploration workshop in June 2014, this book contains new scientific results, up-to-date reviews, and inspiring perspectives in a number of areas related to the energetic aspects of chemical rocket propulsion. This collection covers the entire life of energetic materials from their conceptual formulation to practical manufacturing; it includes coverage of theoretical and experimental ballistics, performance properties, as well as laboratory-scale and full system-scale, handling, hazards, environment, ageing, and disposal. Chemical Rocket Propulsion is a unique work, where a selection of accomplished experts from the pioneering era of space propulsion and current technologists from the most advanced international laboratories discuss the future of chemical rocket propulsion for access to, and exploration of, space. It will be of interest to both postgraduate and final-year undergraduate students in...

  14. Electromechanical Dynamics Simulations of Superconducting LSM Rocket Launcher System in Attractive-Mode

    Science.gov (United States)

    Yoshida, Kinjiro; Hayashi, Kengo; Takami, Hiroshi

    1996-01-01

    Further feasibility study on a superconducting linear synchronous motor (LSM) rocket launcher system is presented on the basis of dynamic simulations of electric power, efficiency and power factor as well as the ascending motions of the launcher and rocket. The advantages of attractive-mode operation are found from comparison with repulsive-mode operation. It is made clear that the LSM rocket launcher system, of which the long-stator is divided optimally into 60 sections according to launcher speeds, can obtain high efficiency and power factor.

  15. Numerical Calculation of Effect of Elastic Deformation on Aerodynamic Characteristics of a Rocket

    OpenAIRE

    Abbas, Laith K.; Chen, Dongyang; Rui, Xiaoting

    2014-01-01

    The application and workflow of Computational Fluid Dynamics (CFD)/Computational Structure Dynamics (CSD) on solving the static aeroelastic problem of a slender rocket are introduced. To predict static aeroelastic behavior accurately, two-way coupling and inertia relief methods are used to calculate the static deformations and aerodynamic characteristics of the deformed rocket. The aerodynamic coefficients of rigid rocket are computed firstly and compared with the experimental data, which ver...

  16. The German scientific balloon and sounding rocket projects

    International Nuclear Information System (INIS)

    Dalh, A.F.

    1978-01-01

    This report contains information on the sounding rocket projects: experiment preparation for spacelab (astronomy), aeronomy, magnetosphere, and material science. Except for material science the scientific balloon projects are performed in the some scientific fields, but with a strong emphasis on astronomical research. It is tried to provide by means of tables a survey as complete as possible of the projects for the time since the last symposium in Elmau and of the plans for the future until 1981. The scientific balloon and sounding rocket projects form a small succesful part of the German space research programme. (author)

  17. Acoustic field generated by flight of rocket at the Earth surface

    International Nuclear Information System (INIS)

    Drobzheva, Ya.V.; Krasnov, V.M.; Maslov, A.N.

    2006-01-01

    In this paper we present a model, which describes the propagation of acoustic impulses produced by explosion of carrier rocket at the active part of trajectory, down through the atmosphere. Calculations of acoustic field parameters on the earth surface were made for altitudes of rocket flight from 2.8 to 92.3 km and yield of explosions from 0.001 to 0.5 t tnt. It was shown the infrasound accompaniment of rocket flight with the goal to register the explosion it is possible only for an altitude about 70 km. For this case, test set should be situated at the distance not exceeding 120 km from the starting place. (author)

  18. The Temperature of the Dimethylhydrazine Drops Moving in the Atmosphere after Depressurization of the Fuel Tank Rockets

    Directory of Open Access Journals (Sweden)

    Bulba Elena

    2016-01-01

    Full Text Available This work includes the results of the numerical modeling of temperature changes process of the dimethylhydrazine (DMH drops, taking into account the radial temperature gradient in the air after the depressurization of the fuel compartments rockets at high altitude. There is formulated a mathematical model describing the process of DMH drops thermal state modifying when it's moving to the Earth's surface. There is the evaluation of the influence of the characteristic size of heptyl drops on the temperature distribution. It's established that the temperatures of the small size droplets practically completely coincide with the distribution of temperature in the atmosphere at altitudes of up to 40 kilometers.

  19. A Low Cost GPS System for Real-Time Tracking of Sounding Rockets

    Science.gov (United States)

    Markgraf, M.; Montenbruck, O.; Hassenpflug, F.; Turner, P.; Bull, B.; Bauer, Frank (Technical Monitor)

    2001-01-01

    This paper describes the development as well as the on-ground and the in-flight evaluation of a low cost Global Positioning System (GPS) system for real-time tracking of sounding rockets. The flight unit comprises a modified ORION GPS receiver and a newly designed switchable antenna system composed of a helical antenna in the rocket tip and a dual-blade antenna combination attached to the body of the service module. Aside from the flight hardware a PC based terminal program has been developed to monitor the GPS data and graphically displays the rocket's path during the flight. In addition an Instantaneous Impact Point (IIP) prediction is performed based on the received position and velocity information. In preparation for ESA's Maxus-4 mission, a sounding rocket test flight was carried out at Esrange, Kiruna, on 19 Feb. 2001 to validate existing ground facilities and range safety installations. Due to the absence of a dedicated scientific payload, the flight offered the opportunity to test multiple GPS receivers and assess their performance for the tracking of sounding rockets. In addition to the ORION receiver, an Ashtech G12 HDMA receiver and a BAE (Canadian Marconi) Allstar receiver, both connected to a wrap-around antenna, have been flown on the same rocket as part of an independent experiment provided by the Goddard Space Flight Center. This allows an in-depth verification and trade-off of different receiver and antenna concepts.

  20. A six degree-of-freedom thrust sensor for a labscale hybrid rocket

    Science.gov (United States)

    Wright, Ann M.; Wright, Andrew B.; Born, Traig; Strickland, Ryan

    2013-12-01

    A six degree-of-freedom thrust sensor was designed, constructed, calibrated, and tested using the labscale hybrid rocket at the University of Arkansas at Little Rock. The system consisted of six independent legs: one parallel to the axis of symmetry of the rocket for main thrust measurement, two vertical legs near the nozzle end of the rocket, one vertical leg near the oxygen input end of the rocket, and two separated horizontal legs near the nozzle end. Each leg was composed of a rotational bearing, a load cell, and a universal joint above and below the load cell. The leg was designed to create point contact along only one direction and minimize the non-axial forces applied to the load cell. With this system, force in each direction and moments for roll, pitch, and yaw can be measured. The system was calibrated and tested using a labscale hybrid rocket using gaseous oxygen and hydroxyl-terminated polybutadiene solid fuel. The thrust stand proved to be stable during calibration tests. Thrust force vector components and roll, pitch, and yaw moments were calculated for test firings with an oxygen mass flow rate range of 0.0174-0.0348 kg s-1.

  1. Nuclear thermal propulsion transportation systems for lunar/Mars exploration

    International Nuclear Information System (INIS)

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

    1992-09-01

    Nuclear thermal propulsion technology development is underway at NASA and DoE for Space Exploration Initiative (SEI) missions to Mars, with initial near-earth flights to validate flight readiness. Several reactor concepts are being considered for these missions, and important selection criteria will be evaluated before final selection of a system. These criteria include: safety and reliability, technical risk, cost, and performance, in that order. Of the concepts evaluated to date, the Nuclear Engine for Rocket Vehicle Applications (NERVA) derivative (NDR) is the only concept that has demonstrated full power, life, and performance in actual reactor tests. Other concepts will require significant design work and must demonstrate proof-of-concept. Technical risk, and hence, development cost should therefore be lowest for the concept, and the NDR concept is currently being considered for the initial SEI missions. As lighter weight, higher performance systems are developed and validated, including appropriate safety and astronaut-rating requirements, they will be considered to support future SEI application. A space transportation system using a modular nuclear thermal rocket (NTR) system for lunar and Mars missions is expected to result in significant life cycle cost savings. Finally, several key issues remain for NTR's, including public acceptance and operational issues. Nonetheless, NTR's are believed to be the next generation of space propulsion systems - the key to space exploration

  2. An X-ray Experiment with Two-Stage Korean Sounding Rocket

    Directory of Open Access Journals (Sweden)

    Uk-Won Nam

    1998-12-01

    Full Text Available The test result of the X-ray observation system is presented which have been developed at Korea Astronomy Observatory for 3 years (1995-1997. The instrument, which is composed of detector and signal processing parts, is designed for the future observations of compact X-ray sources. The performance of the instrument was tested by mounting on the two-stage Korean Sounding Rocket, which was launched from Taean rocket flight center on June 11 at 10:00 KST 1998. Telemetry data were received from individual parts of the instrument for 32 and 55.7 sec, respectively, since the launch of the rocket. In this paper, the result of the data analysis based on the telemetry data and discussion about the performance of the instrument is reported.

  3. LOX/Methane Regeneratively-Cooled Rocket Engine Development

    Data.gov (United States)

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

  4. 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.

  5. An introduction to the water recovery x-ray rocket

    Science.gov (United States)

    Miles, Drew M.; McEntaffer, Randall L.; Schultz, Ted B.; Donovan, Benjamin D.; Tutt, James H.; Yastishock, Daniel; Steiner, Tyler; Hillman, Christopher R.; McCoy, Jake A.; Wages, Mitchell; Hull, Sam; Falcone, Abe; Burrows, David N.; Chattopadhyay, Tanmoy; Anderson, Tyler; McQuaide, Maria

    2017-08-01

    The Water Recovery X-ray Rocket (WRXR) is a sounding rocket payload that will launch from the Kwajalein Atoll in April 2018 and seeks to be the first astrophysics sounding rocket payload to be water recovered by NASA. WRXR's primary instrument is a grating spectrometer that consists of a mechanical collimator, X-ray reflection gratings, grazing-incidence mirrors, and a hybrid CMOS detector. The instrument will obtain a spectrum of the diffuse soft X-ray emission from the northern part of the Vela supernova remnant and is optimized for 3rd and 4th order OVII emission. Utilizing a field of view of 3.25° × 3.25° and resolving power of λ/δλ ≍40-50 in the lines of interest, the WRXR spectrometer aims to achieve the most highly-resolved spectrum of Vela's diffuse soft X-ray emission. This paper presents introductions to the payload and the science target.

  6. Development of the Hawk/Nike Hawk sounding rocket vehicles

    Science.gov (United States)

    Flowers, B. J.

    1976-01-01

    A new sounding rocket family, the Hawk and Nike-Hawk Vehicles, have been developed, flight tested and added to the NASA Sounding Rocket Vehicle Stable. The Hawk is a single-stage vehicle that will carry 35.6 cm diameter payloads weighing 45.5 kg to 91 kg to altitudes of 78 km to 56 km, respectively. The two-stage Nike-Hawk will carry payloads weighing 68 kg to 136 kg to altitudes of 118 km to 113 km, respectively. Both vehicles utilize the XM22E8 Hawk rocket motor which is available in large numbers as a surplus item from the U.S. Army. The Hawk fin and tail can hardware were designed in-house. The Nike tail can and fin hardware are surplus Nike-Ajax booster hardware. Development objectives were to provide a vehicle family with a larger diameter, larger volume payload capability than the Nike-Apache and Nike-Tomahawk vehicles at comparable cost. Both vehicles performed nominally in flight tests.

