WorldWideScience

Sample records for hydrocarbon-fueled rocket engines

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

    Science.gov (United States)

    Tejwani, Gopal D.

    2010-01-01

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

  2. Raman Gas Species Measurements in Hydrocarbon-Fueled Rocket Engine Injector Flows

    Science.gov (United States)

    Wehrmeyer, Joseph; Hartfield, Roy J., Jr.; Trinh, Huu P.; Dobson, Chris C.; Eskridge, Richard H.

    2000-01-01

    Rocket engine propellent injector development at NASA-Marshall includes experimental analysis using optical techniques, such as Raman, fluorescence, or Mie scattering. For the application of spontaneous Raman scattering to hydrocarbon-fueled flows a technique needs to be developed to remove the interfering polycyclic aromatic hydrocarbon fluorescence from the relatively weak Raman signals. A current application of such a technique is to the analysis of the mixing and combustion performance of multijet, impinging-jet candidate fuel injectors for the baseline Mars ascent engine, which will burn methane and liquid oxygen produced in-situ on Mars to reduce the propellent mass transported to Mars for future manned Mars missions. The Raman technique takes advantage of the strongly polarized nature of Raman scattering. It is shown to be discernable from unpolarized fluorescence interference by subtracting one polarized image from another. Both of these polarized images are obtained from a single laser pulse by using a polarization-separating calcite rhomb mounted in the imaging spectrograph. A demonstration in a propane-air flame is presented, as well as a high pressure demonstration in the NASA-Marshall Modular Combustion Test Artice, using the liquid methane-liquid oxygen propellant system

  3. Design and evaluation of high performance rocket engine injectors for use with hydrocarbon fuels

    Science.gov (United States)

    Pavli, A. J.

    1979-01-01

    An experimental program to determine the feasibility of using a heavy hydrocarbon fuel as a rocket propellant is reported herein. A method of predicting performance of a heavy hydrocarbon in terms of vaporization effectiveness is described and compared to other fuels and to experimental test results. The work was done at a chamber pressure of 4137 KN/sq M (600 psia) with RP-1, JP-10, and liquefied natural gas as fuels, and liquid oxygen as the oxidizer. Combustion length effects were explored over a range of 21.6 cm (8 1/2 in.) to 55.9 cm (22 in.). Four injector types were tested, each over a range of mixture ratios. Further configuration modifications were obtained by 'reaming' each injector several times to provide test data over a range of injector pressure drop.

  4. Design and evaluation of high performance rocket engine injectors for use with hydrocarbon fuels

    Science.gov (United States)

    Pavli, A. J.

    1979-01-01

    The feasibility of using a heavy hydrocarbon fuel as a rocket propellant is examined. A method of predicting performance of a heavy hydrocarbon in terms of vaporization effectiveness is described and compared to other fuels and to experimental test results. Experiments were done at a chamber pressure of 4137 KN/sq M (600 psia) with RP-1, JP-10, and liquefied natural gas as fuels, and liquid oxygen as the oxidizer. Combustion length effects were explored over a range of 21.6 cm (8 1/2 in) to 55.9 cm (22 in). Four injector types were tested, each over a range of mixture ratios. Further configuration modifications were obtained by reaming each injector several times to provide test data over a range of injector pressure drop.

  5. Calculation of Free-Atom Fractions in Hydrocarbon-Fueled Rocket Engine Plume

    Science.gov (United States)

    Verma, Satyajit

    2006-01-01

    Free atom fractions (Beta) of nine elements are calculated in the exhaust plume of CH4- oxygen and RP-1-oxygen fueled rocket engines using free energy minimization method. The Chemical Equilibrium and Applications (CEA) computer program developed by the Glenn Research Center, NASA is used for this purpose. Data on variation of Beta in both fuels as a function of temperature (1600 K - 3100 K) and oxygen to fuel ratios (1.75 to 2.25 by weight) is presented in both tabular and graphical forms. Recommendation is made for the Beta value for a tenth element, Palladium. The CEA computer code was also run to compare with experimentally determined Beta values reported in literature for some of these elements. A reasonable agreement, within a factor of three, between the calculated and reported values is observed. Values reported in this work will be used as a first approximation for pilot rocket engine testing studies at the Stennis Space Center for at least six elements Al, Ca, Cr, Cu, Fe and Ni - until experimental values are generated. The current estimates will be improved when more complete thermodynamic data on the remaining four elements Ag, Co, Mn and Pd are added to the database. A critique of the CEA code is also included.

  6. Metabolic engineering for the production of hydrocarbon fuels.

    Science.gov (United States)

    Lee, Sang Yup; Kim, Hye Mi; Cheon, Seungwoo

    2015-06-01

    Biofuels have been attracting increasing attention to provide a solution to the problems of climate change and our dependence on limited fossil oil. During the last decade, metabolic engineering has been performed to develop superior microorganisms for the production of so called advanced biofuels. Among the advanced biofuels, hydrocarbons possess high-energy content and superior fuel properties to other biofuels, and thus have recently been attracting much research interest. Here we review the recent advances in the microbial production of hydrocarbon fuels together with the metabolic engineering strategies employed to develop their production strains. Strategies employed for the production of long-chain and short-chain hydrocarbons derived from fatty acid metabolism along with the isoprenoid-derived hydrocarbons are reviewed. Also, the current limitations and future prospects in hydrocarbon-based biofuel production are discussed.

  7. Introduction to rocket science and engineering

    CERN Document Server

    Taylor, Travis S

    2009-01-01

    What Are Rockets? The History of RocketsRockets of the Modern EraRocket Anatomy and NomenclatureWhy Are Rockets Needed? Missions and PayloadsTrajectoriesOrbitsOrbit Changes and ManeuversBallistic Missile TrajectoriesHow Do Rockets Work? ThrustSpecific ImpulseWeight Flow RateTsiolkovsky's Rocket EquationStagingRocket Dynamics, Guidance, and ControlHow Do Rocket Engines Work? The Basic Rocket EngineThermodynamic Expansion and the Rocket NozzleExit VelocityRocket Engine Area Ratio and LengthsRocket Engine Design ExampleAre All Rockets the Same? Solid Rocket EnginesLiquid Propellant Rocket Engines

  8. Study of NOx Emissions of S.I. Engine Fueled with Different Kinds of Hydrocarbon Fuels and Hydrogen

    Directory of Open Access Journals (Sweden)

    Qahtan A. Abass

    2010-01-01

    Full Text Available Liquefied petroleum gas (LPG, Natural gas (NG and hydrogen were used to operate spark ignition internal combustion engine Ricardo E6, to compare NOx emissions emitted from the engine, with that emitted from engine fueled with gasoline as a fuel.The study was done when engine operated at HUCR for gasoline, compared with its operation at HUCR for each fuel. Compression ratio, equivalence ratio and spark timing were studied at constant speed 25rps.The results appeared that NOx concentrations will be at maximum value in the lean side near the stoichiometric ratio, and reduced with moving away from this ratio for mixture at both sides, these concentrations were at its highest value when hydrogen used at CR=8:1, and got near to each other for the three hydrocarbon fuels used in the study, when the engine operated at HUCR for each fuel, but still hydrogen had maximum value, the main variable affect these concentrations was spark timing

  9. Liquid rocket engine injectors

    Science.gov (United States)

    Gill, G. S.; Nurick, W. H.

    1976-01-01

    The injector in a liquid rocket engine atomizes and mixes the fuel with the oxidizer to produce efficient and stable combustion that will provide the required thrust without endangering hardware durability. Injectors usually take the form of a perforated disk at the head of the rocket engine combustion chamber, and have varied from a few inches to more than a yard in diameter. This monograph treats specifically bipropellant injectors, emphasis being placed on the liquid/liquid and liquid/gas injectors that have been developed for and used in flight-proven engines. The information provided has limited application to monopropellant injectors and gas/gas propellant systems. Critical problems that may arise during injector development and the approaches that lead to successful design are discussed.

  10. Liquid Rocket Engine Testing

    Science.gov (United States)

    2016-10-21

    booster rocket engines • 6000-10000 psia capabilities – Can use gaseous nitrogen, helium, or hydrogen to pressurize propellant tanks 9Distribution A...Approved for Public Release; Distribution Unlimited. PA Clearance 16493 Simplified Test Stand Layout Oxidizer  TankFuel  Tank High  Pressure   Gas (GN2...requires large, complex facilities to deliver propellant at the proper pressure , temperature, and flow rates • The enormous energies involved

  11. Problems of providing completeness of the methane-containing block-jet combustion in a rocket-ramjet engine's combustion chamber

    Science.gov (United States)

    Timoshenko, Valeriy I.; Belotserkovets, Igor S.; Gusinin, Vjacheslav P.

    2009-11-01

    Some problems of methane-containing hydrocarbon fuel combustion are discussed. It seems that reduction of methane burnout zone length is one from main problems of designing new type engine. It is very important at the creation of combustion chambers of a rocket-ramjet engine for prospective space shuttle launch vehicles.

  12. 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.

  13. Combustion characteristics of thermally stressed hydrocarbon fuels

    Science.gov (United States)

    Curtis, Colin William

    Liquid propelled propulsion systems, which range from rocket systems to hypersonic scramjet and ramjet engines, require active cooling in order to prevent additional payload requirements. In these systems, the liquid fuel is used as a coolant and is delivered through micro-channels that surround the combustion chambers, nozzles, as well as the exterior surfaces in order to extract heat from these affected areas. During this process, heat exchange occurs through phase change, sensible heat extraction, and endothermic reactions experienced by the liquid fuel. Previous research has demonstrated the significant modifications in fuel composition and changes to the fuel's physical properties that can result from these endothermic reactions. As a next step, we are experimentally investigating the effect that endothermic reactions have on fundamental flame behavior for real hydrocarbon fuels that are used as rocket and jet propellants. To achieve this goal, we have developed a counter-flow flame burner to measure extinction limits of the thermally stressed fuels. The counter-flow flame system is to be coupled with a high pressure reactor, capable of subjecting the fuel to 170 atm and 873 K, effectively simulating the extreme environment that cause the liquid fuel to experience endothermic reactions. The fundamental flame properties of the reacted fuels will be compared to those of unreacted fuels, allowing us to determine the role of endothermic reactions on the combustion behavior of current hydrocarbon jet and rocket propellants. To quantify the change in transport properties and chemical kinetics of the reacting mixture, simultaneous numerical simulations of the reactor portion of the experiment coupled with a counterflow flame simulation are performed using n-heptane and n-dodecane.

  14. CFD Simulation of Liquid Rocket Engine Injectors

    Science.gov (United States)

    Farmer, Richard; Cheng, Gary; Chen, Yen-Sen; Garcia, Roberto (Technical Monitor)

    2001-01-01

    Detailed design issues associated with liquid rocket engine injectors and combustion chamber operation require CFD methodology which simulates highly three-dimensional, turbulent, vaporizing, and combusting flows. The primary utility of such simulations involves predicting multi-dimensional effects caused by specific injector configurations. SECA, Inc. and Engineering Sciences, Inc. have been developing appropriate computational methodology for NASA/MSFC for the past decade. CFD tools and computers have improved dramatically during this time period; however, the physical submodels used in these analyses must still remain relatively simple in order to produce useful results. Simulations of clustered coaxial and impinger injector elements for hydrogen and hydrocarbon fuels, which account for real fluid properties, is the immediate goal of this research. The spray combustion codes are based on the FDNS CFD code' and are structured to represent homogeneous and heterogeneous spray combustion. The homogeneous spray model treats the flow as a continuum of multi-phase, multicomponent fluids which move without thermal or velocity lags between the phases. Two heterogeneous models were developed: (1) a volume-of-fluid (VOF) model which represents the liquid core of coaxial or impinger jets and their atomization and vaporization, and (2) a Blob model which represents the injected streams as a cloud of droplets the size of the injector orifice which subsequently exhibit particle interaction, vaporization, and combustion. All of these spray models are computationally intensive, but this is unavoidable to accurately account for the complex physics and combustion which is to be predicted, Work is currently in progress to parallelize these codes to improve their computational efficiency. These spray combustion codes were used to simulate the three test cases which are the subject of the 2nd International Workshop on-Rocket Combustion Modeling. Such test cases are considered by

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

    Directory of Open Access Journals (Sweden)

    A. V. Rudinskii

    2015-01-01

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

  16. Summarization on variable liquid thrust rocket engines

    Institute of Scientific and Technical Information of China (English)

    2009-01-01

    The technology actuality and development trend of variable thrust rocket engines at home and abroad are summarized. Key technologies of developing variable thrust rocket engines are analyzed. Development advices on developing variable thrust rocket engines that are adapted to the situation of our country are brought forward.