  7. Formulation, Casting, and Evaluation of Paraffin-Based Solid Fuels Containing Energetic and Novel Additives for Hybrid Rockets

    Science.gov (United States)

    Larson, Daniel B.; Desain, John D.; Boyer, Eric; Wachs, Trevor; Kuo, Kenneth K.; Borduin, Russell; Koo, Joseph H.; Brady, Brian B.; Curtiss, Thomas J.; Story, George

    2012-01-01

    This investigation studied the inclusion of various additives to paraffin wax for use in a hybrid rocket motor. Some of the paraffin-based fuels were doped with various percentages of LiAlH4 (up to 10%). Addition of LiAlH4 at 10% was found to increase regression rates between 7 - 10% over baseline paraffin through tests in a gaseous oxygen hybrid rocket motor. Mass burn rates for paraffin grains with 10% LiAlH4 were also higher than those of the baseline paraffin. RDX was also cast into a paraffin sample via a novel casting process which involved dissolving RDX into dimethylformamide (DMF) solvent and then drawing a vacuum on the mixture of paraffin and RDX/DMF in order to evaporate out the DMF. It was found that although all DMF was removed, the process was not conducive to generating small RDX particles. The slow boiling generated an inhomogeneous mixture of paraffin and RDX. It is likely that superheating the DMF to cause rapid boiling would likely reduce RDX particle sizes. In addition to paraffin/LiAlH4 grains, multi-walled carbon nanotubes (MWNT) were cast in paraffin for testing in a hybrid rocket motor, and assorted samples containing a range of MWNT percentages in paraffin were imaged using SEM. The fuel samples showed good distribution of MWNT in the paraffin matrix, but the MWNT were often agglomerated, indicating that a change to the sonication and mixing processes were required to achieve better uniformity and debundled MWNT. Fuel grains with MWNT fuel grains had slightly lower regression rate, likely due to the increased thermal conductivity to the fuel subsurface, reducing the burning surface temperature.

  8. Launch Excitement with Water Rockets

    Science.gov (United States)

    Sanchez, Juan Carlos; Penick, John

    2007-01-01

    Explosions and fires--these are what many students are waiting for in science classes. And when they do occur, students pay attention. While we can't entertain our students with continual mayhem, we can catch their attention and cater to their desires for excitement by saying, "Let's make rockets." In this activity, students make simple, reusable…

  9. Ignition and Flame Stabilization of a Strut-Jet RBCC Combustor with Small Rocket Exhaust

    Directory of Open Access Journals (Sweden)

    Jichao Hu

    2014-01-01

    Full Text Available A Rocket Based Combined Cycle combustor model is tested at a ground direct connected rig to investigate the flame holding characteristics with a small rocket exhaust using liquid kerosene. The total temperature and the Mach number of the vitiated air flow, at exit of the nozzle are 1505 K and 2.6, respectively. The rocket base is embedded in a fuel injecting strut and mounted in the center of the combustor. The wall of the combustor is flush, without any reward step or cavity, so the strut-jet is used to make sure of the flame stabilization of the second combustion. Mass flow rate of the kerosene and oxygen injected into the rocket is set to be a small value, below 10% of the total fuel when the equivalence ratio of the second combustion is 1. The experiment has generated two different kinds of rocket exhaust: fuel rich and pure oxygen. Experiment result has shown that, with a relative small total mass flow rate of the rocket, the fuel rich rocket plume is not suitable for ignition and flame stabilization, while an oxygen plume condition is suitable. Then the paper conducts a series of experiments to investigate the combustion characteristics under this oxygen pilot method and found that the flame stabilization characteristics are different at different combustion modes.

  10. High-temperature turbopump assembly for space nuclear thermal propulsion

    Science.gov (United States)

    Overholt, David M.

    1993-01-01

    The development of a practical, high-performance nuclear rocket by the U.S. Air Force Space Nuclear Thermal Propulsion (SNTP) program places high priority on maximizing specific impulse (ISP) and thrust-to-weight ratio. The operating parameters arising from these goals drive the propellant-pump design. The liquid hydrogen propellant is pressurized and pumped to the reactor inlet by the turbopump assembly (TPA). Rocket propulsion is effected by rapid heating of the propellant from 100 K to thousands of degrees in the particle-bed reactor (PBR). The exhausted propellant is then expanded through a high-temperature nozzle. One approach to achieve high performance is to use an uncooled carbon-carbon nozzle and duct turbine inlet. The high-temperature capability is obtained by using carbon-carbon throughout the TPA hot section. Carbon-carbon components in development include structural parts, turbine nozzles/stators, and turbine rotors. The technology spinoff is applicable to conventional liquid propulsion engines plus a wide variety of other turbomachinery applications.

  11. Carbon-carbon turbopump concept for Space Nuclear Thermal Propulsion

    Science.gov (United States)

    Overholt, David M.

    1993-06-01

    The U.S. Air Force Space Nuclear Thermal Propulsion (SNTP) program is placing high priority on maximizing specific impulse (ISP) and thrust-to-weight ratio in the development of a practical high-performance nuclear rocket. The turbopump design is driven by these goals. The liquid hydrogen propellant is pressurized and pumped to the reactor inlet by the turbopump assembly (TPA). Rocket propulsion is from rapid heating of the propellant from 180 R to thousands of degrees in the particle bed reactor (PBR). The exhausted propellant is then expanded through a high-temperature nozzle. A high-performance approach is to use an uncooled carbon-carbon nozzle and duct turbine inlet. Carbon-carbon components are used throughout the TPA hot section to obtain the high-temperature capability. Several carbon-carbon components are in development including structural parts, turbine nozzles/stators, and turbine rotors. The technology spinoff is applicable to conventional liquid propulsion engines and many other turbomachinery applications.

  12. High-temperature turbopump assembly for space nuclear thermal propulsion

    International Nuclear Information System (INIS)

    Overholt, D.M.

    1993-01-01

    The development of a practical, high-performance nuclear rocket by the U.S. Air Force Space Nuclear Thermal Propulsion (SNTP) program places high priority on maximizing specific impulse (ISP) and thrust-to-weight ratio. The operating parameters arising from these goals drive the propellant-pump design. The liquid hydrogen propellant is pressurized and pumped to the reactor inlet by the turbopump assembly (TPA). Rocket propulsion is effected by rapid heating of the propellant from 100 K to thousands of degrees in the particle-bed reactor (PBR). The exhausted propellant is then expanded through a high-temperature nozzle. One approach to achieve high performance is to use an uncooled carbon-carbon nozzle and duct turbine inlet. The high-temperature capability is obtained by using carbon-carbon throughout the TPA hot section. Carbon-carbon components in development include structural parts, turbine nozzles/stators, and turbine rotors. The technology spinoff is applicable to conventional liquid propulsion engines plus a wide variety of other turbomachinery applications

  13. Carbon-carbon turbopump concept for Space Nuclear Thermal Propulsion

    International Nuclear Information System (INIS)

    Overholt, D.M.

    1993-06-01

    The U.S. Air Force Space Nuclear Thermal Propulsion (SNTP) program is placing high priority on maximizing specific impulse (ISP) and thrust-to-weight ratio in the development of a practical high-performance nuclear rocket. The turbopump design is driven by these goals. The liquid hydrogen propellant is pressurized and pumped to the reactor inlet by the turbopump assembly (TPA). Rocket propulsion is from rapid heating of the propellant from 180 R to thousands of degrees in the particle bed reactor (PBR). The exhausted propellant is then expanded through a high-temperature nozzle. A high-performance approach is to use an uncooled carbon-carbon nozzle and duct turbine inlet. Carbon-carbon components are used throughout the TPA hot section to obtain the high-temperature capability. Several carbon-carbon components are in development including structural parts, turbine nozzles/stators, and turbine rotors. The technology spinoff is applicable to conventional liquid propulsion engines and many other turbomachinery applications. 3 refs

  14. 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

  15. Optimal Reference Strain Structure for Studying Dynamic Responses of Flexible Rockets

    Science.gov (United States)

    Tsushima, Natsuki; Su, Weihua; Wolf, Michael G.; Griffin, Edwin D.; Dumoulin, Marie P.

    2017-01-01

    In the proposed paper, the optimal design of reference strain structures (RSS) will be performed targeting for the accurate observation of the dynamic bending and torsion deformation of a flexible rocket. It will provide the detailed description of the finite-element (FE) model of a notional flexible rocket created in MSC.Patran. The RSS will be attached longitudinally along the side of the rocket and to track the deformation of the thin-walled structure under external loads. An integrated surrogate-based multi-objective optimization approach will be developed to find the optimal design of the RSS using the FE model. The Kriging method will be used to construct the surrogate model. For the data sampling and the performance evaluation, static/transient analyses will be performed with MSC.Natran/Patran. The multi-objective optimization will be solved with NSGA-II to minimize the difference between the strains of the launch vehicle and RSS. Finally, the performance of the optimal RSS will be evaluated by checking its strain-tracking capability in different numerical simulations of the flexible rocket.

  16. Nuclear vapor thermal reactor propulsion technology

    International Nuclear Information System (INIS)

    Maya, I.; Diaz, N.J.; Dugan, E.T.; Watanabe, Y.; McClanahan, J.A.; Wen-Hsiung Tu; Carman, R.L.

    1993-01-01

    The conceptual design of a nuclear rocket based on the vapor core reactor is presented. The Nuclear Vapor Thermal Rocket (NVTR) offers the potential for a specific impulse of 1000 to 1200 s at thrust-to-weight ratios of 1 to 2. The design is based on NERVA geometry and systems with the solid fuel replaced by uranium tetrafluoride (UF 4 ) vapor. The closed-loop core does not rely on hydrodynamic confinement of the fuel. The hydrogen propellant is separated from the UF 4 fuel gas by graphite structure. The hydrogen is maintained at high pressure (∼100 atm), and exits the core at 3,100 K to 3,500 K. Zirconium carbide and hafnium carbide coatings are used to protect the hot graphite from the hydrogen. The core is surrounded by beryllium oxide reflector. The nuclear reactor core has been integrated into a 75 klb engine design using an expander cycle and dual turbopumps. The NVTR offers the potential for an incremental technology development pathway to high performance gas core reactors. Since the fuel is readily available, it also offers advantages in the initial cost of development, as it will not require major expenditures for fuel development

  17. Space Power Experiments Aboard Rockets SPEAR-3

    National Research Council Canada - National Science Library

    Raitt, W. J

    1997-01-01

    The SPEAR-3 program was a sounding rocket payload designed to study the interaction of a charged body with the Earth's upper atmosphere with particular reference to the discharging ability of selected...

  18. Nuclear rocket propulsion

    International Nuclear Information System (INIS)

    Clark, J.S.; Miller, T.J.

    1991-01-01

    NASA has initiated planning for a technology development project for nuclear rocket propulsion systems for Space Exploration Initiative (SEI) human and robotic missions to the Moon and to Mars. An Interagency project is underway that includes the Department of Energy National Laboratories for nuclear technology development. This paper summarizes the activities of the project planning team in FY 1990 and FY 1991, discusses the progress to date, and reviews the project plan. Critical technology issues have been identified and include: nuclear fuel temperature, life, and reliability; nuclear system ground test; safety; autonomous system operation and health monitoring; minimum mass and high specific impulse

  19. 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.

  20. 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.

  1. Stochastic rocket dynamics under random nozzle side loads: Ornstein-Uhlenbeck boundary layer separation and its coarse grained connection to side loading and rocket response

    Energy Technology Data Exchange (ETDEWEB)

    Keanini, R.G.; Srivastava, N.; Tkacik, P.T. [Department of Mechanical Engineering, University of North Carolina at Charlotte, 9201 University City Blvd., Charlotte, NC 28078 (United States); Weggel, D.C. [Department of Civil and Environmental Engineering, University of North Carolina at Charlotte, 9201 University City Blvd., Charlotte, NC 28078 (United States); Knight, P.D. [Mitchell Aerospace and Engineering, Statesville, North Carolina 28677 (United States)

    2011-06-15

    A long-standing, though ill-understood problem in rocket dynamics, rocket response to random, altitude-dependent nozzle side-loads, is investigated. Side loads arise during low altitude flight due to random, asymmetric, shock-induced separation of in-nozzle boundary layers. In this paper, stochastic evolution of the in-nozzle boundary layer separation line, an essential feature underlying side load generation, is connected to random, altitude-dependent rotational and translational rocket response via a set of simple analytical models. Separation line motion, extant on a fast boundary layer time scale, is modeled as an Ornstein-Uhlenbeck process. Pitch and yaw responses, taking place on a long, rocket dynamics time scale, are shown to likewise evolve as OU processes. Stochastic, altitude-dependent rocket translational motion follows from linear, asymptotic versions of the full nonlinear equations of motion; the model is valid in the practical limit where random pitch, yaw, and roll rates all remain small. Computed altitude-dependent rotational and translational velocity and displacement statistics are compared against those obtained using recently reported high fidelity simulations [Srivastava, Tkacik, and Keanini, J. Appl. Phys. 108, 044911 (2010)]; in every case, reasonable agreement is observed. As an important prelude, evidence indicating the physical consistency of the model introduced in the above article is first presented: it is shown that the study's separation line model allows direct derivation of experimentally observed side load amplitude and direction densities. Finally, it is found that the analytical models proposed in this paper allow straightforward identification of practical approaches for: (i) reducing pitch/yaw response to side loads, and (ii) enhancing pitch/yaw damping once side loads cease. (Copyright copyright 2011 WILEY-VCH Verlag GmbH and Co. KGaA, Weinheim)

  2. An echelle spectrograph for middle ultraviolet solar spectroscopy from rockets.

    Science.gov (United States)

    Tousey, R; Purcell, J D; Garrett, D L

    1967-03-01

    An echelle grating spectrograph is ideal for use in a rocket when high resolution is required becaus itoccupies a minimum of space. The instrument described covers the range 4000-2000 A with a resolution of 0.03 A. It was designed to fit into the solar biaxial pointing-control section of an Aerobee-150 rocket. The characteristics of the spectrograph are illustrated with laboratory spectra of iron and carbon are sources and with solar spectra obtained during rocket flights in 1961 and 1964. Problems encountered in analyzing the spectra are discussed. The most difficult design problem was the elimination of stray light when used with the sun. Of the several methods investigated, the most effective was a predispersing system in the form of a zero-dispersion double monochromator. This was made compact by folding the beam four times.

  3. A six degree-of-freedom thrust sensor for a labscale hybrid rocket

    International Nuclear Information System (INIS)

    Wright, Ann M; Born, Traig; Strickland, Ryan; Wright, Andrew B

    2013-01-01

    A six degree-of-freedom thrust sensor was designed, constructed, calibrated, and tested using the labscale hybrid rocket at the University of Arkansas at Little Rock. The system consisted of six independent legs: one parallel to the axis of symmetry of the rocket for main thrust measurement, two vertical legs near the nozzle end of the rocket, one vertical leg near the oxygen input end of the rocket, and two separated horizontal legs near the nozzle end. Each leg was composed of a rotational bearing, a load cell, and a universal joint above and below the load cell. The leg was designed to create point contact along only one direction and minimize the non-axial forces applied to the load cell. With this system, force in each direction and moments for roll, pitch, and yaw can be measured. The system was calibrated and tested using a labscale hybrid rocket using gaseous oxygen and hydroxyl-terminated polybutadiene solid fuel. The thrust stand proved to be stable during calibration tests. Thrust force vector components and roll, pitch, and yaw moments were calculated for test firings with an oxygen mass flow rate range of 0.0174–0.0348 kg s −1 . (paper)

  4. Rocket measurements of positive ions during polar mesosphere winter echo conditions

    Directory of Open Access Journals (Sweden)

    A. Brattli

    2006-01-01

    Full Text Available On 18 January 2005, two small, instrumented rockets were launched from Andøya Rocket Range (69.3° N, 16° E during conditions with Polar Mesosphere Winter Echoes (PMWE. Each of the rockets was equipped with a Positive Ion Probe (PIP and a Faraday rotation/differential absorption experiment, and was launched as part of a salvo of meteorological rockets measuring temperature and wind using falling spheres and chaff. Layers of PMWE were detected between 55 and 77 km by the 53.5 MHz ALWIN radar. The rockets were launched during a solar proton event, and measured extremely high ion densities, of order 1010 m−3, in the region where PMWE were observed. The density measurements were analyzed with the wavelet transform technique. At large length scales, ~103 m, the power spectral density can be fitted with a k−3 wave number dependence, consistent with saturated gravity waves. Outside the PMWE layers the k−3 spectrum extends down to approximately 102 m where the fluctuations are quickly damped and disappear into the instrumental noise. Inside the PMWE layers the spectrum at smaller length scales is well fitted with a k−5/3 dependence over two decades of scales. The PMWE are therefore clearly indicative of turbulence, and the data are consistent with the turbulent dissipation of breaking gravity waves. We estimate a lower limit for the turbulent energy dissipation rate of about 10−2 W/kg in the upper (72 km layer.

  5. Cambridge Rocketry Simulator – A Stochastic Six-Degrees-of-Freedom Rocket Flight Simulator

    OpenAIRE

    Eerland, Willem J.; Box, Simon; Sóbester, András

    2017-01-01

    The Cambridge Rocketry Simulator can be used to simulate the flight of unguided rockets for both design and operational applications. The software consists of three parts: The first part is a GUI that enables the user to design a rocket. The second part is a verified and peer-reviewed physics model that simulates the rocket flight. This includes a Monte Carlo wrapper to model the uncertainty in the rocket’s dynamics and the atmospheric conditions. The third part generates visualizations of th...

  6. Radiophysical and geomagnetic effects of rocket burn and launch in the near-the-earth environment

    CERN Document Server

    Chernogor, Leonid F

    2013-01-01

    Radiophysical and Geomagnetic Effects of Rocket Burn and Launch in the Near-the-Earth Environment describes experimental and theoretical studies on the effects of rocket burns and launchings on the near-the-Earth environment and geomagnetic fields. It illuminates the main geophysical and radiophysical effects on the ionosphere and magnetosphere surrounding the Earth that accompany rocket or cosmic apparatus burns and launchings from 1,000 to 10,000 kilometers.The book analyzes the disturbances of plasma and the ambient magnetic and electric fields in the near-Earth environment from rocket burn

  7. Measurement of the electrostatic field in aurora by antarctic rocket

    International Nuclear Information System (INIS)

    Takeya, Yoshio; Minami, Shigeyuki

    1974-01-01

    The direct measurement of the electrostatic field produced by the flow of charged particles and geomagnetic field in aurora has been carried out by means of rockets or satellites. The construction of an electric field meter and its characteristics are described, which measures the vectors of electric field with antarctic rockets. New scheme is presented: three components of an electric field are directly obtained through the probes set in three directions. (Mori, K.)

  8. THE STERN PROJECT–HANDS ON ROCKETS SCIENCE FOR UNIVERSITY STUDENT

    OpenAIRE

    Schüttauf, Katharina; Stamminger, Andreas; Lappöhn, Karsten

    2017-01-01

    In April 2012, the German Aerospace Center DLR initiated a sponsorship program for university students to develop, build and launch their own rockets over a period of three years. The program designation STERN was abbreviated from the German “STudentische Experimental-RaketeN”, which translates to Student- Experimental-Rockets. The primary goal of the STERN program is to inspire students in the subject of space transportation through hands-on activities within a pro...

  9. Microcomputers, Model Rockets, and Race Cars.

    Science.gov (United States)

    Mirus, Edward A., Jr.

    1985-01-01

    The industrial education orientation program at Wisconsin School for the Deaf (WSD) presents problem-solving situations to all seventh- and eighth-grade hearing-impaired students. WSD developed user-friendly microcomputer software to guide students individually through complex computations involving model race cars and rockets while freeing…

  10. Space fireworks for upper atmospheric wind measurements by sounding rocket experiments

    Science.gov (United States)

    Yamamoto, M.

    2016-01-01

    Artificial meteor trains generated by chemical releases by using sounding rockets flown in upper atmosphere were successfully observed by multiple sites on ground and from an aircraft. We have started the rocket experiment campaign since 2007 and call it "Space fireworks" as it illuminates resonance scattering light from the released gas under sunlit/moonlit condition. By using this method, we have acquired a new technique to derive upper atmospheric wind profiles in twilight condition as well as in moonlit night and even in daytime. Magnificent artificial meteor train images with the surrounding physics and dynamics in the upper atmosphere where the meteors usually appear will be introduced by using fruitful results by the "Space firework" sounding rocket experiments in this decade.

  11. Nitrates and Glucosinolates as Strong Determinants of the Nutritional Quality in Rocket Leafy Salads

    Directory of Open Access Journals (Sweden)

    Marina Cavaiuolo

    2014-04-01

    Full Text Available Rocket is an important leafy vegetable crop and a good source of antioxidants and anticancer molecules such as glucosinolates and other sulfur compounds. Rocket is also a hyper-accumulator of nitrates which have been considered for long time the main factors that cause gastro-intestinal cancer. In this review, the content of these compounds in rocket tissues and their levels at harvest and during storage are discussed. Moreover, the effect of these compounds in preventing or inducing human diseases is also highlighted. This review provides an update to all the most recent studies carried out on rocket encouraging the consumption of this leafy vegetable to reduce the risk of contracting cancer and other cardiovascular diseases.

  12. Nitrates and glucosinolates as strong determinants of the nutritional quality in rocket leafy salads.

    Science.gov (United States)

    Cavaiuolo, Marina; Ferrante, Antonio

    2014-04-14

    Rocket is an important leafy vegetable crop and a good source of antioxidants and anticancer molecules such as glucosinolates and other sulfur compounds. Rocket is also a hyper-accumulator of nitrates which have been considered for long time the main factors that cause gastro-intestinal cancer. In this review, the content of these compounds in rocket tissues and their levels at harvest and during storage are discussed. Moreover, the effect of these compounds in preventing or inducing human diseases is also highlighted. This review provides an update to all the most recent studies carried out on rocket encouraging the consumption of this leafy vegetable to reduce the risk of contracting cancer and other cardiovascular diseases.

  13. Using PDV to Understand Damage in Rocket Motor Propellants

    Science.gov (United States)

    Tear, Gareth; Chapman, David; Ottley, Phillip; Proud, William; Gould, Peter; Cullis, Ian

    2017-06-01

    There is a continuing requirement to design and manufacture insensitive munition (IM) rocket motors for in-service use under a wide range of conditions, particularly due to shock initiation and detonation of damaged propellant spalled across the central bore of the rocket motor (XDT). High speed photography has been crucial in determining this behaviour, however attempts to model the dynamic behaviour are limited by the lack of precision particle and wave velocity data with which to validate against. In this work Photonic Doppler Velocimetery (PDV) has been combined with high speed video to give accurate point velocity and timing measurements of the rear surface of a propellant block impacted by a fragment travelling upto 1.4 km s-1. By combining traditional high speed video with PDV through a dichroic mirror, the point of velocity measurement within the debris cloud has been determined. This demonstrates a new capability to characterise the damage behaviour of a double base rocket motor propellant and hence validate the damage and fragmentation algorithms used in the numerical simulations.

  14. A research on polyether glycol replaced APCP rocket propellant

    Science.gov (United States)

    Lou, Tianyou; Bao, Chun Jia; Wang, Yiyang

    2017-08-01

    Ammonium perchlorate composite propellant (APCP) is a modern solid rocket propellant used in rocket vehicles. It differs from many traditional solid rocket propellants by the nature of how it is processed. APCP is cast into shape, as opposed to powder pressing it with black powder. This provides manufacturing regularity and repeatability, which are necessary requirements for use in the aerospace industry. For traditional APCP, ingredients normally used are ammonium peroxide, aluminum, Hydroxyl-terminated polybutadiene(HTPB), curing agency and other additives, the greatest disadvantage is that the fuel is too expensive. According to the price we collected in our country, a single kilogram of this fuel will cost 200 Yuan, which is about 35 dollars, for a fan who may use tons of the fuel in a single year, it definitely is a great deal of money. For this reason, we invented a new kind of APCP fuel. Changing adhesive agency from cross-linked htpb to cross linked polyether glycol gives a similar specific thrust, density and mechanical property while costs a lower price.

  15. Nuclear propulsion tradeoffs for manned Mars missions

    International Nuclear Information System (INIS)

    Walton, L.A.; Malloy, J.D.