  17. Progress Toward Quality Assurance Standards for Advanced Hydrocarbon Fuels Based on Thermal Performance Testing and Chemometric Modeling

    Science.gov (United States)

    2015-12-15

    analyses; (2) Identify class distinguishing features, i.e., chemical compounds in the chromatographic data, that contribute to a fuel’s group ...implemented primarily to identify distinguishing chemical compounds that contribute to a fuel’s group assignment. F-ratio analysis was performed using...hydrocarbon-fueled liquid rocket engines, combustion enthalpy is transferred at high rates to thrust chamber surfaces, which are maintained at acceptably

  18. Unique nuclear thermal rocket engine

    Energy Technology Data Exchange (ETDEWEB)

    Culver, D.W. (Aerojet Propulsion Division, P.O. Box 13222, Sacramento, California 95813-6000 (United States)); Rochow, R. (Babcock Wilcox Space Nuclear Systems, P.O. Box 11165, Lynchburg, Virginia 24506-1165 (United States))

    1993-01-15

    Earlier this year Aerojet Propulsion Division (APD) introduced a new, advanced nuclear thermal rocket engine (NTRE) concept intended for manned missions to the moon and to Mars. 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 (E-D) 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)Reactor design and cooling scheme simplification while retaining a high pressure power balance option; (2)Eliminate need for a new, uncooled nozzle throat material suitable for long life application; (3)Practical provision for reactor power control; and (4)Use near term, long life turbopumps.

  19. Unique nuclear thermal rocket engine

    Science.gov (United States)

    Culver, Donald W.; Rochow, Richard

    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.

  20. 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…

  1. 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…

  2. Reusable rocket engine optical condition monitoring

    Science.gov (United States)

    Wyett, L.; Maram, J.; Barkhoudarian, S.; Reinert, J.

    1987-01-01

    Plume emission spectrometry and optical leak detection are described as two new applications of optical techniques to reusable rocket engine condition monitoring. Plume spectrometry has been used with laboratory flames and reusable rocket engines to characterize both the nominal combustion spectra and anomalous spectra of contaminants burning in these plumes. Holographic interferometry has been used to identify leaks and quantify leak rates from reusable rocket engine joints and welds.

  3. A rocket-based combined-cycle engine prototype demonstrating comprehensive component compatibility and effective mode transition

    Science.gov (United States)

    Shi, Lei; He, Guoqiang; Liu, Peijin; Qin, Fei; Wei, Xianggeng; Liu, Jie; Wu, Lele

    2016-11-01

    A rocket-based combined cycle (RBCC) engine was designed to demonstrate its broad applicability in the ejector and ramjet modes within the flight range from Mach 0 to Mach 4.5. To validate the design, a prototype was fabricated and tested as a freejet engine operating at flight Mach 3 using hydrocarbon fuel. The proposed design was a single module, heat sink steel alloy model with an interior fuel supply and active control system and a fully integrated flowpath that was comprehensively instrumented with pressure sensors. The mass capture and back pressure resistance of the inlet were numerically investigated and experimentally calibrated. The combustion process and rocket operation during mode transition were investigated by direct-connect tests. Finally, the comprehensive component compatibility and multimodal operational capability of the RBCC engine prototype was validated through freejet tests. This paper describes the design of the RBCC engine prototype, reviews the testing procedures, and discusses the experimental results of these efforts in detail.

  4. Fire-safe hydrocarbon fuels

    Energy Technology Data Exchange (ETDEWEB)

    Fodor, G.E.; Weatherford, W.D. Jr.; Wright, B.R.

    1979-11-06

    A stabilized, fire-safe, aqueous hydrocarbon fuel emulsion prepared by mixing: a diesel fuel; an emulsifier (consisting of oleyl diethanolamide, diethanolamine, and diethanolamine soap of oleic acid) which has been treated with about 0 to 7 1/2 of oleic acid. A modified version of this fuel also contains 0 to 0.5% of an antimisting agent, and water.

  5. Experimental investigation on laser-induced plasma ignition of hydrocarbon fuel in scramjet engine at takeover flight conditions

    Science.gov (United States)

    Li, Xipeng; Liu, Weidong; Pan, Yu; Yang, Leichao; An, Bin

    2017-09-01

    Laser-induced plasma ignition of an ethylene fuelled cavity is successfully conducted in a model scramjet engine combustor with dual cavities. The simulated flight condition corresponds to takeover flight Mach 4, with isolator entrance Mach number of 2.1, the total pressure of 0.65 MPa and stagnation temperature of 947 K. Ethylene is injected 35 mm upstream of cavity flameholder from four orifices with 2-mm-diameter. The 1064 nm laser beam, from a Q-switched Nd:YAG laser source running at 10 Hz and 940 mJ per pulse, is focused into cavity for ignition. High speed photography is used to capture the transient ignition process. The laser-induced gas breakdown, flame kernel generation and propagation are all recorded and ensuing stable supersonic combustion is established in cavity. The highly ionized plasma zone is almost round at starting, and then the surface of the flame kernel is wrinkled severely in 150 μs after the laser pulse due to the strong turbulence flow in cavity. The flame kernel is found rotating anti-clockwise and gradually moves upstream as the entrainment of circulation flow in cavity. The flame is stabilized at the corner of the cavity for about 200 μs, and then spreads from leading edge to trailing edge via the under part of shear layer to fully fill the entire cavity. The corner recirculation zone of cavity is of great importance for flame spreading. Eventually, a cavity shear-layer stabilized combustion is established in the supersonic flow roughly 2.9 ms after the laser pulse. Both the temporal evolution of normalized chemiluminescence intensity and normalized flame area show that the entire ignition process can be divided into four stages, which are referred as turbulent dissipation stage, combustion enhancement stage, reverting stage and combustion stabilization stage. The results show promising potentials of laser induced plasma for ignition in real scramjets.

  6. 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.

  7. Advanced Vortex Hybrid Rocket Engine (AVHRE) Project

    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...

  8. Advanced Vortex Hybrid Rocket Engine (AVHRE) Project

    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...

  9. Electrodynamic actuators for rocket engine valves

    Science.gov (United States)

    Fiet, O.; Doshi, D.

    1972-01-01

    Actuators, employed in acoustic loudspeakers, operate liquid rocket engine valves by replacing light paper cones with flexible metal diaphragms. Comparative analysis indicates better response time than solenoid actuators, and improved service life and reliability.

  10. 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.

  11. Deep desulfurization of hydrocarbon fuels

    Science.gov (United States)

    Song, Chunshan [State College, PA; Ma, Xiaoliang [State College, PA; Sprague, Michael J [Calgary, CA; Subramani, Velu [State College, PA

    2012-04-17

    The invention relates to processes for reducing the sulfur content in hydrocarbon fuels such as gasoline, diesel fuel and jet fuel. The invention provides a method and materials for producing ultra low sulfur content transportation fuels for motor vehicles as well as for applications such as fuel cells. The materials and method of the invention may be used at ambient or elevated temperatures and at ambient or elevated pressures without the need for hydrogen.

  12. LIQUID HYDROCARBON FUEL CELL DEVELOPMENT.

    Science.gov (United States)

    A compound anode consists of a reforming catalyst bed in direct contact with a palladium-silver fuel cell anode. The objective of this study was to...prove the feasibility of operating a compound anode fuel cell on a liquid hydrocarbon and to define the important parameters that influence cell...performance. Both reformer and fuel cell tests were conducted with various liquid hydrocarbon fuels. Included in this report is a description of the

  13. Heat Transfer and Thermal Stability Research for Advanced Hydrocarbon Fuel Technologies

    Science.gov (United States)

    DeWitt, Kenneth; Stiegemeier, Benjamin

    2005-01-01

    In recent years there has been increased interest in the development of a new generation of high performance boost rocket engines. These efforts, which will represent a substantial advancement in boost engine technology over that developed for the Space Shuttle Main Engines in the early 1970s, are being pursued both at NASA and the United States Air Force. NASA, under its Space Launch Initiative s Next Generation Launch Technology Program, is investigating the feasibility of developing a highly reliable, long-life, liquid oxygen/kerosene (RP-1) rocket engine for launch vehicles. One of the top technical risks to any engine program employing hydrocarbon fuels is the potential for fuel thermal stability and material compatibility problems to occur under the high-pressure, high-temperature conditions required for regenerative fuel cooling of the engine combustion chamber and nozzle. Decreased heat transfer due to carbon deposits forming on wetted fuel components, corrosion of materials common in engine construction (copper based alloys), and corrosion induced pressure drop increases have all been observed in laboratory tests simulating rocket engine cooling channels. To mitigate these risks, the knowledge of how these fuels behave in high temperature environments must be obtained. Currently, due to the complexity of the physical and chemical process occurring, the only way to accomplish this is empirically. Heated tube testing is a well-established method of experimentally determining the thermal stability and heat transfer characteristics of hydrocarbon fuels. The popularity of this method stems from the low cost incurred in testing when compared to hot fire engine tests, the ability to have greater control over experimental conditions, and the accessibility of the test section, facilitating easy instrumentation. These benefits make heated tube testing the best alternative to hot fire engine testing for thermal stability and heat transfer research. This investigation

  14. 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

  15. Regenerative Cooling for Liquid Rocket Engines

    Institute of Scientific and Technical Information of China (English)

    QiFeng

    1995-01-01

    Heat transfer in the thrust chamber is of great importance in the design of liquid propellant rocket engines.Regenerative cooling is and advanced method which can ensure not only the proper running but also higher performance of a rocket engine.The theoretical model is complicated,it relates to fluid bynamics,heat transfer,combustion.etc…,In this paper,a regenerative cooling model is presented.Effects such as radiation,heat transfer to environment,variable thermal properties and coking are included in the model.This model can be applied to all kinds of liquid propellant rocket engines as well as similar constructions.The modularized computer code is completed in the work.

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

    Data.gov (United States)

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

  17. Influence of Rocket Engine Characteristics on Shaft Sealing Technology Needs

    Science.gov (United States)

    Keba, John E.

    1999-01-01

    This paper presents viewgraphs of The Influence of Rocket Engine Characteristics on Shaft Sealing Technology Needs. The topics include: 1) Rocket Turbomachinery Shaft Seals (Inter-Propellant-Seal (IPS) Systems, Lift-off Seal Systems, and Technology Development Needs); 2) Rocket Engine Characteristics (Engine cycles, propellants, missions, etc., Influence on shaft sealing requirements); and 3) Conclusions.

  18. Conversion of hydrocarbon fuel in thermal protection reactors of hypersonic aircraft

    Science.gov (United States)

    Kuranov, A. L.; Mikhaylov, A. M.; Korabelnikov, A. V.

    2016-07-01

    Thermal protection of heat-stressed surfaces of a high-speed vehicle flying in dense layers of atmosphere is one of the topical issues. Not of a less importance is also the problem of hydrocarbon fuel combustion in a supersonic air flow. In the concept under development, it is supposed that in the most high-stressed parts of airframe and engine, catalytic thermochemical reactors will be installed, wherein highly endothermic processes of steam conversion of hydrocarbon fuel take place. Simultaneously with heat absorption, hydrogen generation will occur in the reactors. This paper presents the results of a study of conversion of hydrocarbon fuel in a slit reactor.

  19. The next generation rocket engines

    Science.gov (United States)

    Beichel, Rudi; O'Brien, Charles J.; Taylor, James P.

    This paper examines propulsion system technologies for earth-to-orbit vehicles, and describes several propulsion system concepts which could support the recommendations of the Commission for Space Development for the year 2000. The hallmark of that system must and will be reliability. Reliability will be obtained through a very structured design approach, coupled with a rational, cost effective, development and qualification program. To improve the next generation space transportation propulsion systems we need to select the very best of alternative power and performance cycles and engine physical concepts with a rigid requirement to achieve a robust, dependable, affordable propulsion system. For example, engine concepts using either propellants or non-propellant fluids for cooling and/or power drive offer the potential to provide smooth, controlled engine starts, low turbine temperatures, etc. as required for long life turbomachinery. Concepts examined are LOX/LH 2, |LOX/LH 2 + hydrocarbon, and LOX/LH 2 + hydrocarbon + Al dual expander engines, separate LOX/LH 2 and LOX/hydrocarbon engines, and variable mixture ratio engines. A fully reusable propulsion system that is perceived to be very low risk and low in operation cost is described.

  20. Nuclear thermal rocket engine operation and control

    Science.gov (United States)

    Gunn, Stanley V.; Savoie, Margarita T.; Hundal, Rolv

    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.

  1. Analysis of a Radioisotope Thermal Rocket Engine

    Science.gov (United States)

    Machado-Rodriguez, Jonathan P.; Landis, Geoffrey A.

    2017-01-01

    The Triton Hopper is a concept for a vehicle to explore the surface of Neptunes moon Triton, which uses a radioisotope heated rocket engine and in-situ propellant acquisition. The initial Triton Hopper conceptual design stores pressurized Nitrogen in a spherical tank to be used as the propellant. The aim of the research was to investigate the benefits of storing propellant at ambient temperature and heating it through a thermal block during engine operation, as opposed to storing gas at a high temperature.

  2. Software for Collaborative Engineering of Launch Rockets

    Science.gov (United States)

    Stanley, Thomas Troy

    2003-01-01

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

  3. Rocket Engine Innovations Advance Clean Energy

    Science.gov (United States)

    2012-01-01

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

  4. Parallelization of Rocket Engine System Software (Press)

    Science.gov (United States)

    Cezzar, Ruknet

    1996-01-01

    The main goal is to assess parallelization requirements for the Rocket Engine Numeric Simulator (RENS) project which, aside from gathering information on liquid-propelled rocket engines and setting forth requirements, involve a large FORTRAN based package at NASA Lewis Research Center and TDK software developed by SUBR/UWF. The ultimate aim is to develop, test, integrate, and suitably deploy a family of software packages on various aspects and facets of rocket engines using liquid-propellants. At present, all project efforts by the funding agency, NASA Lewis Research Center, and the HBCU participants are disseminated over the internet using world wide web home pages. Considering obviously expensive methods of actual field trails, the benefits of software simulators are potentially enormous. When realized, these benefits will be analogous to those provided by numerous CAD/CAM packages and flight-training simulators. According to the overall task assignments, Hampton University's role is to collect all available software, place them in a common format, assess and evaluate, define interfaces, and provide integration. Most importantly, the HU's mission is to see to it that the real-time performance is assured. This involves source code translations, porting, and distribution. The porting will be done in two phases: First, place all software on Cray XMP platform using FORTRAN. After testing and evaluation on the Cray X-MP, the code will be translated to C + + and ported to the parallel nCUBE platform. At present, we are evaluating another option of distributed processing over local area networks using Sun NFS, Ethernet, TCP/IP. Considering the heterogeneous nature of the present software (e.g., first started as an expert system using LISP machines) which now involve FORTRAN code, the effort is expected to be quite challenging.

  5. Liquid fuel injection elements for rocket engines

    Science.gov (United States)

    Cox, George B., Jr. (Inventor)

    1993-01-01

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

  6. 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.

  7. Advancing the State-of-the-Practice for Liquid Rocket Engine Injector Design

    Science.gov (United States)

    Tucker, P. K.; Kenny, R. J.; Richardson, B. R.; Anderso, W. E.; Austin, B. J.; Schumaker, S. A.; Muss, J. A.

    2015-01-01

    Current shortcomings in both the overall injector design process and its underlying combustion stability assessment methodology are rooted in the use of empirically based or low fidelity representations of complex physical phenomena and geometry details that have first order effects on performance, thermal environments and combustion stability. The result is a design and analysis capability that is often inadequate to reliably arrive at a suitable injector design in an efficient manner. Specifically, combustion instability has been particularly difficult to predict and mitigate. Large hydrocarbon-fueled booster engines have been especially problematic in this regard. Where combustion instability has been a problem, costly and time-consuming redesign efforts have often been an unfortunate consequence. This paper presents an overview of a recently completed effort at NASA Marshall Space Flight Center to advance the state-of-the-practice for liquid rocket engine injector design. Multiple perturbations of a gas-centered swirl coaxial (GCSC) element that burned gaseous oxygen and RP-1 were designed, assessed for combustion stability, and tested. Three designs, one stable, one marginally unstable and one unstable, were used to demonstrate both an enhanced overall injector design process and an improved combustion stability assessment process. High-fidelity results from state-of-the-art computational fluid dynamics CFD simulations were used to substantially augment and improve the injector design methodology. The CFD results were used to inform and guide the overall injector design process. They were also used to upgrade selected empirical or low-dimensional quantities in the ROCket Combustor Interactive Design (ROCCID) stability assessment tool. Hot fire single element injector testing was used to verify both the overall injector designs and the stability assessments. Testing was conducted at the Air Force Research Laboratory and at Purdue University. Companion papers

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

    OpenAIRE

    2015-01-01

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

  9. Numerical heat transfer analysis of transcritical hydrocarbon fuel flow in a tube partially filled with porous media

    Directory of Open Access Journals (Sweden)

    Jiang Yuguang

    2016-01-01

    Full Text Available Hydrocarbon fuel has been widely used in air-breathing scramjets and liquid rocket engines as coolant and propellant. However, possible heat transfer deterioration and threats from local high heat flux area in scramjet make heat transfer enhancement essential. In this work, 2-D steady numerical simulation was carried out to study different schemes of heat transfer enhancement based on a partially filled porous media in a tube. Both boundary and central layouts were analyzed and effects of gradient porous media were also compared. The results show that heat transfer in the transcritical area is enhanced at least 3 times with the current configuration compared to the clear tube. Besides, the proper use of gradient porous media also enhances the heat transfer compared to homogenous porous media, which could help to avoid possible over-temperature in the thermal protection.

  10. Comparison of Practical Investigations for CO Emissions Emitted From Single Cylinder S. I. Engine Fueled With Different Kinds of Hydrocarbon Fuels and Hydrogen

    Directory of Open Access Journals (Sweden)

    Khalil Ibrahim Abaas

    2011-01-01

    Full Text Available Liquefied petroleum gas (LPG, Natural gas (NG and hydrogen were all used to operate spark ignition internal combustion engine Ricardo E6. A comparison of CO emissions emitted from each case, with emissions emitted from engine fueled with gasoline as a fuel is conducted.The study was accomplished when engine operated at HUCR for gasoline n(8:1, was compared with its operation at HUCR for each fuel. Compression ratio, equivalence ratio and spark timing were studied at constant speed 1500 rpm.CO concentrations were little at lean ratios; it appeared to be effected a little with equivalence ratio in this side, at rich side its values became higher, and it appeared to be effected by equivalence ratio highly, the results showed that CO emissions resulted from gasoline engine were higher than that resulted from using LPG and NG all the time; while hydrogen engine emitted extremely low CO concentrations.

  11. Hydrocarbon Fuel Thermal Performance Modeling based on Systematic Measurement and Comprehensive Chromatographic Analysis

    Science.gov (United States)

    2016-07-27

    Conference Paper 3. DATES COVERED (From - To) 10 June 2016 - 27 July 2016 4. TITLE AND SUBTITLE Hydrocarbon Fuel Thermal Performance Modeling based on...The Johns Hopkins University Energetics Research Group (JHU/ERG), Columbia, MD and University of Washington, Seattle, WA 14. ABSTRACT Ensuring fuel ...is a common requirement for aircraft, rockets, and hypersonic vehicles. The Aerospace Fuels Quality Test and Model Development (AFQTMoDev) project

  12. MHD thrust vectoring of a rocket engine

    Science.gov (United States)

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

    2016-09-01

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

  13. Injector for liquid fueled rocket engine

    Science.gov (United States)

    Cornelius, Charles S. (Inventor); Myers, W. Neill (Inventor); Shadoan, Michael David (Inventor); Sparks, David L. (Inventor)

    2000-01-01

    An injector for liquid fueled rocket engines wherein a generally flat core having a frustoconical dome attached to one side of the core to serve as a manifold for a first liquid, with the core having a generally circular configuration having an axis. The other side of the core has a plurality of concentric annular first slots and a plurality of annular concentric second slots alternating with the first slots, the second slots having a greater depth than said first slots. A bore extends through the core for inletting a second liquid into said core, the bore intersecting the second slots to feed the second liquid into the second slots. The core also has a plurality of first passageways leading from the manifold to the first annular slots for feeding the first liquid into said first slots. A faceplate brazed to said other side of the core is provided with apertures extending from the first and second slots through said face plate, these apertures being positioned to direct fuel and liquid oxygen into contact with each other in the combustion chamber. The first liquid may be liquid oxygen and the second liquid may be kerosene or liquid hydrogen.

  14. Potential Climate and Ozone Impacts From Hybrid Rocket Engine Emissions

    Science.gov (United States)

    Ross, M.

    2009-12-01

    Hybrid rocket engines that use N2O as an oxidizer and a solid hydrocarbon (such as rubber) as a fuel are relatively new. Little is known about the composition of such hybrid engine emissions. General principles and visual inspection of hybrid plumes suggest significant soot and possibly NO emissions. Understanding hybrid rocket emissions is important because of the possibility that a fleet of hybrid powered suborbital rockets will be flying on the order of 1000 flights per year by 2020. The annual stratospheric emission for these rockets would be about 10 kilotons, equal to present day solid rocket motor (SRM) emissions. We present a preliminary analysis of the magnitude of (1) the radiative forcing from soot emissions and (2) the ozone depletion from soot and NO emissions associated with such a fleet of suborbital hybrid rockets. Because the details of the composition of hybrid emissions are unknown, it is not clear if the ozone depletion caused by these hybrid rockets would be more or less than the ozone depletion from SRMs. We also consider the climate implications associated with the N2O production and use requirements for hybrid rockets. Finally, we identify the most important data collection and modeling needs that are required to reliably assess the complete range of environmental impacts of a fleet of hybrid rockets.

  15. Safety Analysis of Liquid Rocket Engine Using Bayesian Networks

    Institute of Scientific and Technical Information of China (English)

    WANG Hua-wei; YAN Zhi-qiang

    2007-01-01

    Safety analysis for liquid rocket engine has a great meaning for shortening development cycle, saving development expenditure and reducing development risk. The relationship between the structure and component of liquid rocket engine is much more complex, furthermore test data are absent in development phase. Thereby, the uncertainties exist in safety analysis for liquid rocket engine. A safety analysis model integrated with FMEA(failure mode and effect analysis)based on Bayesian networks (BN) is brought forward for liquid rocket engine, which can combine qualitative analysis with quantitative decision. The method has the advantages of fusing multi-information, saving sample amount and having high veracity. An example shows that the method is efficient.