    1991-01-01

    A conjunction class split/sprint manned Mars exploration mission was studied to evaluate tradeoffs in performance characteristics of nuclear thermal rockets. A Particle Bed Reactor-based nuclear thermal rocket was found to offer a 38% to 52% total mass savings compared with a NERVA-based nuclear thermal rocket for this mission. This advantage is primarily due to the higher thrust-to-weight ratio of the Particle Bed Reactor nuclear rocket. The mission is enabled by nuclear thermal rockets. It cannot be performed practically using chemical propulsion

  16. Unsupervised Anomaly Detection for Liquid-Fueled Rocket Prop...

    Data.gov (United States)

    National Aeronautics and Space Administration — Title: Unsupervised Anomaly Detection for Liquid-Fueled Rocket Propulsion Health Monitoring. Abstract: This article describes the results of applying four...

  17. Hezbollah Rockets and Missiles: Deployment and Defence Programme

    Directory of Open Access Journals (Sweden)

    Ladislav Kulhánek

    2018-01-01

    Full Text Available The paper addresses the issue of the proliferation of the rocket technology deployed in guerrilla warfare throughout the Near East. It evaluates assumptions and identifies potential regressions with respect to the most recent hot phase in the violent conflict between Hezbollah and Israel, with a focus on primary methods based on rocket shelling and resistance efforts. In comparing the two military forces, Israel emerges as the stronger, even if artillery and tactical ballistic missiles with warheads containing substances defined as weapons of mass destruction were to be deployed as an extreme first measure by Hezbollah, in which case an Israeli or American response would destroy at least parts of Lebanon.

  18. The Space Nuclear Thermal Propulsion Program: Propulsion for the twenty first century

    International Nuclear Information System (INIS)

    Bleeker, G.; Moody, J.; Kesaree, M.

    1993-01-01

    As mission requirements approach the limits of the chemical propulsion systems, new engines must be investigated that can meet the advanced mission requirements of higher payload fractions, higher velocities, and consequently higher specific Impulses (Isp). The propulsion system that can meet these high demands is a nuclear thermal rocket engine. This engine generates the thrust by expanding/existing the hydrogen, heated from the energy derived from the fission process in a reactor, through a nozzle. The Department of Defense (DoD), however, initiated a new nuclear rocket development program in 1987 for ballistic missile defense application. The Space Nuclear Thermal Propulsion (SNTP) Program that seeks to improve on the technology of ROVER/NERVA grew out of this beginning and has been managed by the Air Force, with the involvement of DoE and NASA. The goal of the SNTP Program is to develop an engine to meet potential Air Force requirements for upper stage engine, bimodal propulsion/power applications, and orbital transfer vehicles, as well as the NASA requirements for possible missions to the Moon and Mars. During the entire life of the program, the DoD has considered safety to be of paramount importance, and is following all national environmental policies

  19. Seismic tests at the HDR facility using explosives and solid propellant rockets

    International Nuclear Information System (INIS)

    Corvin, P.; Steinhilber, H.

    1981-01-01

    In blast tests the HDR reactor building and its mechanical equipment were subjected to earthquake-type excitations through the soil and the foundation. A series of six tests was carried out, two tests being made with HDR facility under operating conditions (BWR conditions, 285 0 C, 70 bar). The charges were placed in boreholes at a depth of 4 to 10 m and a distance of 16 to 25 m from the reactor building. The tests with solid propellant rockets were made in order to try a new excitation technique. The rockets used in these tests were of compact design and had a short combustion period (500 ms) at high constant thrust (100 kN per combustion chamber). These rockets were fixed to the concrete dome of the building in such a way that the thrust generated during the combustion of the propellant resulted in an impulsive load acting on the building. This type of excitation was selected with a view to investigating the global effects of the load case 'aircraft impact' on the building and the mechanical equipment. In the four tests made so far, up to four rockets were ignited simultaneously, so that the maximum impulse was 2 x 10 5 Ns. The excitation level can be markedly increased by adding further rockets. This excitation technique was characterised by excellent reproducibility of the load parameters. (orig./HP)

  20. Sounding rocket study of auroral electron precipitation

    International Nuclear Information System (INIS)

    McFadden, J.P.

    1985-01-01

    Measurement of energetic electrons in the auroral zone have proved to be one of the most useful tools in investigating the phenomena of auroral arc formation. This dissertation presents a detailed analysis of the electron data from two sounding rocket campaigns and interprets the measurements in terms of existing auroral models. The Polar Cusp campaign consisted of a single rocket launched from Cape Parry, Canada into the afternoon auroral zone at 1:31:13 UT on January 21, 1982. The results include the measurement of a narrow, magnetic field aligned electron flux at the edge of an arc. This electron precipitation was found to have a remarkably constant 1.2 eV temperature perpendicular to the magnetic field over a 200 to 900 eV energy range. The payload also made simultaneous measurements of both energetic electrons and 3-MHz plasma waves in an auroral arc. Analysis has shown that the waves are propagating in the upper hybrid band and should be generated by a positive slope in the parallel electron distribution. A correlation was found between the 3-MHz waves and small positive slopes in the parallel electron distribution but experimental uncertainties in the electron measurement were large enough to influence the analysis. The BIDARCA campaign consisted of two sounding rockets launched from Poker Flat and Fort Yukon, Alaska at 9:09:00 UT and 9:10:40 UT on February 7, 1984

  1. X-ray scanning of overhead aurorae from rockets

    International Nuclear Information System (INIS)

    Barcus, J.R.; Goldberg, R.A.

    1981-01-01

    Two Nike Tomahawk rocket payloads were launched into energetic auroral events to investigate their structure and effects on the atmosphere. The instrument complement included X-ray scintillation detectors with energy discrimination in four ranges to measure the deposition of bremsstrahlung produced X-rays within the stratosphere and mesosphere. For this purpose, each instrument was designed for wide angle viewing; however, properties of the rocket motion have permitted coarse observation of distinct spatial X-ray structure. The detectors were mounted at a 45 0 angle with respect to the payload axis to permit scanning of the upper hemisphere, with rocket spin rates near 5 c/s during the upleg portion of each flight. Here, atmospheric shielding reduced energetic particle contamination effects to insignificant values below 65 to 75 km. Iterative computer techniques were used to reconstruct X-ray source maps at 100 km, taking atmospheric absorption effects into account. Payload 18.178 was launched on 21 September (0302 LMT) into an aurora observed to have two distinct azimuthal regions of optical brightness. Payload 18.179 (23 September, 0147 LMT) was launched into an aurora of more diffuse character. The presence of a two component spectrum is indicated for each event with the hard component originating in the more diffuse, optically faint regions. (author)

  2. Ultra-fast Escape of a Octopus-inspired Rocket

    Science.gov (United States)

    Weymouth, Gabriel; Triantafyllou, Michael

    2013-11-01

    The octopus, squid, and other cephalopods inflate with water and then release a jet to accelerate in the opposite direction. This escape mechanism is particularly interesting in the octopus because they become initially quite bluff, yet this does not hinder them in achieving impressive bursts of speed. We examine this somewhat paradoxical maneuver using a simple deflating spheroid model in both potential and viscous flow. We demonstrate that the dynamic reduction of the width of the body completely changes the flow and forces acting on the escaping rocket in three ways. First, a body which reduces in size can generate an added mass thrust which counteracts the added mass inertia. Second, the motion of the shrinking wall acts similar to suction on a static wall, reducing separation and drag forces in a viscous fluid, but that this effects depends on the rate of size change. Third, using a combination of these two features it is possible to initially load the fluid with kinetic energy when heavy and bluff and then recover that energy when streamlined and light, enabling ultra-fast accelerations. As a notable example, these mechanisms allow a shrinking spheroid rocket in a heavy inviscid fluid to achieve speeds greater than an identical rocket in the vacuum of space. Southampton Marine and Maritime Institute.

  3. 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

  4. Studies of small-scale plasma inhomogeneities in the cusp ionosphere using sounding rocket data

    Science.gov (United States)

    Chernyshov, Alexander A.; Spicher, Andres; Ilyasov, Askar A.; Miloch, Wojciech J.; Clausen, Lasse B. N.; Saito, Yoshifumi; Jin, Yaqi; Moen, Jøran I.

    2018-04-01

    Microprocesses associated with plasma inhomogeneities are studied on the basis of data from the Investigation of Cusp Irregularities (ICI-3) sounding rocket. The ICI-3 rocket is devoted to investigating a reverse flow event in the cusp F region ionosphere. By numerical stability analysis, it is demonstrated that inhomogeneous-energy-density-driven (IEDD) instability can be a mechanism for the excitation of small-scale plasma inhomogeneities. The Local Intermittency Measure (LIM) method also applied the rocket data to analyze irregular structures of the electric field during rocket flight in the cusp. A qualitative agreement between high values of the growth rates of the IEDD instability and the regions with enhanced LIM is observed. This suggests that IEDD instability is connected to turbulent non-Gaussian processes.

  5. 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

  6. Simulation and experimental research on line throwing rocket with flight

    OpenAIRE

    Wen-bin Gu; Ming Lu; Jian-qing Liu; Qin-xing Dong; Zhen-xiong Wang; Jiang-hai Chen

    2014-01-01

    The finite segment method is used to model the line throwing rocket system. A dynamic model of line throwing rocket with flight motion based on Kane's method is presented by the kinematics description of the system and the consideration of the forces acting on the system. The experiment designed according to the parameters of the dynamic model is made. The simulation and experiment results, such as range, velocity and flight time, are compared and analyzed. The simulation results are basicall...

  7. Nitrates and Glucosinolates as Strong Determinants of the Nutritional Quality in Rocket Leafy Salads

    OpenAIRE

    Cavaiuolo, Marina; Ferrante, Antonio

    2014-01-01

    Rocket is an important leafy vegetable crop and a good source of antioxidants and anticancer molecules such as glucosinolates and other sulfur compounds. Rocket is also a hyper-accumulator of nitrates which have been considered for long time the main factors that cause gastro-intestinal cancer. In this review, the content of these compounds in rocket tissues and their levels at harvest and during storage are discussed. Moreover, the effect of these compounds in preventing or inducing human di...

  8. Pocket rocket: An electrothermal plasma micro-thruster

    Science.gov (United States)

    Greig, Amelia Diane

    Recently, an increase in use of micro-satellites constructed from commercial off the shelf (COTS) components has developed, to address the large costs associated with designing, testing and launching satellites. One particular type of micro-satellite of interest are CubeSats, which are modular 10 cm cubic satellites with total weight less than 1.33 kg. To assist with orbit boosting and attitude control of CubeSats, micro-propulsion systems are required, but are currently limited. A potential electrothermal plasma micro-thruster for use with CubeSats or other micro-satellites is under development at The Australian National University and forms the basis for this work. The thruster, known as ‘Pocket Rocket’, utilises neutral gas heating from ion-neutral collisions within a weakly ionised asymmetric plasma discharge, increasing the exhaust thermal velocity of the propellant gas, thereby producing higher thrust than if the propellant was emitted cold. In this work, neutral gas temperature of the Pocket Rocket discharge is studied in depth using rovibrational spectroscopy of the nitrogen (N2) second positive system (C3Πu → B3Πg), using both pure N2 and argon/N2 mixtures as the operating gas. Volume averaged steady state gas temperatures are measured for a range of operating conditions, with an analytical collisional model developed to verify experimental results. Results show that neutral gas heating is occurring with volume averaged steady state temperatures reaching 430 K in N2 and 1060 K for argon with 1% N2 at standard operating conditions of 1.5 Torr pressure and 10 W power input, demonstrating proof of concept for the Pocket Rocket thruster. Spatiotemporal profiles of gas temperature identify that the dominant heating mechanisms are ion-neutral collisions within the discharge and wall heating from ion bombardment of the thruster walls. To complement the experimental results, computational fluid dynamics (CFD) simulations using the commercial CFD

  9. Government Relations: It's Not Rocket Science

    Science.gov (United States)

    Radway, Mike

    2007-01-01

    Many people in the early childhood education field are afraid of government relations work, intimidated by politicians, and believe the whole process is unseemly. The author asserts that they should not be afraid nor be intimidated because government relations is not rocket science and fundamentally officeholders are no different from the rest of…

  10. Description and Flight Performance Results of the WASP Sounding Rocket

    Science.gov (United States)

    De Pauw, J. F.; Steffens, L. E.; Yuska, J. A.