  16. Propellant Flow Actuated Piezoelectric Rocket Engine Igniter Project

    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,...

  17. Distributed Rocket Engine Testing Health Monitoring System Project

    Data.gov (United States)

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

  18. Distributed Rocket Engine Testing Health Monitoring System Project

    Data.gov (United States)

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

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

    Science.gov (United States)

    Ellis, David L.

    2016-01-01

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

  20. An experimental investigation of liquid methane convection and boiling in rocket engine cooling channels

    Science.gov (United States)

    Trujillo, Abraham Gerardo

    In the past decades, interest in developing hydrocarbon-fueled rocket engines for deep spaceflight missions has continued to grow. In particular, liquid methane (LCH4) has been of interest due to the weight efficiency, storage, and handling advantages it offers over several currently used propellants. Deep space exploration requires reusable, long life rocket engines. Due to the high temperatures reached during combustion, the life of an engine is significantly impacted by the cooling system's efficiency. Regenerative (regen) cooling is presented as a viable alternative to common cooling methods such as film and dump cooling since it provides improved engine efficiency. Due to limited availability of experimental sub-critical liquid methane cooling data for regen engine design, there has been an interest in studying the heat transfer characteristics of the propellant. For this reason, recent experimental studies at the Center for Space Exploration Technology Research (cSETR) at the University of Texas at El Paso (UTEP) have focused on investigating the heat transfer characteristics of sub-critical CH4 flowing through sub-scale cooling channels. To conduct the experiments, the csETR developed a High Heat Flux Test Facility (HHFTF) where all the channels are heated using a conduction-based thermal concentrator. In this study, two smooth channels with cross sectional geometries of 1.8 mm x 4.1 mm and 3.2 mm x 3.2 mm were tested. In addition, three roughened channels all with a 3.2 mm x 3.2 mm square cross section were also tested. For the rectangular smooth channel, Reynolds numbers ranged between 68,000 and 131,000, while the Nusselt numbers were between 40 and 325. For the rough channels, Reynolds numbers ranged from 82,000 to 131,000, and Nusselt numbers were between 65 and 810. Sub-cooled film-boiling phenomena were confirmed for all the channels presented in this work. Film-boiling onset at Critical Heat Flux (CHF) was correlated to a Boiling Number (Bo) of

  1. Yuzhnoye's new liquid rocket engines as enablers for space exploration

    Science.gov (United States)

    Degtyarev, Alexander; Kushnaryov, Alexander; Shulga, Vladimir; Ventskovsky, Oleg

    2016-10-01

    Advanced liquid rocket engines (LREs) are being created by Yuzhnoye Design Office of Ukraine based on the fifty-year experience of rocket engines' and propulsion systems' development. These LREs use both hypergolic (NTO+UDMH) and cryogenic (liquid oxygen+kerosene) propellants. First stage engines have a range of thrust from 40 to 250 t, while the upper stage (used in space) engines - from several kilograms to 50 t and a re-ignition feature. The engines are intended for both Ukraine"s independent access to space and international market.

  2. Oxidizer heat exchangers for rocket engine operation in idle modes

    Science.gov (United States)

    Kanic, P. G.; Kmiec, T. D.

    1987-01-01

    The heat exchanger concept is discussed together with its role in rocket engine operation in idle modes. Two heat exchanger designs (low and high heat transfer) utilizing different approaches to achieve stable oxygen vaporization are presented as well as their performance test results. It is concluded that compact and lightweight heat exchangers can be used in a stable manner under the 'idle' operating conditions expected with the RL10 rocket engine.

  3. Rocketdyne/Westinghouse nuclear thermal rocket engine modeling

    Science.gov (United States)

    Glass, James F.

    1993-01-01

    The topics are presented in viewgraph form and include the following: systems approach needed for nuclear thermal rocket (NTR) design optimization; generic NTR engine power balance codes; rocketdyne nuclear thermal system code; software capabilities; steady state model; NTR engine optimizer code-logic; reactor power calculation logic; sample multi-component configuration; NTR design code output; generic NTR code at Rocketdyne; Rocketdyne NTR model; and nuclear thermal rocket modeling directions.

  4. 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.

  5. Mixing and reaction processes in rocket based combined cycle and conventional rocket engines

    Science.gov (United States)

    Lehman, Matthew Kurt

    Raman spectroscopy was used to make species measurements in two rocket engines. An airbreathing rocket, the rocket based combined cycle (RBCC) engine, and a conventional rocket were investigated. A supersonic rocket plume mixing with subsonic coflowing air characterizes the ejector mode of the RBCC engine. The mixing length required for the air and plume to become homogenous is a critical dimension. For the conventional rocket experiments, a gaseous oxygen/gaseous hydrogen single-element shear coaxial injector was used. Three chamber Mach number conditions, 0.1, 0.2 and 0.3, were chosen to assess the effect of Mach number on mixing. The flow within the chamber was entirely subsonic. For the RBCC experiments, vertical Raman line measurements were made at multiple axial locations downstream from the rocket nozzle plane. Species profiles assessed the mixing progress between the supersonic plume and subsonic air. For the conventional rocket, Raman line measurements were made downstream from the injector face. The goal was to evaluate the effect of increased chamber Mach number on injector mixing/reaction. For both engines, quantitative and qualitative information was collected for computational fluid dynamics (CFD development. The RBCC experiments were conducted for three distinct geometries. The primary flow path was a diffuse and afterburner design with a direct-connect air supply. A sea-level static (SLS) version and a thermally choked variant were also tested. The experimental results show that mixing length increases with additional coflow air in the DAB geometry. Operation of variable rocket mixture ratios at identical air flow rates did not significantly affect the mixing length. The thermally choked variant had a longer mixing length compared to the DAB geometry, and the SLS modification had a shorter mixing length due to a reduced air flow. The conventional rocket studies focused on the effect of chamber Mach number on primary injector mixing. Chamber Mach

  6. National Gas Survey. Synthesized gaseous hydrocarbon fuels

    Energy Technology Data Exchange (ETDEWEB)

    None

    1978-06-01

    The supply-Technical Advisory Task Force-Synthesized Gaseous Hydrocarbon Fuels considered coal, hydrocarbon liquids, oil shales, tar sands, and bioconvertible materials as potential feedstocks for gaseous fuels. Current status of process technology for each feedstock was reviewed, economic evaluations including sensitivity analysis were made, and constraints for establishment of a synthesized gaseous hydrocarbon fuels industry considered. Process technology is presently available to manufacture gaseous hydrocarbon fuels from each of the feedstocks. In 1975 there were eleven liquid feedstock SNG plants in the United States having a capacity of 1.1 billion SCFD. There can be no contribution of SNG before 1982 from plants using feedstocks other than liquids because there are no plants in operation or under construction as of 1977. Costs for SNG are higher than current regulated prices for U.S. natural gas. Because of large reserves, coal is a prime feedstock candidate although there are major constraints in the area of coal leases, mining and water permits, and others. Commercial technology is available and several new gasification processes are under development. Oil shale is also a feedstock in large supply and commercial process technology is available. There are siting and permit constraints, and water availability may limit the ultimate size of an oil shale processing industry. Under projected conditions, bioconvertible materials are not expected to support the production of large quantities of pipeline quality gas during the next decade. Production of low or medium Btu gas from municipal solid wastes can be expected to be developed in urban areas in conjunction with savings in disposal costs. In the economic evaluations presented, the most significant factor for liquid feedstock plants is the anticipated cost of feedstock and fuel. The economic viability of plants using other feedstocks is primarily dependent upon capital requirements.

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

    Science.gov (United States)

    Thomas, Matt; Bossard, John; Early, Jim; Trinh, Huu; Dennis, Jay; Turner, James (Technical Monitor)

    2001-01-01

    This viewgraph presentation gives an overview of laser ignition technology for bipropellant rocket engines applications. The objectives of this project include: (1) the selection test chambers and flows; (2) definition of the laser ignition setup; (3) pulse format optimization; (4) fiber optic coupled laser ignition system analysis; and (5) chamber integration issues definition. The testing concludes that rocket combustion chamber laser ignition is imminent. Support technologies (multiplexing, window durability/cleaning, and fiber optic durability) are feasible.

  8. Waste Plastic Converting into Hydrocarbon Fuel Materials

    Energy Technology Data Exchange (ETDEWEB)

    Sarker, Moinuddin; Mamunor Rashid, Mohammad; Molla, Mohammad

    2010-09-15

    The increased demand and high prices for energy sources are driving efforts to convert organic compounds into useful hydrocarbon fuels. Although much of this work has focused on biomass, there are strong benefits to deriving fuels from waste plastic material. Natural State Research Inc. (NSR) has invented a simple and economically viable process to decompose the hydrocarbon polymers of waste plastic into the shorter chain hydrocarbon of liquid fuel (patent pending). The method and principle of the production / process will be discussed. Initial tests with several widely used polymers indicate a high potential for commercialization.

  9. 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.

  10. Investigations of Rocket Engine Combustion Emissions During ACCENT

    Science.gov (United States)

    Ross, M. N.; Friedl, R. R.

    2001-12-01

    The composition of rocket combustion emissions and the atmospheric processes that determine their stratospheric impacts are poorly understood. While present day rocket emissions do not significantly affect stratospheric chemistry, the potential for vigorous growth of the space transportation industry in coming decades suggests that rocket emissions and their stratospheric impacts should be better understood. A variety of in-situ measurements and modeling results were obtained during the Atmospheric Chemistry of Combustion Emissions Near the Tropopause (ACCENT) effort that will be used to evaluate the role of rocket exhaust in perturbing ozone chemistry in plume wakes and in the global stratosphere. We present a review of the ACCENT rocket emissions science objectives, summarize data obtained during the WB-57F plume wake sorties, and briefly discuss how the data will help resolve several outstanding questions regarding the impact of rocket emissions on the stratosphere. These include measurement of the emission indices for several important rocket engine combustion products and validation of plume wake chemistry models.

  11. 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.

  12. Investigation Status of Heat Exchange while Boiling Hydrocarbon Fuel

    Directory of Open Access Journals (Sweden)

    D. S. Obukhov

    2006-01-01

    Full Text Available The paper contains analysis of heat exchange investigations while boiling hydrocarbon fuel. The obtained data are within the limits of the S.S. Kutateladze dependence proposed in 1939. Heat exchange at non-stationary heat release has not been investigated. The data for hydrocarbon fuel with respect to critical density of heat flow are not available even for stationary conditions.

  13. Pyrochlore catalysts for hydrocarbon fuel reforming

    Science.gov (United States)

    Berry, David A.; Shekhawat, Dushyant; Haynes, Daniel; Smith, Mark; Spivey, James J.

    2012-08-14

    A method of catalytically reforming a reactant gas mixture using a pyrochlore catalyst material comprised of one or more pyrochlores having the composition A2B2-y-zB'yB"zO7-.DELTA., where y>0 and z.gtoreq.0. Distribution of catalytically active metals throughout the structure at the B site creates an active and well dispersed metal locked into place in the crystal structure. This greatly reduces the metal sintering that typically occurs on supported catalysts used in reforming reactions, and reduces deactivation by sulfur and carbon. Further, oxygen mobility may also be enhanced by elemental exchange of promoters at sites in the pyrochlore. The pyrochlore catalyst material may be utilized in catalytic reforming reactions for the conversion of hydrocarbon fuels into synthesis gas (H2+CO) for fuel cells, among other uses.

  14. Vacuum plasma spray applications on liquid fuel rocket engines

    Science.gov (United States)

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

    1992-01-01

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

  15. Additive Manufacturing a Liquid Hydrogen Rocket Engine

    Science.gov (United States)

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

    2016-01-01

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

  16. Near-term lunar nuclear thermal rocket engine options

    Science.gov (United States)

    Pelaccio, Dennis G.; Scheil, Christine M.; Collins, John T.

    1991-01-01

    The Nuclear Thermal Rocket (NTR) is an attractive candidate propulsion system option for manned planetary missions. Its high performance capability for such missions translates into a substantial reduction in low-earth-orbit (LEO) required mass and trip times with increased operational flexibility. This study examined NTR engine options that could support near-term lunar mission operations. Expander and gas generator cycle, solid-core NERVA derivative reactor-based NTR engines were investigated. Weight, size, operational characteristics, and design features for representative NTR engine concepts are presented. The impact of using these NTR engines for a typical lunar mission scenario is also examined.

  17. Hydrocarbon Rocket Technology Impact Forecasting

    Science.gov (United States)

    Stuber, Eric; Prasadh, Nishant; Edwards, Stephen; Mavris, Dimitri N.

    2012-01-01

    Forecasting method is a normative forecasting technique that allows the designer to quantify the effects of adding new technologies on a given design. This method can be used to assess and identify the necessary technological improvements needed to close the gap that exists between the current design and one that satisfies all constraints imposed on the design. The TIF methodology allows for more design knowledge to be brought to the earlier phases of the design process, making use of tools such as Quality Function Deployments, Morphological Matrices, Response Surface Methodology, and Monte Carlo Simulations.2 This increased knowledge allows for more informed decisions to be made earlier in the design process, resulting in shortened design cycle time. This paper will investigate applying the TIF method, which has been widely used in aircraft applications, to the conceptual design of a hydrocarbon rocket engine. In order to reinstate a manned presence in space, the U.S. must develop an affordable and sustainable launch capability. Hydrocarbon-fueled rockets have drawn interest from numerous major government and commercial entities because they offer a low-cost heavy-lift option that would allow for frequent launches1. However, the development of effective new hydrocarbon rockets would likely require new technologies in order to overcome certain design constraints. The use of advanced design methods, such as the TIF method, enables the designer to identify key areas in need of improvement, allowing one to dial in a proposed technology and assess its impact on the system. Through analyses such as this one, a conceptual design for a hydrocarbon-fueled vehicle that meets all imposed requirements can be achieved.

  18. Reusable rocket engine turbopump condition monitoring

    Science.gov (United States)

    Hampson, M. E.; Barkhoudarian, S.

    1985-01-01

    Significant improvements in engine readiness with attendant reductions in maintenance costs and turnaround times can be achieved with an engine condition monitoring system (CMS). The CMS provides real time health status of critical engine components, without disassembly, through component monitoring with advanced sensor technologies. Three technologies were selected to monitor the rotor bearings and turbine blades: the isotope wear detector and fiber optic deflectometer (bearings), and the fiber optic pyrometer (blades). Signal processing algorithms were evaluated and ranked for their utility in providing useful component health data to unskilled maintenance personnel. Design modifications to current configuration Space Shuttle Main Engine (SSME) high pressure turbopumps and the MK48-F turbopump were developed to incorporate the sensors.

  19. Developing Avionics Hardware and Software for Rocket Engine Testing

    Science.gov (United States)

    Aberg, Bryce Robert

    2014-01-01

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

  20. 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.

  1. Dual-fuel, dual-mode rocket engine

    Science.gov (United States)

    Martin, James A. (Inventor)

    1989-01-01

    The invention relates to a dual fuel, dual mode rocket engine designed to improve the performance of earth-to-orbit vehicles. For any vehicle that operates from the earth's surface to earth orbit, it is advantageous to use two different fuels during its ascent. A high density impulse fuel, such as kerosene, is most efficient during the first half of the trajectory. A high specific impulse fuel, such as hydrogen, is most efficient during the second half of the trajectory. The invention allows both fuels to be used with a single rocket engine. It does so by adding a minimum number of state-of-the-art components to baseline single made rocket engines, and is therefore relatively easy to develop for near term applications. The novelty of this invention resides in the mixing of fuels before exhaust nozzle cooling. This allows all of the engine fuel to cool the exhaust nozzle, and allows the ratio of fuels used throughout the flight depend solely on performance requirements, not cooling requirements.

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

    Science.gov (United States)

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

    1998-01-01

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

  3. Investigation of heat sink of endothermic hydrocarbon fuels

    Institute of Scientific and Technical Information of China (English)

    GUO Yong-sheng; LIN Rui-sen

    2005-01-01

    Endothermic hydrocarbon fuels are advanced coolants for high-temperature structures of spacecraft. No data of tested-cooling-ability of endothermic fuels have been broadly discussed in literature. In this work a high-temperature flow calorimeter was designed, and the cooling capacity of six different hydrocarbon fuels were measured. Experimental results showed that these hydrocarbon fuels have capacity for cooling high-temperature structures, and that the cooling capacity of fuel N-1 can reach 3.15 M J/kg, which can nearly satisfy the requirement of thermal management for a Mach 3 cruise aircraft, whose heat sink requirement is about 3.5 M J/kg. The endothermic velocity of hydrocarbon fuels was also measured by the calorimeter.

  4. Method of fabricating a rocket engine combustion chamber

    Science.gov (United States)

    Holmes, Richard R. (Inventor); Mckechnie, Timothy N. (Inventor); Power, Christopher A. (Inventor); Daniel, Ronald L., Jr. (Inventor); Saxelby, Robert M. (Inventor)

    1993-01-01

    A process for making a combustion chamber for a rocket engine wherein a copper alloy in particle form is injected into a stream of heated carrier gas in plasma form which is then projected onto the inner surface of a hollow metal jacket having the configuration of a rocket engine combustion chamber is described. The particles are in the plasma stream for a sufficient length of time to heat the particles to a temperature such that the particles will flatten and adhere to previously deposited particles but will not spatter or vaporize. After a layer is formed, cooling channels are cut in the layer, then the channels are filled with a temporary filler and another layer of particles is deposited.

  5. Syngas Upgrading to Hydrocarbon Fuels Technology Pathway

    Energy Technology Data Exchange (ETDEWEB)

    Talmadge, M.; Biddy, Mary J.; Dutta, Abhijit; Jones, Susanne B.; Meyer, Pimphan A.

    2013-03-31

    In support of the Bioenergy Technologies Office, the National Renewable Energy Laboratory (NREL) and the Pacific Northwest National Laboratory (PNNL) are undertaking studies of biomass conversion technologies to hydrocarbon fuels to identify barriers and target research toward reducing conversion costs. Process designs and preliminary economic estimates for each of these pathway cases were developed using rigorous modeling tools (Aspen Plus and Chemcad). These analyses incorporated the best information available at the time of development, including data from recent pilot and bench-scale demonstrations, collaborative industrial and academic partners, and published literature and patents. This pathway case investigates the upgrading of biomass derived synthesis gas (‘syngas’) to hydrocarbon biofuels. While this specific discussion focuses on the conversion of syngas via a methanol intermediate to hydrocarbon blendstocks, there are a number of alternative conversion routes for production of hydrocarbons through a wide array of intermediates from syngas. Future work will also consider the variations to this pathway to determine the most economically viable and risk adverse conversion route. Technical barriers and key research needs have been identified that should be pursued for the syngas to hydrocarbon pathway to be competitive with petroleum-derived gasoline, diesel and jet range blendstocks.

  6. A device for reforming a hydrocarbon fuel

    Energy Technology Data Exchange (ETDEWEB)

    Kendzi, T.; Ikuo, M.

    1984-03-15

    In order to utilize the heat from the reaction of reforming of a hydrocarbon fuel and the heat scattered from a heater, a design is proposed for a fuel reforming reactor in which the gases entering the reactor first pass inside the reactor along the external wall and are heated by the heat dispersed inside the reactor. Then they go in the opposite direction along a clearance between the interior shell of the reactor and the internal body of the reactor itself with a catalyst (Kt) and a heated electrical cylindrical heater. Then the gases, already heated, go directly into the cavity of the reactor filled with the catalyst where the reforming reaction occurs and then the gases and the vapors of the reformed fuel are discharged, passing through a system of heat exchangers. The layout of such a reactor, which contains a cylindrical shell inside, a cylindrical sleeve coaxial with it and the body of the reactor itself with the heater, is given. A system for attaching the internal sleeve and the body of the reactor to the catalyst is cited. The course of the gases inside the reactor is also given.

  7. Approaches to Low Fuel Regression Rate in Hybrid Rocket Engines

    OpenAIRE

    Dario Pastrone

    2012-01-01

    Hybrid rocket engines are promising propulsion systems which present appealing features such as safety, low cost, and environmental friendliness. On the other hand, certain issues hamper the development hoped for. The present paper discusses approaches addressing improvements to one of the most important among these issues: low fuel regression rate. To highlight the consequence of such an issue and to better understand the concepts proposed, fundamentals are summarized. Two approaches are pre...

  8. Analysis of rocket engine injection combustion processes

    Science.gov (United States)

    Salmon, J. W.

    1976-01-01

    A critique is given of the JANNAF sub-critical propellant injection/combustion process analysis computer models and application of the models to correlation of well documented hot fire engine data bases. These programs are the distributed energy release (DER) model for conventional liquid propellants injectors and the coaxial injection combustion model (CICM) for gaseous annulus/liquid core coaxial injectors. The critique identifies model inconsistencies while the computer analyses provide quantitative data on predictive accuracy. The program is comprised of three tasks: (1) computer program review and operations; (2) analysis and data correlations; and (3) documentation.

  9. Oxidation of Copper Alloy Candidates for Rocket Engine Applications

    Science.gov (United States)

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

    2002-01-01

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

  10. 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.

  11. An expert system for spectroscopic analysis of rocket engine plumes

    Science.gov (United States)

    Reese, Greg; Valenti, Elizabeth; Alphonso, Keith; Holladay, Wendy

    The expert system described in this paper analyzes spectral emissions of rocket engine exhaust plumes and shows major promise for use in engine health diagnostics. Plume emission spectroscopy is an important tool for diagnosing engine anomalies, but it is time-consuming and requires highly skilled personnel. The expert system was created to alleviate such problems. The system accepts a spectral plot in the form of wavelength vs intensity pairs and finds the emission peaks in the spectrum, lists the elemental emitters present in the data and deduces the emitter that produced each peak. The system consists of a conventional language component and a commercially available inference engine that runs on an Apple Macintosh computer. The expert system has undergone limited preliminary testing. It detects elements well and significantly decreases analysis time.

  12. A new comprehensive reaction mechanism for combustion of hydrocarbon fuels

    Energy Technology Data Exchange (ETDEWEB)

    Ranzi, E.; Sogaro, A.; Gaffuri, P.; Pennati, G. [Politecnico di Milano (Italy). Dipt. di Chimica Industriale e Ingegneria Chimica; Westbrook, C.K.; Pitz, W.J. [Lawrence Livermore National Lab., CA (United States)

    1993-12-03

    A chemical kinetic model has been developed which describes pyrolysis, ignition and oxidation of many small hydrocarbon fuels over a wide range of experimental conditions. Fuels include carbon monoxide and hydrogen, methane and other alkane species up to n-butane, ethylene, propene, acetylene, and oxygenated species such as methanol, acetaldehyde and ethanol. Formation of some larger intermediate and product species including benzene, butadiene, large olefins, and cyclopentadiene has been treated in a semi-empirical manner. The reaction mechanism has been tested for conditions that do not involve transport and diffusional processes, including plug flow and stirred reactors, batch reactors and shock tubes. The present kinetic model and its validation differ from previous reaction mechanisms in two ways. First, in addition to conventional combustion data, experiments more commonly associated with chemical engineering problems such as oxidative coupling, oxidative pyrolysis and steam cracking are used to test the reaction mechanism, making it even more general than previous models. In addition, H atom abstraction and some other reaction rates, even for the smaller C{sub 2}, C{sub 3} and C{sub 4} species, are treated using approximations that facilitate future extensions to larger fuels in a convenient manner. Construction of the reaction mechanism and comparisons with experimental data illustrate the generality of the model.