    1968-01-01

    A general description of the design and construction of the WASP sounding rocket and of the performance of its first flight are presented. The purpose of the flight test was to place the 862-pound (391-kg) spacecraft above 250 000 feet (76.25 km) on free-fall trajectory for at least 6 minutes in order to study the effect of "weightlessness" on a slosh dynamics experiment. The WASP sounding rocket fulfilled its intended mission requirements. The sounding rocket approximately followed a nominal trajectory. The payload was in free fall above 250 000 feet (76.25 km) for 6.5 minutes and reached an apogee altitude of 134 nautical miles (248 km). Flight data including velocity, altitude, acceleration, roll rate, and angle of attack are discussed and compared to nominal performance calculations. The effect of residual burning of the second stage motor is analyzed. The flight vibration environment is presented and analyzed, including root mean square (RMS) and power spectral density analysis.

  11. 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...

  12. 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.

  13. Detection of aurorae in light time of the day at rocket investigations of atmospheric radiation

    International Nuclear Information System (INIS)

    Khokhlov, V.N.

    1996-01-01

    Results of rocket observations of aurorae in light time of the day were analyzed. Characteristic features of Rayleigh scattering, day airglow, solar radiation, scattered in the device and near-rocket glow were considered. The contribution of aurorae in the light time of the day was determined on the basis of analyzing results of rocket experiments, laboratory measurements and theoretical simulation. 4 refs., 2 figs

  14. 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.

  15. Thermal characterizations of the paraffin wax/low density polyethylene blends as a solid fuel

    Energy Technology Data Exchange (ETDEWEB)

    Kim, Soojong; Moon, Heejang; Kim, Jinkon, E-mail: jkkim@kau.ac.kr

    2015-08-10

    Highlights: • Regression rate of blends fuel is higher than polymer fuel. • LDPE is an effective mixing ingredient for the combustion efficiency. • Blends fuel is a uniform mixture with two degradation steps. • LDPE plays a positive role for the low sensitivity to the thermal deformation • Blends with low LDPE content can be an effective fuel for hybrid rocket application. - Abstract: Thermal characterizations of a novel solid fuel for hybrid rocket application, based on the paraffin wax blends with low density polyethylene (LDPE) concentration of 5% (SF-5) and 10% (SF-10) were conducted. Both the increased regression rate in comparison with the polymeric fuel, and the improved combustion efficiency in comparison with the pure paraffin fuel reveal that the blend fuels achieve higher combustion performance. The morphology of the shape stabilized paraffin wax/LDPE blends was characterized by the scanning electron microscopy (SEM). Although the SEM observation indicated the blends have uniform mixtures, they showed two degradation steps confirming the immiscibility of components in the crystalline phase from thermogravimetric analysis (TGA). The differential scanning calorimeter (DSC) results showed that the melting temperature of LDPE in the blends decreased with an increase of paraffin wax content. The decreasing total specific melting enthalpy of blended fuels with decreasing paraffin wax content is in fairly good agreement with the additive rule. In thermomechanical analysis (TMA), the linear coefficient of thermal expansion (LCTE) seems to decrease with an increase of LDPE loading, however, the loaded LDPE do merely affect the LCTE in case of the blends with low LDPE concentration. It was found that a blend of low concentration of LDPE with a relatively high concentration of paraffin wax can lead to a potential novel fuel for rocket application, a contrary case with respect to the field of phase change materials (PCM) where a blend of high concentration

  16. Thermal characterizations of the paraffin wax/low density polyethylene blends as a solid fuel

    International Nuclear Information System (INIS)

    Kim, Soojong; Moon, Heejang; Kim, Jinkon

    2015-01-01

    Highlights: • Regression rate of blends fuel is higher than polymer fuel. • LDPE is an effective mixing ingredient for the combustion efficiency. • Blends fuel is a uniform mixture with two degradation steps. • LDPE plays a positive role for the low sensitivity to the thermal deformation • Blends with low LDPE content can be an effective fuel for hybrid rocket application. - Abstract: Thermal characterizations of a novel solid fuel for hybrid rocket application, based on the paraffin wax blends with low density polyethylene (LDPE) concentration of 5% (SF-5) and 10% (SF-10) were conducted. Both the increased regression rate in comparison with the polymeric fuel, and the improved combustion efficiency in comparison with the pure paraffin fuel reveal that the blend fuels achieve higher combustion performance. The morphology of the shape stabilized paraffin wax/LDPE blends was characterized by the scanning electron microscopy (SEM). Although the SEM observation indicated the blends have uniform mixtures, they showed two degradation steps confirming the immiscibility of components in the crystalline phase from thermogravimetric analysis (TGA). The differential scanning calorimeter (DSC) results showed that the melting temperature of LDPE in the blends decreased with an increase of paraffin wax content. The decreasing total specific melting enthalpy of blended fuels with decreasing paraffin wax content is in fairly good agreement with the additive rule. In thermomechanical analysis (TMA), the linear coefficient of thermal expansion (LCTE) seems to decrease with an increase of LDPE loading, however, the loaded LDPE do merely affect the LCTE in case of the blends with low LDPE concentration. It was found that a blend of low concentration of LDPE with a relatively high concentration of paraffin wax can lead to a potential novel fuel for rocket application, a contrary case with respect to the field of phase change materials (PCM) where a blend of high concentration

  17. Manufacturing Advanced Channel Wall Rocket Liners, Phase I

    Data.gov (United States)

    National Aeronautics and Space Administration — This SBIR will adapt and demonstrate a low cost flexible method of manufacturing channel wall liquid rocket nozzles and combustors, while providing developers a...

  18. 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...

  19. Small Reactor Designs Suitable for Direct Nuclear Thermal Propulsion: Interim Report

    International Nuclear Information System (INIS)

    Schnitzler, Bruce G.

    2012-01-01

    Advancement of U.S. scientific, security, and economic interests requires high performance propulsion systems to support missions beyond low Earth orbit. A robust space exploration program will include robotic outer planet and crewed missions to a variety of destinations including the moon, near Earth objects, and eventually Mars. Past studies, in particular those in support of both the Strategic Defense Initiative (SDI) and the Space Exploration Initiative (SEI), have shown nuclear thermal propulsion systems provide superior performance for high mass high propulsive delta-V missions. In NASA's recent Mars Design Reference Architecture (DRA) 5.0 study, nuclear thermal propulsion (NTP) was again selected over chemical propulsion as the preferred in-space transportation system option for the human exploration of Mars because of its high thrust and high specific impulse (∼900 s) capability, increased tolerance to payload mass growth and architecture changes, and lower total initial mass in low Earth orbit. The recently announced national space policy2 supports the development and use of space nuclear power systems where such systems safely enable or significantly enhance space exploration or operational capabilities. An extensive nuclear thermal rocket technology development effort was conducted under the Rover/NERVA, GE-710 and ANL nuclear rocket programs (1955-1973). Both graphite and refractory metal alloy fuel types were pursued. The primary and significantly larger Rover/NERVA program focused on graphite type fuels. Research, development, and testing of high temperature graphite fuels was conducted. Reactors and engines employing these fuels were designed, built, and ground tested. The GE-710 and ANL programs focused on an alternative ceramic-metallic 'cermet' fuel type consisting of UO2 (or UN) fuel embedded in a refractory metal matrix such as tungsten. The General Electric program examined closed loop concepts for space or terrestrial applications as well as

  20. The Off-plane Grating Rocket Experiment

    Science.gov (United States)

    Donovan, Benjamin

    2018-01-01

    The next generation of X-ray spectrometers necessitate significant increases in both resolution and effective area to achieve the science goals set forth in the 2010 Decadal Survey and the 2013 Astrophysics Roadmap. The Off-plane Grating Rocket Experiment (OGRE), an X-ray spectroscopy suborbital rocket payload currently scheduled for launch in Q3 2020, will serve as a testbed for several key technologies which can help achieve the desired performance increases of future spectrometers. OGRE will be the first instrument to fly mono-crystalline silicon X-ray mirrors developed at NASA Goddard Space Flight Center. The payload will also utilize an array of off-plane gratings manufactured at The Pennsylvania State University. Additionally, the focal plane will be populated with an array of four electron-multiplying CCDs developed by the Open University and XCAM Ltd. With these key technologies, OGRE hopes to achieve the highest resolution on-sky soft X-ray spectrum to date. We discuss the optical design, expected performance, and the current status of the payload.

  1. Concentric traveling ionospheric disturbances triggered by the launch of a SpaceX Falcon 9 rocket

    Science.gov (United States)

    Lin, Charles C. H.; Shen, Ming-Hsueh; Chou, Min-Yang; Chen, Chia-Hung; Yue, Jia; Chen, Po-Cheng; Matsumura, Mitsuru

    2017-08-01

    We report the first observation of concentric traveling ionospheric disturbances (CTIDs) triggered by the launch of a SpaceX Falcon 9 rocket on 17 January 2016. The rocket-triggered ionospheric disturbances show shock acoustic wave signature in the time rate change (time derivative) of total electron content (TEC), followed by CTIDs in the 8-15 min band-pass filtering of TEC. The CTIDs propagated northward with phase velocity of 241-617 m/s and reached distances more than 1000 km away from the source on the rocket trajectory. The wave characteristics of CTIDs with periods of 10.5-12.7 min and wavelength 200-400 km agree well with the gravity wave dispersion relation. The optimal wave source searching and gravity wave ray tracing technique suggested that the CTIDs have multiple sources which are originated from 38-120 km altitude before and after the ignition of the second-stage rocket, 200 s after the rocket was launched.

  2. Measurement of IR atmospheric band dayglow by S-520-4 rocket

    International Nuclear Information System (INIS)

    Makino, Tadao; Yamamoto, Hiromasa; Sekiguchi, Hiroyuki

    1984-01-01

    The measurement of IR atmospheric band dayglow was made by rocket S-520-4 flown from Uchinoura at 1000 JST on Sept. 5, 1981. The instrument loaded on the rocket was the same type as the one loaded on EXOS-C satellite which will be launched in 1984 in order to observe the mesospheric ozone. This rocket experiment was performed for the purpose of testing the functions of this instrument in flight. The 1.27 μm filter radiometer consisted of three plane mirros, a camera lens, a chopper and a PbS detector array. The PbS array (4x5=20 elements) was operated at about -4 0 C with a thermoelectric cooler. We obtained the following results from the rocket experiment: (i) this instrument worked well during the flight, (ii) the intensities of the solar radiation scattered by the sea and clouds were obtained at 1.27 μm, and (iii) the baffle designed to permit the daytime measurement of the atmospheric emission could attenuate the off-axis radiation as weak as possible. The altitude distribution of the daytime mesospheric ozone density derived from the downleg data was in agreement with the previous profile obtained in twilight condition. (author)

  3. Non-destructive testing of rocket propellant quality using -X-ray radiography

    International Nuclear Information System (INIS)

    Arayaprecha, W.

    1979-01-01

    Currently, X-rays radiography has been used extensively in various industries. In this thesis, X-rays has been used in the study of compaction of rocket propellant. For a rocket, to gain an accurate guidance result, the propellant used must be mixed and compacted thoroughly. The quality control of the production of propellant sticks must be carefully done. In this study of non-destructive quality testing of rocket propellant, at first the ultrasonic rays was used to test its homogeneity. However, because the density of the propellant was too low, the test equipment could not detect any reflected signals from the propellant being tested. Then the new procedure using X-rays radiography was tried. The variables in the test procedure were voltage, amperage and the focal-film distance. Also different types of films were used. The results of this experiment were then used to construct an exposure chart for testing the homogeneity of the rocket propellant. The advantage of this chart is that a tester can use this table with propellant sticks of different sizes if they have similar density to the density specified in the chart. Also, it is not necessary that the mixture of the testing propellant be the same as the ones used to construct this chart

  4. Aluminum Agglomeration and Trajectory in Solid Rocket Motors

    National Research Council Canada - National Science Library

    Coats, Douglas; Hylin, E. C; Babbitt, Deborah; Tullos, James A; Beckstead, Merrill; Webb, Michael; Davis, I. L; Dang, Anthony

    2007-01-01

    Report developed under STTR contract for Topic AF06-T012. The demand for higher performance rocket motors at a reduced cost requires continuous improvements in understanding and controlling propellant combustion...