  13. Magnetic bearings: A key technology for advanced rocket engines?

    Science.gov (United States)

    Girault, J. PH.

    1992-01-01

    For several years, active magnetic bearings (AMB) have demonstrated their capabilities in many fields, from industrial compressors to control wheel suspension for spacecraft. Despite this broad area, no significant advance has been observed in rocket propulsion turbomachinery, where size, efficiency, and cost are crucial design criteria. To this respect, Societe Europeenne de Propulsion (SEP) had funded for several years significant efforts to delineate the advantages and drawbacks of AMB applied to rocket propulsion systems. Objectives of this work, relative technological basis, and improvements are described and illustrated by advanced turbopump layouts. Profiting from the advantages of compact design in cryogenic environments, the designs show considerable improvements in engine life, performances, and reliability. However, these conclusions should still be tempered by high recurrent costs, mainly due to the space-rated electronics. Development work focused on this point and evolution of electronics show the possibility to decrease production costs by an order of magnitude.

  14. Rocket

    Directory of Open Access Journals (Sweden)

    K. Karmarkar

    1952-09-01

    Full Text Available The rockets of World War II represented, not the invention of a new weapon, but the modernization of a very old one. As early as 1232 A.D, the Chinese launched rockets against the Mongols. About a hundred years later the knowledge of ledge of rockets was quite widespread and they were used to set fire to buildings and to terrorize the enemy. But as cannon developed, rockets declined in warfare. However rockets were used occasionally as weapons till about 1530 A.D. About this time improvements in artillery-rifled gun barrel and mechanism to absorb recoil-established a standard of efficiency with which rockets could not compare until World War II brought pew conditions

  15. 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.

  16. Axisymmetric Numerical Modeling of Pulse Detonation Rocket Engines

    Science.gov (United States)

    Morris, Christopher I.

    2005-01-01

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

  17. Net-Shape HIP Powder Metallurgy Components for Rocket Engines

    Science.gov (United States)

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

    2005-01-01

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

  18. Multiple dopant injection system for small rocket engines

    Science.gov (United States)

    Sakala, G. G.; Raines, N. G.

    1992-07-01

    The Diagnostics Test Facility (DTF) at NASA's Stennis Space Center (SSC) was designed and built to provide a standard rocket engine exhaust plume for use in the research and development of engine health monitoring instrumentation. A 1000 lb thrust class liquid oxygen (LOX)-gaseous hydrogen (GH2) fueled rocket engine is used as the subscale plume source to simulate the SSME during experimentation and instrument development. The ability of the DTF to provide efficient, and low cost test operations makes it uniquely suited for plume diagnostic experimentation. The most unique feature of the DTF is the Multiple Dopant Injection System (MDIS) that is used to seed the exhaust plume with the desired element or metal alloy. The dopant injection takes place at the fuel injector, yielding a very uniform and homogeneous distribution of the seeding material in the exhaust plume. The MDIS allows during a single test firing of the DTF, the seeding of the exhaust plume with up to three different dopants and also provides distilled water base lines between the dopants. A number of plume diagnostic-related experiments have already utilized the unique capabilities of the DTF.

  19. Using Innovative Technologies for Manufacturing and Evaluating Rocket Engine Hardware

    Science.gov (United States)

    Betts, Erin M.; Hardin, Andy

    2011-01-01

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

  20. Evaluation of Vortex Chamber Concepts for Liquid Rocket Engine Applications

    Science.gov (United States)

    Trinh, Huu Phuoc; Knuth, Williams; Michaels, Scott; Turner, James E. (Technical Monitor)

    2000-01-01

    Rocket-based combined-cycle engines (RBBC) being considered at NASA for future generation launch vehicles feature clusters of small rocket thrusters as part of the engine components. Depending on specific RBBC concepts, these thrusters may be operated at various operating conditions including power level and/or propellant mixture ratio variations. To pursue technology developments for future launch vehicles, NASA/Marshall Space Flight Center (MSFC) is examining vortex chamber concepts for the subject cycle engine application. Past studies indicated that the vortex chamber schemes potentially have a number of advantages over conventional chamber methods. Due to the nature of the vortex flow, relatively cooler propellant streams tend to flow along the chamber wall. Hence, the thruster chamber can be operated without the need of any cooling techniques. This vortex flow also creates strong turbulence, which promotes the propellant mixing process. Consequently, the subject chamber concepts not only offer the system simplicity but they also would enhance the combustion performance. The test results showed that the chamber performance was markedly high even at a low chamber length-to- diameter ratio (L/D). This incentive can be translated to a convenience in the thrust chamber packaging.

  1. Effects of rocket engines on laser during lunar landing

    Energy Technology Data Exchange (ETDEWEB)

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

    2013-11-15

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

  2. On the hydrodynamics of rocket propellant engine inducers and turbopumps

    Science.gov (United States)

    d'Agostino, L.

    2013-12-01

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

  3. Distributed Health Monitoring System for Reusable Liquid Rocket Engines

    Science.gov (United States)

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

    2009-01-01

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

  4. Russian Nuclear Rocket Engine Design for Mars Exploration

    Institute of Scientific and Technical Information of China (English)

    Vadim Zakirov; Vladimir Pavshook

    2007-01-01

    This paper is to promote investigation into the nuclear rocket engine (NRE) propulsion option that is considered as a key technology for manned Mars exploration. Russian NRE developed since the 1950 s in the former Soviet Union to a full-scale prototype by the 1990 s is viewed as advantageous and the most suitable starting point concept for manned Mars mission application study. The main features of Russian heterogeneous core NRE design are described and the most valuable experimental performance results are summarized. These results have demonstrated the significant specific impulse performance advantage of the NRE over conventional liquid rocket engine (LRE) propulsion technologies. Based on past experience,the recent developments in the field of high-temperature nuclear fuels, and the latest conceptual studies, the developed NRE concept is suggested to be upgraded to the nuclear power and propulsion system (NPPS),more suitable for future manned Mars missions. Although the NRE still needs development for space application, the problems are solvable with additional effort and funding.

  5. Hyperthermal Environments Simulator for Nuclear Rocket Engine Development

    Science.gov (United States)

    Litchford, Ron J.; Foote, John P.; Clifton, W. B.; Hickman, Robert R.; Wang, Ten-See; Dobson, Christopher C.

    2011-01-01

    An arc-heater driven hyperthermal convective environments simulator was recently developed and commissioned for long duration hot hydrogen exposure of nuclear thermal rocket materials. This newly established non-nuclear testing capability uses a high-power, multi-gas, wall-stabilized constricted arc-heater to produce hightemperature pressurized hydrogen flows representative of nuclear reactor core environments, excepting radiation effects, and is intended to serve as a low-cost facility for supporting non-nuclear developmental testing of hightemperature fissile fuels and structural materials. The resulting reactor environments simulator represents a valuable addition to the available inventory of non-nuclear test facilities and is uniquely capable of investigating and characterizing candidate fuel/structural materials, improving associated processing/fabrication techniques, and simulating reactor thermal hydraulics. This paper summarizes facility design and engineering development efforts and reports baseline operational characteristics as determined from a series of performance mapping and long duration capability demonstration tests. Potential follow-on developmental strategies are also suggested in view of the technical and policy challenges ahead. Keywords: Nuclear Rocket Engine, Reactor Environments, Non-Nuclear Testing, Fissile Fuel Development.

  6. 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

  7. Linear quadratic servo control of a reusable rocket engine

    Science.gov (United States)

    Musgrave, Jeffrey L.

    1991-01-01

    The paper deals with the development of a design method for a servo component in the frequency domain using singular values and its application to a reusable rocket engine. A general methodology used to design a class of linear multivariable controllers for intelligent control systems is presented. Focus is placed on performance and robustness characteristics, and an estimator design performed in the framework of the Kalman-filter formalism with emphasis on using a sensor set different from the commanded values is discussed. It is noted that loop transfer recovery modifies the nominal plant noise intensities in order to obtain the desired degree of robustness to uncertainty reflected at the plant input. Simulation results demonstrating the performance of the linear design on a nonlinear engine model over all power levels during mainstage operation are discussed.

  8. Software for Estimating Costs of Testing Rocket Engines

    Science.gov (United States)

    Hines, Merlon M.

    2004-01-01

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

  9. Using Innovative Technologies for Manufacturing Rocket Engine Hardware

    Science.gov (United States)

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

    2011-01-01

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

  10. Status report on a real time Engine Diagnostics Console for rocket engine exhaust plume monitoring

    Science.gov (United States)

    Bircher, F. E.; Gardner, D. G.; Vandyke, D. B.; Harris, A. B.; Chenevert, D. J.

    1990-01-01

    This paper describes the work done on the Engine Diagnostics Console during the past year of development at Stennis Space Center. The Engine Diagnostics Console (EDC) is a hardware and software package which provides near real time monitoring of rocket engine exhaust plume emissions during ground testing. The long range goal of the EDC development program is to develop an instrument that can detect engine degradation leading to catastrophic failure, and respond by taking preventative measures. The immediate goal for the past year's effort is the ability to process spectral data, taken from a rocket engine's exhaust plume, and to identify in an automated and high speed manner, the elemental species and multielemental materials that are present in the exhaust plume.

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

    Science.gov (United States)

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

    2004-10-01

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

  12. Using Innovative Techniques for Manufacturing Rocket Engine Hardware

    Science.gov (United States)

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

    2011-01-01

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

  13. High energy-density liquid rocket fuel performance

    Science.gov (United States)

    Rapp, Douglas C.

    1990-01-01

    A fuel performance database of liquid hydrocarbons and aluminum-hydrocarbon fuels was compiled using engine parametrics from the Space Transportation Engine Program as a baseline. Propellant performance parameters are introduced. General hydrocarbon fuel performance trends are discussed with respect to hydrogen-to-carbon ratio and heat of formation. Aluminum-hydrocarbon fuel performance is discussed with respect to aluminum metal loading. Hydrocarbon and aluminum-hydrocarbon fuel performance is presented with respect to fuel density, specific impulse and propellant density specific impulse.

  14. Software for Preprocessing Data from Rocket-Engine Tests

    Science.gov (United States)

    Cheng, Chiu-Fu

    2004-01-01

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

  15. Romanian MRE Rocket Engines Program - An Early Endeavor

    Science.gov (United States)

    Rugescu, R. E.

    2002-01-01

    (MRE) was initiated in the years '60 of the past century at the Chair of Aerospace Sciences "Elie Carafoli" from the "Politehnica" University in Bucharest (PUB). Consisting of theoretical and experimental investigations in the form of computational methods and technological solutions for small size MRE-s and the concept of the test stand for these engines, the program ended in the construction of the first Romanian liquid rocket motors. Hermann Oberth and Dorin Pavel, were known from 1923, no experimental practice was yet tempted, at the time level of 1960. It was the intention of the developers at PUB to cover this gap and initiate a feasible, low-cost, demonstrative program of designing and testing experimental models of MRE. The research program was oriented towards future development of small size space carrier vehicles for scientific applications only, as an independent program with no connection to other defense programs imagined by the authorities in Bucharest, at that time. Consequently the entire financial support was assured by "Politehnica" university. computerized methods in the thermochemistry of heterogeneous combustion, for both steady and unsteady flows with chemical reactions and two phase flows. The research was gradually extended to the production of a professional CAD program for steady-state heat transfer simulations and the loading capacity analyses of the double wall, cooled thrust chamber. The resulting computer codes were run on a 360-30 IMB machine, beginning in 1968. Some of the computational methods were first exposed at the 9th International Conference on Applied Mechanics, held in Bucharest between June 23-27, 1969. hot testing of a series of storable propellant, variable thrust, variable geometry, liquid rocket motors, with a maximal thrust of 200N. A remotely controlled, portable test bad, actuated either automatically or manually and consisting of a 6-modules construction was built for this motor series, with a simple 8 analog

  16. Analysis of hydrocarbon fuel properties by means of Raman spectroscopy

    Science.gov (United States)

    Flatley, Martin W.