  5. Three-Axis Gasless Sounding Rocket Payload Attitude Control

    Data.gov (United States)

    National Aeronautics and Space Administration — Gas released by current sounding rocket payload attitude control systems (ACS) has the potential to interfere with some types of science instruments. A single-axis...

  6. 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

  7. 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

  8. Tutorial on nuclear thermal propulsion safety for Mars

    International Nuclear Information System (INIS)

    Buden, D.

    1992-01-01

    Safety is the prime design requirement for nuclear thermal propulsion (NTP). It must be built in at the initiation of the design process. An understanding of safety concerns is fundamental to the development of nuclear rockets for manned missions to Mars and many other applications that will be enabled or greatly enhanced by the use of nuclear propulsion. To provide an understanding of the basic issues, a tutorial has been prepared. This tutorial covers a range of topics including safety requirements and approaches to meet these requirements, risk and safety analysis methodology, NERVA reliability and safety approach, and life cycle risk assessments

  9. Space Processing Applications Rocket project, SPAR 1. Final report

    International Nuclear Information System (INIS)

    Reeves, F.; Chassay, R.

    1976-12-01

    The experiment objectives, design/operational concepts, and final results of each of nine scientific experiments conducted during the first Space Processing Applications Rocket (SPAR) flight are summarized. The nine individual SPAR experiments, covering a wide and varied range of scientific materials processing objectives, were entitled: solidification of Pb-Sb eutectic, feasibility of producing closed-cell metal foams, characterization of rocket vibration environment by measurement of mixing of two liquids, uniform dispersions of crystallization processing, direct observation of solidification as a function of gravity levels, casting thoria dispersion-strengthened interfaces, contained polycrystalline solidification, and preparation of a special alloy for manufacturing of magnetic hard superconductor under zero-g environment

  10. A Method of Initial Velocity Measurement for Rocket Projectile

    Directory of Open Access Journals (Sweden)

    Zhang Jiancheng

    2017-01-01

    Full Text Available In this paper, a novel method is proposed to measure the initial velocity of the rocket based on STFT (the short-time Fourier transform and the WT (wavelet transform. The radar echo signal processing procedure involves the following steps: sampling process, overlapping windows, wavelet decomposition and reconstruction, computing FFT (Fast Fourier Transform and spectrum analysis, power spectrum peak detection. Then, according to the peak of the detection power spectrum, the corresponding Doppler frequency is obtained. Finally, on the basis of the relationship between Doppler frequency and instantaneous velocity, the V-T curve is drawn in MATLAB to obtain the initial velocity of the rocket muzzle.

  11. 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.

  12. Infrasound and Seismic Recordings of Rocket Launches from Kennedy Space Center, 2016-2017

    Science.gov (United States)

    McNutt, S. R.; Thompson, G.; Brown, R. G.; Braunmiller, J.; Farrell, A. K.; Mehta, C.

    2017-12-01

    We installed a temporary 3-station seismic-infrasound network at Kennedy Space Center (KSC) in February 2016 to test sensor calibrations and train students in field deployment and data acquisitions techniques. Each station featured a single broadband 3-component seismometer and a 3-element infrasound array. In May 2016 the network was scaled back to a single station due to other projects competing for equipment. To date 8 rocket launches have been recorded by the infrasound array, as well as 2 static tests, 1 aborted launch and 1 rocket explosion (see next abstract). Of the rocket launches recorded 4 were SpaceX Falcon-9, 2 were ULA Atlas-5 and 2 were ULA Delta-IV. A question we attempt to answer is whether the rocket engine type and launch trajectory can be estimated with appropriate travel-time, amplitude-ratio and spectral techniques. For example, there is a clear Doppler shift in seismic and infrasound spectrograms from all launches, with lower frequencies occurring later in the recorded signal as the rocket accelerates away from the array. Another question of interest is whether there are relationships between jet noise frequency, thrust and/or nozzle velocity. Infrasound data may help answer these questions. We are now in the process of deploying a permanent seismic and infrasound array at the Astronaut Beach House. 10 more rocket launches are schedule before AGU. NASA is also conducting a series of 33 sonic booms over KSC beginning on Aug 21st. Launches and other events at KSC have provided rich sources of signals that are useful to characterize and gain insight into physical processes and wave generation from man-made sources.

  13. Flutter Analysis of RX-420 Balistic Rocket Fin Involving Rigid Body Modes of Rocket Structures

    Directory of Open Access Journals (Sweden)

    Novi Andria

    2013-03-01

    Full Text Available Flutter is a phenomenon that has brought a catastrophic failure to the flight vehicle structure. In this experiment, flutter was analyzed for its symmetric and antisymmetric configuration to understand the effect of rocket rigid modes to the fin flutter characteristic. This research was also expected to find out the safety level of RX-420 structure design. The analysis was performed using half rocket model. Fin structure used in this research was a fin which has semispan 600 mm, thickness 12 mm, chord root 700 mm, chord tip 400 mm, made by Al 6061-T651, double spar configuration with skin thickness of 2 mm. Structural dynamics and flutter stability were analyzed using finite element software implemented on MSC. Nastran. The analysis shows that the antisymmetric flutter mode is more critical than symmetric flutter mode. At sea level altitude, antisymmetric flutter occurs at 6.4 Mach, and symmetric flutter occurs at 10.15 Mach. Compared to maximum speed of RX-420 which is 4.5 Mach at altitude 11 km or equivalent to 2.1 Mach at sea level, it can be concluded that the RX-420 structure design is safe, and flutter will not occur during flight.

  14. Pulsating Heat pipe Only for Space (PHOS): results of the REXUS 18 sounding rocket campaign

    International Nuclear Information System (INIS)

    Creatini, F; Guidi, G M; Belfi, F; Cicero, G; Fioriti, D; Di Prizio, D; Piacquadio, S; Becatti, G; Orlandini, G; Frigerio, A; Fontanesi, S; Nannipieri, P; Rognini, M; Morganti, N; Filippeschi, S; Di Marco, P; Fanucci, L; Baronti, F; Mameli, M; Manzoni, M

    2015-01-01

    Two Closed Loop Pulsating Heat Pipes (CLPHPs) are tested on board REXUS 18 sounding rocket in order to obtain data over a relatively long microgravity period (approximately 90 s). The CLPHPs are partially filled with FC-72 and have, respectively, an inner tube diameter larger (3 mm) and slightly smaller (1.6 mm) than the critical diameter evaluated in static Earth gravity conditions. On ground, the small diameter CLPHP effectively works as a Pulsating Heat Pipe (PHP): the characteristic slug and plug flow pattern forms inside the tube and the heat exchange is triggered by thermally driven self-sustained oscillations of the working fluid. On the other hand, the large diameter CLPHP works as a two- phase thermosyphon in vertical position and doesn't work in horizontal position: in this particular condition, the working fluid stratifies within the device as the surface tension force is no longer able to balance buoyancy. Then, the idea to test the CLPHPs in reduced gravity conditions: as the gravity reduces the buoyancy forces becomes less intense and it is possible to recreate the typical PHP flow pattern also for larger inner tube diameters. This allows to increase the heat transfer rate and, consequently, to decrease the overall thermal resistance. Even though it was not possible to experience low gravity conditions due to a failure in the yoyo de-spin system, the thermal response to the peculiar acceleration field (hyper-gravity) experienced on board are thoroughly described. (paper)

  15. 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.

  16. The Water Recovery X-ray Rocket (WRX-R)

    Science.gov (United States)

    Miles, Drew

    2017-08-01

    The Water Recovery X-ray Rocket (WRX-R) is a diffuse soft X-ray spectrometer that will launch on a sounding rocket from the Kwajalein Atoll. WRX-R has a field of view of >10 deg2 and will observe the Vela supernova remnant. A mechanical collimator, state-of-the-art off-plane reflection grating array and hybrid CMOS detector will allow WRX to achieve the most highly-resolved spectrum of the Vela SNR ever recorded. In addition, this payload will fly a hard X-ray telescope that is offset from the soft X-ray spectrometer in order to observe the pulsar at the center of the remnant. We present here an introduction to the instrument, the expected science return, and an update on the state of the payload as we work towards launch.

  17. Assessment of the advantages and feasibility of a nuclear rocket

    International Nuclear Information System (INIS)

    Howe, S.D.

    1985-01-01

    The feasibility of rebuilding and testing a nuclear thermal rocket (NTR) for the Mars mission has been investigated. Calculations indicate that an NTR would substantially reduce the earth-orbit assembled mass compared to LOX/LH 2 systems. The mass savings were 36% and 65% for the cases of total aerobraking and of total propulsive braking respectively. Consequently, the cost savings for a single mission of using an NTR, if aerobraking is feasible, are probably insufficient to warrant the NTR development. If multiple missions are planned or if propulsive braking is desired at Mars and/or at Earth, then the savings of about $7B will easily pay for the NTR development. Estimates of the cost of rebuilding a NTR were based on the previous NERVA program's budget plus additional costs to develop a flight ready engine. The total cost to build the engine would be between $4 to 5B. The concept of developing a full-power test stand at Johnston Atoll in the Pacific appears very feasible. The added expense of building facilities on the island should be less than $1.4B

  18. Simultaneous rocket and radar measurements of currents in an auroral arc

    International Nuclear Information System (INIS)

    Robinson, R.M.; Bering, E.A.; Vondrak, R.R.; Anderson, H.R.; Cloutier, P.A.

    1981-01-01

    A detailed study of electric field, current and conductivities associated with an auroral arc was made in a coordinated rocket and radar experiment in Alaska on March 9, 1978. The payload, designated 29.007 UE, was launched at 1013 p.m. local time. It penetrated the diffuse aurora on the upleg and at apogee traversed field lines connected to a stable auroral arc of 40 kR intensity. Among the instruments carried by the payload were a vector magnetometer, a set of electrostatic double probes and a set of electron and proton spectrometers. Simultaneous electron density and line-of-sight velocity measurements were made by Chatanika radar operating in an elevation scan mode in the magnetic meridian plane. Both the radar and rocket measurements indicated that the zonal electric field was westward and approximately constant across the arc with a magnitude of about 7 mV/m. Small differences between the rocket and radar zonal electric field measurements indicated the presence of upward drifting ions in the region of the arc. The meridional field was large and northward equatorward of the arc, but negligible within the arc. Conductivities computed from measured fluxes of energetic electrons agreed well with the conductivities derived from the radar measureements of electron density. The electric field and conductivity measurements indicated that the zonal currents were eastward equatorward of the arc and westward within the arc. These electrojet currents agreed well with those inferred from the rocket magnetometer data. Better agreement was obtained when a westward neutral wind was added. The westward wind was also consistent with differences between the rocket and radar meridional electric fields. The meridional currents computed from the electric field measurements were northward over the entire region

  19. 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.

  20. Design Methodology and Performance Evaluation of New Generation Sounding Rockets

    Directory of Open Access Journals (Sweden)

    Marco Pallone

    2018-01-01

    Full Text Available Sounding rockets are currently deployed for the purpose of providing experimental data of the upper atmosphere, as well as for microgravity experiments. This work provides a methodology in order to design, model, and evaluate the performance of new sounding rockets. A general configuration composed of a rocket with four canards and four tail wings is sized and optimized, assuming different payload masses and microgravity durations. The aerodynamic forces are modeled with high fidelity using the interpolation of available data. Three different guidance algorithms are used for the trajectory integration: constant attitude, near radial, and sun-pointing. The sun-pointing guidance is used to obtain the best microgravity performance while maintaining a specified attitude with respect to the sun, allowing for experiments which are temperature sensitive. Near radial guidance has instead the main purpose of reaching high altitudes, thus maximizing the microgravity duration. The results prove that the methodology at hand is straightforward to implement and capable of providing satisfactory performance in term of microgravity duration.

  1. The Rocket Balloon (Rocketball): Applications to Science, Technology, and Education

    Science.gov (United States)

    Esper, Jaime

    2009-01-01

    Originally envisioned to study upper atmospheric phenomena, the Rocket Balloon system (or Rocketball for short) has utility in a range of applications, including sprite detection and in-situ measurements, near-space measurements and calibration correlation with orbital assets, hurricane observation and characterization, technology testing and validation, ground observation, and education. A salient feature includes the need to reach space and near-space within a critical time-frame and in adverse local meteorological conditions. It can also provide for the execution of technology validation and operational demonstrations at a fraction of the cost of a space flight. In particular, planetary entry probe proof-of-concepts can be examined. A typical Rocketball operational scenario consists of a sounding rocket launch and subsequent deployment of a balloon above a desired location. An obvious advantage of this combination is the additional mission 'hang-time' rendered by the balloon once the sounding rocket flight is completed. The system leverages current and emergent technologies at the NASA Goddard Space Flight Center and other organizations.

  2. 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.

  3. 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...

  4. Magnesium Based Rockets for Martian Exploration, Phase I

    Data.gov (United States)

    National Aeronautics and Space Administration — We propose to develop Mg rockets for Martian ascent vehicle applications. The propellant can be acquired in-situ from MgO in the Martian regolith (5.1% Mg by mass)...

  5. Convection measurement package for space processing sounding rocket flights. [low gravity manufacturing - fluid dynamics

    Science.gov (United States)

    Spradley, L. W.

    1975-01-01

    The effects on heated fluids of nonconstant accelerations, rocket vibrations, and spin rates, was studied. A system is discussed which can determine the influence of the convective effects on fluid experiments. The general suitability of sounding rockets for performing these experiments is treated. An analytical investigation of convection in an enclosure which is heated in low gravity is examined. The gravitational body force was taken as a time-varying function using anticipated sounding rocket accelerations, since accelerometer flight data were not available. A computer program was used to calculate the flow rates and heat transfer in fluids with geometries and boundary conditions typical of space processing configurations. Results of the analytical investigation identify the configurations, fluids and boundary values which are most suitable for measuring the convective environment of sounding rockets. A short description of fabricated fluid cells and the convection measurement package is given. Photographs are included.

  6. Color and textural quality of packaged wild rocket measured by multispectral imaging

    DEFF Research Database (Denmark)

    Løkke, Mette Marie; Seefeldt, Helene Fast; Skov, Thomas

    2013-01-01

    Green color and texture are important attributes for the perception of freshness of wild rocket. Packaging of green leafy vegetables can postpone senescence and yellowing, but a drawback is the risk of anaerobic respiration leading to loss of tissue integrity and development of an olive-brown color....... The hypothesis underlying this paper is that color and textural quality of packaged wild rocket leaves can be predicted by multispectral imaging for faster evaluation of visual quality of leafy green vegetables in scientific experiments. Multispectral imaging was correlated to sensory evaluation of packaged wild...... rocket quality. CIELAB values derived from the multispectral images and from a spectrophotometer changed during storage, but the data were insufficient to describe variation in sensory perceived color and texture. CIELAB values from the multispectral images allowed for a more detailed determination...

  7. Preliminary Results of the VLFE Quadrupole Instrumentation From The PARX Sounding Rocket

    Science.gov (United States)

    Reinleitner, L. A.; Holzworth, R. H.; Meadows, A. L.

    2003-12-01

    The NASA Pulsating Auroral Rocket eXperiment (PARX - March '97 from Poker Flat, AK) was equipped with 4 electric field probes oriented (X and Y) perpendicular to the ambient magnetic field, and one probe (along the Z axis) to obtain the parallel electric field. The rocket also included a three-axis VLF search coil magnetometer. The VLF measurements for both instruments were from 100 Hz - 8 KHz. Additionally, the electric field information was used onboard the rocket to obtain the "quadrupole" electric field, defined to be {(V1+V2) - (V3+V4)}/2d, which shows significant response only to short wavelength waves. This instrumentation clearly shows the long wavelength nature of features tentatively described as auroral hiss, and the shorter wavelength nature of the electrostatic and/or quasi-electrostatic waves.

  8. A new sounding rocket payload for solar plasma studies

    Science.gov (United States)

    Bruner, Marilyn E.; Brown, William A.; Appert, Kevin L.

    1989-01-01

    A sounding rocket payload developed for studies of high-temperature plasmas associated with solar active regions and flares is described. The payload instruments will record both spectra and images in the UV, EUV, and soft X-ray regions of the spectrum. The instruments, including the Dual Range Spectrograph, the Flat Field Soft X-ray Spectrograph, the Normal Incidence Soft X-ray Imager, the UV Filtergraph, and the H-alpha Imaging system, are described. Attention is also given to the new structural system of the payload, based on a large optical table suspended within the payload cavity, which will support the optical elements in their correct positions and orientations and will maintain these alignments throughout the rocket launch environment.

  9. Hard X-ray Vela supernova observation on rocket experiment WRX-R

    Science.gov (United States)

    Stehlikova, V.; Urban, M.; Nentvich, O.; Daniel, V.; Sieger, L.; Tutt, J.

    2017-07-01

    This paper presents a hard X-ray telescope for the Vela nebula observation during a sounding rocket flight. The Water Recovery X-ray Rocket (WRX-R) experiment is organised by the Pennsylvania State University (PSU), USA with a primary payload of a soft X-ray spectroscope. The Czech team developed a hard X-ray Lobster-eye telescope as a secondary payload. The Czech experiment’s astrophysical object of study is the Vela pulsar in the centre of the Vela nebula.

  10. Charging and the cross-field discharge during electron accelerator operation on a rocket

    International Nuclear Information System (INIS)

    Kellogg, P.J.; Monson, S.J.

    1988-01-01

    We present some limited results obtained from the flight of SCEX II, from Poker Flat, Alaska, on January 31, 1987. Some of the experiments were aimed at understanding neutralization processes around an electron beam emitting rocket. It was expected that electrons drifting in the strong electric fields around the charged rocket would acquire sufficient energy to ionize neutrals, and that the resulting ions would be hurled outward at energies up to the rocket potential. Three hemispherical retarding potential analyzers were ejected from the main payload to measure these ions. This experiment was successful, in spite of arcs which developed around the batteries for the electron guns, which degraded the emitted electron beam to unusable levels except for about 8 sec of the flight. Ions were observed at energies up to 175 eV, the limit of the analyzers. The main payload carried, in addition to the electron accelerator, two arms with conducting elements to act as Langmuir probes, and to measure floating potentials. These measurements show that fields sufficient to accelerate electrons to ionizing energies were present around the rocket. (author)

  11. Ceramic Matrix Composite (CMC) Thermal Protection Systems (TPS) and Hot Structures for Hypersonic Vehicles

    Science.gov (United States)

    Glass, David E.

    2008-01-01

    Thermal protection systems (TPS) and hot structures are required for a range of hypersonic vehicles ranging from ballistic reentry to hypersonic cruise vehicles, both within Earth's atmosphere and non-Earth atmospheres. The focus of this paper is on air breathing hypersonic vehicles in the Earth's atmosphere. This includes single-stage to orbit (SSTO), two-stage to orbit (TSTO) accelerators, access to space vehicles, and hypersonic cruise vehicles. This paper will start out with a brief discussion of aerodynamic heating and thermal management techniques to address the high heating, followed by an overview of TPS for rocket-launched and air-breathing vehicles. The argument is presented that as we move from rocket-based vehicles to air-breathing vehicles, we need to move away from the insulated airplane approach used on the Space Shuttle Orbiter to a wide range of TPS and hot structure approaches. The primary portion of the paper will discuss issues and design options for CMC TPS and hot structure components, including leading edges, acreage TPS, and control surfaces. The current state-of-the-art will be briefly discussed for some of the components. The two primary technical challenges impacting the use of CMC TPS and hot structures for hypersonic vehicles are environmental durability and fabrication, and will be discussed briefly.

  12. Improved Background Removal in Sounding Rocket Neutral Atom Imaging Data

    Science.gov (United States)

    Smith, M. R.; Rowland, D. E.

    2017-12-01

    The VISIONS sounding rocket, launched into a substorm on Feb 7, 2013 from Poker Flat, Alaska had a novel miniaturized energetic neutral atom (ENA) imager onboard. We present further analysis of the ENA data from this rocket flight, including improved removal of ultraviolet and electron contamination. In particular, the relative error source contributions due to geocoronal, auroral, and airglow UV, as well as energetic electrons from 10 eV to 3 keV were assessed. The resulting data provide a more clear understanding of the spatial and temporal variations of the ion populations that are energized to tens or hundreds of eV.

  13. FASTSAT-HSV01 Thermal Math Model Correlation

    Science.gov (United States)

    McKelvey, Callie

    2011-01-01

    This paper summarizes the thermal math model correlation effort for the Fast Affordable Science and Technology SATellite (FASTSAT-HSV01), which was designed, built and tested by NASA's Marshall Space Flight Center (MSFC) and multiple partners. The satellite launched in November 2010 on a Minotaur IV rocket from the Kodiak Launch Complex in Kodiak, Alaska. It carried three Earth science experiments and two technology demonstrations into a low Earth circular orbit with an inclination of 72deg and an altitude of 650 kilometers. The mission has been successful to date with science experiment activities still taking place daily. The thermal control system on this spacecraft was a passive design relying on thermo-optical properties and six heaters placed on specific components. Flight temperature data is being recorded every minute from the 48 Resistance Temperature Devices (RTDs) onboard the satellite structure and many of its avionics boxes. An effort has been made to correlate the thermal math model to the flight temperature data using Cullimore and Ring's Thermal Desktop and by obtaining Earth and Sun vector data from the Attitude Control System (ACS) team to create an "as-flown" orbit. Several model parameters were studied during this task to understand the spacecraft's sensitivity to these changes. Many "lessons learned" have been noted from this activity that will be directly applicable to future small satellite programs.

  14. 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.

  15. SCORE - Sounding-rocket Coronagraphic Experiment

    Science.gov (United States)

    Fineschi, Silvano; Moses, Dan; Romoli, Marco

    The Sounding-rocket Coronagraphic Experiment - SCORE - is a The Sounding-rocket Coronagraphic Experiment - SCORE - is a coronagraph for multi-wavelength imaging of the coronal Lyman-alpha lines, HeII 30.4 nm and HI 121.6 nm, and for the broad.band visible-light emission of the polarized K-corona. SCORE has flown successfully in 2009 acquiring the first images of the HeII line-emission from the extended corona. The simultaneous observation of the coronal Lyman-alpha HI 121.6 nm, has allowed the first determination of the absolute helium abundance in the extended corona. This presentation will describe the lesson learned from the first flight and will illustrate the preparations and the science perspectives for the second re-flight approved by NASA and scheduled for 2016. The SCORE optical design is flexible enough to be able to accommodate different experimental configurations with minor modifications. This presentation will describe one of such configurations that could include a polarimeter for the observation the expected Hanle effect in the coronal Lyman-alpha HI line. The linear polarization by resonance scattering of coronal permitted line-emission in the ultraviolet (UV) can be modified by magnetic fields through the Hanle effect. Thus, space-based UV spectro-polarimetry would provide an additional new tool for the diagnostics of coronal magnetism.

  16. Registration of ELF waves in rocket-satellite experiment with plasma injection

    Science.gov (United States)

    Korobeinikov, V. G.; Oraevskii, V. N.; Ruzhin, Iu. Ia.; Sobolev, Ia. P.; Skomarovskii, V. S.; Chmyrev, V. M.; Namazov, C. A.; Pokhunkov, A. A.; Nesmeianov, V. I.