    The project is focused on the determination of Raman spectra of hydrocarbon fuel samples using a spectrometer employing a silicon linear array detector which has a spectral range of 400 nm to 1.1 mum. The spectra are processed using chemometric techniques in order to determine the concentrations of the tracked blend components and analytical values that are used to ensure that desired specifications are achieved. The verification is based on the American Standard Testing Methods procedures for the determination of the motor, research, and road octane numbers, simulated distillation and Reid vapour pressure. Blending is one of the most important steps in the final production of hydrocarbon fuels; as many as ten complex components are mixed to achieve the desired properties of the final product. Traditionally, blending relies on well-established analytical methods such as gas chromatography for component and simulated distillation analysis, knock engines and near infrared spectroscopy for octane analysis. All of these methods are reliable and accurate, but their results are not available in real time but rather with a substantial delay, since it is in the nature of the methods that the sample must be transported from a test site to the site where the instrument is located. Additional time is required for performing the analytical procedure; e.g. the results of a gas chromatography analysis are only available from minutes to hours after the sample has been introduced into the instrument. Consequently, the results, although accurate, become only available after the process of blending has been completed. The thesis describes an implementation of a Raman spectroscopic method, which is novel in the given context, since it allows monitoring and control of the blending process online, in real time. A Raman spectrometer was designed, using a solid state laser for excitation (785 nm, 800 mW), a blazed grating for the diffraction (600 lines-per-millimeter, 750 nm blaze, 635

  17. The open-cycle gas-core nuclear rocket engine - Some engineering considerations.

    Science.gov (United States)

    Taylor, M. F.; Whitmarsh, C. L., Jr.; Sirocky, P. J., Jr.; Iwanczyk, L. C.

    1971-01-01

    A preliminary design study of a conceptual 6000-MW open-cycle gas-core nuclear rocket engine system was made. The engine has a thrust of 44,200 lb and a specific impulse of 4400 sec. The nuclear fuel is uranium-235 and the propellant is hydrogen. Critical fuel mass was calculated for several reactor configurations. Major components of the reactor (reflector, pressure vessel) and the waste heat rejection system were considered conceptually and were sized.

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

    Science.gov (United States)

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

    2010-01-01

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

  19. Approaches to Low Fuel Regression Rate in Hybrid Rocket Engines

    Directory of Open Access Journals (Sweden)

    Dario Pastrone

    2012-01-01

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

  20. Low loss injector for liquid propellant rocket engines

    Science.gov (United States)

    Vonpragenau, G. L. (Inventor)

    1986-01-01

    A low pressure loss injector element is disclosed for the main combustion chamber of a rocket engine which includes a lox post terminating in a cylindrical barrel. Received within the barrel is a lox plug which is threaded in the lox post and includes an interchangeable lox metering sieve which meters the lox into an annular lox passage. A second annular gas passage is coaxial with the annular lox passage. A cylindrical sleeve surrounds the annular gas passage and includes an interchangeable gas metering seive having metering orifices through which a hot gas passes into the annular passage. The jets which emerge from the annular lox passage and annular gas passage intersect in a recessed area away from the combustion area. Thus, mixing and combustion stability are enhanced.

  1. Materials for advanced rocket engine turbopump turbine blades

    Science.gov (United States)

    Chandler, W. T.

    1985-01-01

    A study program was conducted to identify those materials that will provide the greatest benefits as turbine blades for advanced liquid propellant rocket engine turbines and to prepare technology plans for the development of those materials for use in the 1990 through 1995 period. The candidate materials were selected from six classes of materials: single-crystal (SC) superalloys, oxide dispersion-strengthened (ODS) superalloys, rapid solidification processed (RSP) superalloys, directionally solidified eutectic (DSE) superalloys, fiber-reinforced superalloy (FRS) composites, and ceramics. Properties of materials from the six classes were compiled and evaluated and property improvements were projected approximately 5 years into the future for advanced versions of materials in each of the six classes.

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

    CERN Document Server

    Zhang, Wei

    2016-01-01

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

  3. A Design Tool for Liquid Rocket Engine Injectors

    Science.gov (United States)

    Farmer, R.; Cheng, G.; Trinh, H.; Tucker, K.

    2000-01-01

    A practical design tool which emphasizes the analysis of flowfields near the injector face of liquid rocket engines has been developed and used to simulate preliminary configurations of NASA's Fastrac and vortex engines. This computational design tool is sufficiently detailed to predict the interactive effects of injector element impingement angles and points and the momenta of the individual orifice flows and the combusting flow which results. In order to simulate a significant number of individual orifices, a homogeneous computational fluid dynamics model was developed. To describe sub- and supercritical liquid and vapor flows, the model utilized thermal and caloric equations of state which were valid over a wide range of pressures and temperatures. The model was constructed such that the local quality of the flow was determined directly. Since both the Fastrac and vortex engines utilize RP-1/LOX propellants, a simplified hydrocarbon combustion model was devised in order to accomplish three-dimensional, multiphase flow simulations. Such a model does not identify drops or their distribution, but it does allow the recirculating flow along the injector face and into the acoustic cavity and the film coolant flow to be accurately predicted.

  4. Evaluation of Impinging Stream Vortex Chamber Concepts for Liquid Rocket Engine Applications

    Science.gov (United States)

    Trinh, Huu P.; Bullard, Brad; Kopicz, Charles; Michaels, Scott

    2002-01-01

    To pursue technology developments for future launch vehicles, NASA/Marshall Space Flight Center (MSFC) is examining vortex chamber concepts for liquid rocket engine applications. Past studies indicated that the vortex chamber schemes potentially have a number of advantages over conventional chamber methods. Due to the nature of the vortex flow, relatively cooler propellant streams tend to flow along the chamber wall. Hence, the thruster chamber can be operated without the need of any cooling techniques. This vortex flow also creates strong turbulence, which promotes the propellant mixing process. Consequently, the subject chamber concepts not only offer system simplicity, but also enhance the combustion performance. Test results have shown that chamber performance is markedly high even at a low chamber length-to-diameter ratio (LD). This incentive can be translated to a convenience in the thrust chamber packaging. Variations of the vortex chamber concepts have been introduced in the past few decades. These investigations include an ongoing work at Orbital Technologies Corporation (ORBITEC). By injecting the oxidizer tangentially at the chamber convergence and fuel axially at the chamber head end, Knuth et al. were able to keep the wall relatively cold. A recent investigation of the low L/D vortex chamber concept for gel propellants was conducted by Michaels. He used both triplet (two oxidizer orifices and one fuel orifice) and unlike impinging schemes to inject propellants tangentially along the chamber wall. Michaels called the subject injection scheme an Impinging Stream Vortex Chamber (ISVC). His preliminary tests showed that high performance, with an Isp efficiency of 9295, can be obtained. MSFC and the U. S. Army are jointly investigating an application of the ISVC concept for the cryogenic oxygen/hydrocarbon propellant system. This vortex chamber concept is currently tested with gel propellants at AMCOM at Redstone Arsenal, Alabama. A version of this concept

  5. Performance Increase Verification for a Bipropellant Rocket Engine

    Science.gov (United States)

    Alexander, Leslie; Chapman, Jack; Wilson, Reed; Krismer, David; Lu, Frank; Wilson, Kim; Miller, Scott; England, Chris

    2008-01-01

    Component performance assessment testing for a, pressure-fed earth storable bipropellant rocket engine was successfully completed at Aerojet's Redmond test facility. The primary goal of the this development project is to increase the specific impulse of an apogee class bi-propellant engine to greater than 330 seconds with nitrogen tetroxide and monomethylhydrazine propellants and greater than 335 seconds with nitrogen tetroxide and hydrazine. The secondary goal of the project is to take greater advantage of the high temperature capabilities of iridium/rhenium chambers. In order to achieve these goals, the propellant feed pressures were increased to 400 psia, nominal, which in turn increased the chamber pressure and temperature, allowing for higher c*. The tests article used a 24-on-24 unlike doublet injector design coupled with a copper heat sink chamber to simulate a flight configuration combustion chamber. The injector is designed to produce a nominal 200 lbf of thrust with a specific impulse of 335 seconds (using hydrazine fuel). Effect of Chamber length on engine C* performance was evaluated with the use of modular, bolt-together test hardware and removable chamber inserts. Multiple short duration firings were performed to characterize injector performance across a range of thrust levels, 180 to 220 lbf, and mixture ratios, from 1.1 to 1.3. During firing, ignition transient, chamber pressure, and various temperatures were measured in order to evaluate the performance of the engine and characterize the thermal conditions. The tests successfully demonstrated the stable operation and performance potential of a full scale engine with a measured c* of XXXX ft/sec (XXXX m/s) under nominal operational conditions.

  6. Real-time diagnostics for a reusable rocket engine

    Science.gov (United States)

    Guo, T. H.; Merrill, W.; Duyar, A.

    1992-01-01

    A hierarchical, decentralized diagnostic system is proposed for the Real-Time Diagnostic System component of the Intelligent Control System (ICS) for reusable rocket engines. The proposed diagnostic system has three layers of information processing: condition monitoring, fault mode detection, and expert system diagnostics. The condition monitoring layer is the first level of signal processing. Here, important features of the sensor data are extracted. These processed data are then used by the higher level fault mode detection layer to do preliminary diagnosis on potential faults at the component level. Because of the closely coupled nature of the rocket engine propulsion system components, it is expected that a given engine condition may trigger more than one fault mode detector. Expert knowledge is needed to resolve the conflicting reports from the various failure mode detectors. This is the function of the diagnostic expert layer. Here, the heuristic nature of this decision process makes it desirable to use an expert system approach. Implementation of the real-time diagnostic system described above requires a wide spectrum of information processing capability. Generally, in the condition monitoring layer, fast data processing is often needed for feature extraction and signal conditioning. This is usually followed by some detection logic to determine the selected faults on the component level. Three different techniques are used to attack different fault detection problems in the NASA LeRC ICS testbed simulation. The first technique employed is the neural network application for real-time sensor validation which includes failure detection, isolation, and accommodation. The second approach demonstrated is the model-based fault diagnosis system using on-line parameter identification. Besides these model based diagnostic schemes, there are still many failure modes which need to be diagnosed by the heuristic expert knowledge. The heuristic expert knowledge is

  7. Heat transfer in rocket engine combustion chambers and nozzles

    Science.gov (United States)

    Anderson, P. G.; Cheng, G. C.; Farmer, R. C.

    1993-01-01

    Complexities of liquid rocket engine heat transfer which involve the injector faceplate and regeneratively and film cooled walls are being investigated by computational analysis. A conjugate heat transfer analysis will be used to describe localized heating phenomena associated with particular injector configurations and coolant channels and film coolant dumps. These components are being analyzed, and the analyses verified with appropriate test data. Finally, the component analyses will be synthesized into an overall flowfield/heat transfer model. The FDNS code is being used to make the component analyses. Particular attention is being given to the representation of the thermodynamic properties of the fluid streams and to the method of combining the detailed models to represent overall heating. Unit flow models of specific coaxial injector elements have been developed and will be described. Since test data from the NLS development program are not available, new validation heat transfer data have been sought. Suitable data were obtained from a Rocketdyne test program on a model hydrocarbon/oxygen engine. Simulations of these test data will be presented. Recent interest in the hybrid motor have established the need for analyses of ablating solid fuels in the combustion chamber. Analysis of a simplified hybrid motor will also be presented.

  8. Plume spectrometry for liquid rocket engine health monitoring

    Science.gov (United States)

    Powers, William T.; Sherrell, F. G.; Bridges, J. H., III; Bratcher, T. W.