    1992-12-01

    Two rocket KOMBI-SAMA experiments with plasma injection at height 100-240 km were performed in August 1987 in the region of Brazilian magnetic anomaly (L = 1.25). The launching time of the rocket was determined so that plasma injection was at the time when COSMOS 1809 satellite passed as close as possible to magnetic tube of injection. Caesium plasma jet was produced during not less than 300 s by an electric plasma generator separated from the payload. When the satellite passed the geomagnetic tube intersecting the injection region an enhancement of ELF emission at 140 Hz, 450 Hz by a factor of 2 was registered on board the satellite. An enhancement of energetic particle flux by a factor of 4-5 was registered on board the rocket. Observed ELF emission below 100 Hz is interpreted as the generation of oblique electromagnetic ion-cyclotron waves due to drift plasma instability at the front of the plasma jet.

  17. 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.

  18. Biodegradable protein-based rockets for drug transportation and light-triggered release.

    Science.gov (United States)

    Wu, Zhiguang; Lin, Xiankun; Zou, Xian; Sun, Jianmin; He, Qiang

    2015-01-14

    We describe a biodegradable, self-propelled bovine serum albumin/poly-l-lysine (PLL/BSA) multilayer rocket as a smart vehicle for efficient anticancer drug encapsulation/delivery to cancer cells and near-infrared light controlled release. The rockets were constructed by a template-assisted layer-by-layer assembly of the PLL/BSA layers, followed by incorporation of a heat-sensitive gelatin hydrogel containing gold nanoparticles, doxorubicin, and catalase. These rockets can rapidly deliver the doxorubicin to the targeted cancer cell with a speed of up to 68 μm/s, through a combination of biocatalytic bubble propulsion and magnetic guidance. The photothermal effect of the gold nanoparticles under NIR irradiation enable the phase transition of the gelatin hydrogel for rapid release of the loaded doxorubicin and efficient killing of the surrounding cancer cells. Such biodegradable and multifunctional protein-based microrockets provide a convenient and efficient platform for the rapid delivery and controlled release of therapeutic drugs.

  19. Rocket measurements of X-rays and energetic electrons through an auroral arc

    International Nuclear Information System (INIS)

    Aarsnes, K.; Stadsnes, J.; Soeraas, F.

    1976-01-01

    Preliminary results from rocket measurements on auroral electron precipitation are discussed as far as the spatial structure and time and space variations in the primary electron fluxes are concerned. The analysis demonstrates that there was a good overall correspondence between the X-ray and electron data. By using a well collimated X-ray detector on a spinning rocket, it was possible to get additional information on the overall electron precipitation pattern

  20. Measurements of auroral particles by means of sounding rockets of mother-daughter type

    International Nuclear Information System (INIS)

    Falck, A.

    1985-11-01

    The scientific objective of the S17 payloads was to study the ionosphere during auroral situations and especially with regards to the local fine structure and a possible separation of spatial and temporal variations of auroral phenomena. The intensities of 8 keV and 2 keV electrons have been measured from one sounding rocket launched into a breakup aurora of moderate activity and from another rocket launched into a very active substorm situation. Both the rockets were of mother-daughter type i.e. had two separated payloads. The general features in the data of different particle energies were very similar over the whole flight time of the rockets. Special events and gradients and well identifiable shapes in the particle intensities were studied to see if the intensity fluctuations obtained from two detectors in one payload or from detectors into separate payloads were time delayed. Such time delays in the particle flux intensities were obvious in both of the rocket measurements and most of these time shifts could be understood as caused by spatial variations in the particle precipitation. In parts of the rocket flights the particle intensity variations were true temporal changes. The time lags between 8 keV and 2 keV electron intensities detected in the same payload, which could be observed and were obtained by crosscorrelation analyses, were in the range less than 0.3 s and most of them less than 0.1 s. If the time differences are assumed to be caused by the velocity dispersion of the particles, the particle data reported here placed the modulation source at a distance of less than 10 000 km from the rocket position. Measurements at the S17-1 mother payload of the electric field have been compared with data of precipitating electrons and low-light-level-TV-recording of the auroral situation. An inverted-V precipitation event was observed and was associated with auroral arcs and with reversals of the measured electric field components implicating the possibility of

  1. 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.

  2. Evaluation of thermal sprayed metallic coatings for use on the structures at Launch Complex 39

    Science.gov (United States)

    Welch, Peter J.

    1990-01-01

    The current status of the evaluation program is presented. The objective was to evaluate the applicability of Thermal Sprayed Coatings (TSC) to protect the structures in the high temperature acid environment produced by exhaust of the Solid Rocket Boosters during the launches of the Shuttle Transportation System. Only the relatively low cost aluminum TSC which provides some cathodic protection for steel appears to be a practical candidate for further investigation.

  3. 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...

  4. 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...

  5. Development of efficient finite elements for structural integrity analysis of solid rocket motor propellant grains

    International Nuclear Information System (INIS)

    Marimuthu, R.; Nageswara Rao, B.

    2013-01-01

    Solid propellant rocket motors (SRM) are regularly used in the satellite launch vehicles which consist of mainly three different structural materials viz., solid propellant, liner, and casing materials. It is essential to assess the structural integrity of solid propellant grains under the specified gravity, thermal and pressure loading conditions. For this purpose finite elements developed following the Herrmann formulation are: twenty node brick element (BH20), eight node quadrilateral plane strain element (PH8) and, eight node axi-symmetric solid of revolution element (AH8). The time-dependent nature of the solid propellant grains is taken into account utilizing the direct inverse method of Schepary to specify the effective Young's modulus and Poisson's ratio. The developed elements are tested considering various problems prior to implementation in the in-house software package (viz., Finite Element Analysis of STructures, FEAST). Several SRM configurations are analyzed to assess the structural integrity under different loading conditions. Finite element analysis results are found to be in good agreement with those obtained earlier from MARC software. -- Highlights: • Developed efficient Herrmann elements. • Accuracy of finite elements demonstrated solving several known solution problems. • Time dependent structural response obtained using the direct inverse method of Schepary. • Performed structural analysis of grains under gravity, thermal and pressure loads

  6. 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.

  7. 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.

  8. 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.

  9. The Rocket Investigation of Current Closure in the Ionosphere (RICCI) mission: A novel application of CubeSats from a sounding rocket platform

    Science.gov (United States)

    Cohen, I. J.; Anderson, B. J.; Lessard, M.; Bonnell, J. W.; Bounds, S. R.; Lysak, R. L.; Erlandson, R. E.

    2017-12-01

    The transfer of energy and momentum between the terrestrial magnetosphere and ionosphere is substantially mediated by large-scale field-aligned currents (FACs), driven by magnetopause dynamics and magnetospheric pressures and closing through the ionosphere where the dissipation and drag are governed. While significant insight into ionospheric electrodynamics and the nature of magnetosphere-ionosphere (M-I) coupling have been gained by rocket and satellite measurements, in situ measurement of these ionospheric closure currents remains challenging. To date the best estimates of ionospheric current densities are inferred from ground-based radar observations combining electric fields calculated from drifts with conductivities derived from densities. RICCI aims to observe the structure of the ionospheric currents in situ to determine how the altitude structure of these currents is related to precipitation and density cavities, electromagnetic dynamics, and governs energy dissipation in the ionosphere. In situ measurement of the current density using multi-point measurements of the magnetic field requires precise attitude knowledge for which the only demonstrated technique is the use of star camera systems. The low vehicle rotation rates required for miniature commercial off-the-shelf (COTS) star cameras prohibit the use of available rocket sub-payload technologies at Wallops Flight Facility (WFF) which use high rates of spin to stabilize attitude. However, CubeSat attitude systems are already designed to achieve low vehicle rotation rates, so RICCI will use a set of three CubeSat sub-payloads deployed from a main low altitude payload with apogee of 160 km to provide precise current density measurement through the ionospheric closure altitude regime, together with a second rocket with apogee near 320 km to measure the incident input energy flux and convection electric field. The two rocket payloads and CubeSate sub-payloads are all instrumented with star cameras and

  10. Straw Rockets Are out of This World

    Science.gov (United States)

    Gillman, Joan

    2013-01-01

    To capture students' excitement and engage their interest in rocketships and visiting planets in the solar system, the author designed lessons that give students the opportunity to experience the joys and challenges of developing straw rockets, and then observing which design can travel the longest distance. The lessons are appropriate for…

  11. Nuclear thermal propulsion workshop overview

    International Nuclear Information System (INIS)

    Clark, J.S.

    1991-01-01

    NASA is planning an Exploration Technology Program as part of the Space Exploration Initiative to return U.S. astronauts to the moon, conduct intensive robotic exploration of the moon and Mars, and to conduct a piloted mission to Mars by 2019. Nuclear Propulsion is one of the key technology thrust for the human mission to Mars. The workshop addresses NTP (Nuclear Thermal Rocket) technologies with purpose to: assess the state-of-the-art of nuclear propulsion concepts; assess the potential benefits of the concepts for the mission to Mars; identify critical, enabling technologies; lay-out (first order) technology development plans including facility requirements; and estimate the cost of developing these technologies to flight-ready status. The output from the workshop will serve as a data base for nuclear propulsion project planning

  12. 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

  13. A new sounding rocket payload for solar plasma studies

    International Nuclear Information System (INIS)

    Bruner, M.E.; Brown, W.A.; Appert, K.L.

    1989-01-01

    A sounding rocket payload developed for studies of high-temperature plasmas associated with solar active regions and flares is described. The payload instruments will record both spectra and images in the UV, EUV, and soft X-ray regions of the spectrum. The instruments, including the Dual Range Spectrograph, the Flat Field Soft X-ray Spectrograph, the Normal Incidence Soft X-ray Imager, the UV Filtergraph, and the H-alpha Imaging system, are described. Attention is also given to the new structural system of the payload, based on a large optical table suspended within the payload cavity, which will support the optical elements in their correct positions and orientations and will maintain these alignments throughout the rocket launch environment. 8 refs

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

    Science.gov (United States)

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

    2010-01-01

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

  15. The National Aeronautics and Space Administration (NASA)/Goddard Space Flight Center (GSFC) sounding-rocket program

    Science.gov (United States)

    Guidotti, J. G.

    1976-01-01

    An overall introduction to the NASA sounding rocket program as managed by the Goddard Space Flight Center is presented. The various sounding rockets, auxiliary systems (telemetry, guidance, etc.), launch sites, and services which NASA can provide are briefly described.

  16. Rocket potential measurements during electron beam injection into the ionosphere

    International Nuclear Information System (INIS)

    Gringauz, K.I.; Shutte, N.M.

    1981-01-01

    Electron flux measurements were made during pulsed injection of electron beams at a current of about 0.5 A and energy of 15 or 27 keV, using a retarding potential analyzer which was mounted on the lateral surface of the Eridan rocket during the ARAKS experiment of January 26, 1975. The general character of the retardation curves was found to be the same regardless of the electron injection energy, and regardless of the fact whether the plasma generator, injecting quasineutral cesium plasma with an ion current of about 10 A, was switched on. A sharp current increase in the interval between 10 to the -7th and 10 to the -6th A was observed with a decrease of the retarding potential. The rocket potential did not exceed approximately 150 V at about 130 to 190 km, and decreased to 20 V near 100 km. This was explained by the formation of a highly conducting region near the rocket, which was formed via intense plasma waves generated by the beam. Measurements of electron fluxes with energies of 1 to 3 keV agree well with estimates based on the beam plasma discharge theory

  17. 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.

  18. Dynamical Model of Rocket Propellant Loading with Liquid Hydrogen

    Data.gov (United States)

    National Aeronautics and Space Administration — A dynamical model describing the multi-stage process of rocket propellant loading has been developed. It accounts for both the nominal and faulty regimes of...

  19. 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,...

  20. The behavior of fission products during nuclear rocket reactor tests

    International Nuclear Information System (INIS)

    Bokor, P.C.; Kirk, W.L.; Bohl, R.J.

    1991-01-01

    Fission product release from nuclear rocket propulsion reactor fuel is an important consideration for nuclear rocket development and application. Fission product data from the last six reactors of the Rover program are collected in this paper to provide as basis for addressing development and testing issues. Fission product loss from the fuel will depend on fuel composition and reactor design and operating parameters. During ground testing, fission products can be contained downstream of the reactor. The last Rover reactor tested, the Nuclear Furnance, was mated to an effluent clean-up system that was effective in preventing the discharge of fission products into the atmosphere