    1988-01-01

    An investigation of Space Shuttle Main Engine (SSME) testing failures identified optical events which appeared to be precursors of those failures. A program was therefore undertaken to detect plume trace phenomena characteristic of the engine and to design a monitoring system, responsive to excessive activity in the plume, capable of delivering a warning of an anomalous condition. By sensing the amount of extraneous material entrained in the plume and considering engine history, it may be possible to identify wearing of failing components in time for a safe shutdown and thus prevent a catastrophic event. To investigate the possibilities of safe shutdown and thus prevent a monitor to initiate the shutdown procedure, a large amount of plume data were taken from SSME firings using laboratory instrumentation. Those data were used to design a more specialized instrument dedicated to rocket plume diagnostics. The spectral wavelength range of the baseline data was about 220 nanometers (nm) to 15 micrometer with special attention given to visible and near UV. The data indicates that a satisfactory design will include a polychromator covering the range of 250 nM to 1000 nM, along with a continuous coverage spectrometer, each having a resolution of at least 5A degrees. The concurrent requirements for high resolution and broad coverage are normally at odds with one another in commercial instruments, therefore necessitating the development of special instrumentation. The design of a polychromator is reviewed herein, with a detailed discussion of the continuous coverage spectrometer delayed to a later forum. The program also requires the development of applications software providing detection, variable background discrimination, noise reduction, filtering, and decision making based on varying historical data.

  9. 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.

  10. Studies of Fission Fragment Rocket Engine Propelled Spacecraft

    Science.gov (United States)

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

    2014-01-01

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

  11. Hyperthermal Environments Simulator for Nuclear Rocket Engine Development

    Science.gov (United States)

    Litchford, R. J.; Foote, J. P.; Clifton, W. B.; Hickman, R. R.; Wang, T.-S.; Dobson, C. C.

    An arc-heater driven hyperthermal convective environments simulator was recently developed and commissioned for long duration hot hydrogen exposure of nuclear thermal rocket materials. This newly established non-nuclear testing capability uses a high-power, multi-gas, wall-stabilised constricted arc-heater to produce high-temperature pressurised hydrogen flows representative of nuclear reactor core environments, excepting radiation effects, and is intended to serve as a low-cost facility for supporting non-nuclear developmental testing of high-temperature fissile fuels and structural materials. The resulting reactor environments simulator represents a valuable addition to the available inventory of non-nuclear test facilities and is uniquely capable of investigating and characterising candidate fuel/structural materials, improving associated processing/ fabrication techniques, and simulating reactor thermal hydraulics. This paper summarizes facility design and engineering development efforts and reports baseline operational characteristics as determined from a series of performance mapping and long duration capability demonstration tests. Potential follow-on developmental strategies are also suggested in view of the technical and policy challenges ahead.

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

    Science.gov (United States)

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

    2010-09-01

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

  13. Regeneratively-Cooled, Turbopump-Fed, Small-Scale Cryogenic Rocket Engines Project

    Data.gov (United States)

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

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

    Data.gov (United States)

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

  15. 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.

  16. Magnetohydrodynamic Augmentation of Pulse Detonation Rocket Engines (Preprint)

    Science.gov (United States)

    2010-09-28

    Detonation Rocket-Induced MHD Ejector (PDRIME) concept, energy could be extracted from the high speed portion of the system, e.g., through an MHD...but with some challenges associated with achieving these gains, suggesting further analysis and optimization are required. 15. SUBJECT TERMS 16...mentation, such as in the Pulse Detonation Rocket-Induced MHD Ejector (PDRIME) concept, energy could be extracted from the high speed por- tion of the system

  17. Monomethylhydrazine versus hydrazine fuels - Test results using a 100 pound thrust bipropellant rocket engine

    Science.gov (United States)

    Smith, J. A.; Stechman, R. C.

    1981-01-01

    A test program was performed to evaluate hydrazine (N2H4) as a fuel for a 445 Newton (100 lbf) thrust bipropellant rocket engine. Results of testing with an identical thruster utilizing monomethylhydrazine (MMH) are included for comparison. Engine performance with hydrazine fuel was essentially identical to that experienced with monomethylhydrazine although higher combustor wall temperatures (approximately 400 F) were obtained with hydrazine. Results are presented which indicate that hydrazine as a fuel is compatible with Marquardt bipropellant rocket engines which use monomethylhydrazine as a baseline fuel.

  18. Catalytic co-pyrolysis of waste vegetable oil and high density polyethylene for hydrocarbon fuel production.

    Science.gov (United States)

    Wang, Yunpu; Dai, Leilei; Fan, Liangliang; Cao, Leipeng; Zhou, Yue; Zhao, Yunfeng; Liu, Yuhuan; Ruan, Roger

    2017-03-01

    In this study, a ZrO2-based polycrystalline ceramic foam catalyst was prepared and used in catalytic co-pyrolysis of waste vegetable oil and high density polyethylene (HDPE) for hydrocarbon fuel production. The effects of pyrolysis temperature, catalyst dosage, and HDPE to waste vegetable oil ratio on the product distribution and hydrocarbon fuel composition were examined. Experimental results indicate that the maximum hydrocarbon fuel yield of 63.1wt. % was obtained at 430°C, and the oxygenates were rarely detected in the hydrocarbon fuel. The hydrocarbon fuel yield increased when the catalyst was used. At the catalyst dosage of 15wt.%, the proportion of alkanes in the hydrocarbon fuel reached 97.85wt.%, which greatly simplified the fuel composition and improved the fuel quality. With the augment of HDPE to waste vegetable oil ratio, the hydrocarbon fuel yield monotonously increased. At the HDPE to waste vegetable oil ratio of 1:1, the maximum proportion (97.85wt.%) of alkanes was obtained. Moreover, the properties of hydrocarbon fuel were superior to biodiesel and 0(#) diesel due to higher calorific value, better low-temperature low fluidity, and lower density and viscosity. Copyright © 2017 Elsevier Ltd. All rights reserved.

  19. Technique for the optimization of the powerhead configuration and performance of liquid rocket engines

    Science.gov (United States)

    St. Germain, Brad David

    The development and optimization of liquid rocket engines is an integral part of space vehicle design, since most Earth-to-orbit launch vehicles to date have used liquid rockets as their main propulsion system. Rocket engine design tools range in fidelity from very simple conceptual level tools to full computational fluid dynamics (CFD) simulations. The level of fidelity of interest in this research is a design tool that determines engine thrust and specific impulse as well as models the powerhead of the engine. This is the highest level of fidelity applicable to a conceptual level design environment where faster running analyses are desired. The optimization of liquid rocket engines using a powerhead analysis tool is a difficult problem, because it involves both continuous and discrete inputs as well as a nonlinear design space. Example continuous inputs are the main combustion chamber pressure, nozzle area ratio, engine mixture ratio, and desired thrust. Example discrete variable inputs are the engine cycle (staged-combustion, gas generator, etc.), fuel/oxidizer combination, and engine material choices. Nonlinear optimization problems involving both continuous and discrete inputs are referred to as Mixed-Integer Nonlinear Programming (MINLP) problems. Many methods exist in literature for solving MINLP problems; however none are applicable for this research. All of the existing MINLP methods require the relaxation of the discrete variables as part of their analysis procedure. This means that the discrete choices must be evaluated at non-discrete values. This is not possible with an engine powerhead design code. Therefore, a new optimization method was developed that uses modified response surface equations to provide lower bounds of the continuous design space for each unique discrete variable combination. These lower bounds are then used to efficiently solve the optimization problem. The new optimization procedure was used to find optimal rocket engine designs

  20. NUMERICAL STUDIES ON HYDROGEN COMBUSTION IN A FILM COOLED CRYOGENIC ROCKET ENGINE

    Directory of Open Access Journals (Sweden)

    ARSHAD A.

    2012-07-01

    Full Text Available Liquid rocket engines have variety of propellant combinations which produces very high specific impulses. It is due to this fact; very high heat fluxes are incident on the combustion chamber and the nozzle walls. In order to deal with these heat fluxes, a wide range of cooling techniques have been employed, out of which a combination of film cooling and regenerative cooling promises to be the most effective one. The present study involves the numerical analysis of combustion in a typical film cooled cryogenic rocket engine thrust chamber considering the combustion of the fuel, heat transfer through the chamber walls and the fluid flow simultaneously. Analysis was done for a typical rocket engine thrust chamber with a single coaxial injector which uses gaseous hydrogen as the fuel and liquid oxygen as the oxidizer.

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

    Science.gov (United States)

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

    2016-12-01

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

  2. Optimisation Study of a Homogeneously-Catalysed HTP Rocket Engine

    Science.gov (United States)

    Musker, A.; Roberts, G.; Chandler, P.; Grayson, J.; Holdsworth, J.

    2004-10-01

    The decomposition of hydrogen peroxide (HTP) using a liquid catalyst offers an alternative to the employment of traditional catalyst packs. A test rig has been used to estimate the length of chamber required for complete decomposition to take place for the case of HTP with a phosphorus content some 10000 times the level normally associated with rocket-grade HTP. Complete decomposition within a 10 mm advection length was achieved.

  3. Storable Hypergolic Solid Fuel for Hybrid Rocket Engines

    Directory of Open Access Journals (Sweden)

    R. V. Singh

    1976-07-01

    Full Text Available A solid fuel was synthesised by condensing aniline with furfuraldehyde. The product was directly cast in the rocket motor casing. After curing a hard solid mass was obtained. This was found to have good hypergolicity with RFNA (Red Fuming Nitric Acid, good storability at room temperature and the mechanical properties. The paper presented the techniques of casting, ignition delay measurements and indicates the future programme for this study.

  4. SMC Standard: Evaluation and Test Requirements for Liquid Rocket Engines

    Science.gov (United States)

    2017-07-26

    Company, 2002. 23. NASA SP-8123, Liquid Rocket Lines, Bellows, Flexible Hoses, and Filters, National Aeronautics and Space Administration, April 1977...boundaries established by design requirements. For example, the test screens for malfunctions, failure to execute, sequence of action, interruption...Test: A static load or pressure test performed as an acceptance workmanship screen to prove the structural integrity of a unit or assembly. Gives

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

    Science.gov (United States)

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

    2013-01-01

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

  6. Pyrochlore-type catalysts for the reforming of hydrocarbon fuels

    Science.gov (United States)

    Berry, David A [Morgantown, WV; Shekhawat, Dushyant [Morgantown, WV; Haynes, Daniel [Morgantown, WV; Smith, Mark [Morgantown, WV; Spivey, James J [Baton Rouge, LA

    2012-03-13

    A method of catalytically reforming a reactant gas mixture using a pyrochlore catalyst material comprised of one or more pyrochlores having the composition A.sub.2-w-xA'.sub.wA''.sub.xB.sub.2-y-zB'.sub.yB''.sub.zO.sub.7-.DELTA.. Distribution of catalytically active metals throughout the structure at the B site creates an active and well dispersed metal locked into place in the crystal structure. This greatly reduces the metal sintering that typically occurs on supported catalysts used in reforming reactions, and reduces deactivation by sulfur and carbon. Further, oxygen mobility may also be enhanced by elemental exchange of promoters at sites in the pyrochlore. The pyrochlore catalyst material may be utilized in catalytic reforming reactions for the conversion of hydrocarbon fuels into synthesis gas (H.sub.2+CO) for fuel cells, among other uses.

  7. Microbial contamination of stored hydrocarbon fuels and its control

    Directory of Open Access Journals (Sweden)

    Gaylarde Christine C.

    1999-01-01

    Full Text Available The major microbial problem in the petroleum refining industry is contamination of stored products, which can lead to loss of product quality, formation of sludge and deterioration of pipework and storage tanks, both in the refinery and at the end-user. Three major classes of fuel are discussed in this article - gasoline, aviation kerosene and diesel, corresponding to increasingly heavy petroleum fractions. The fuel that presents the most serious microbiological problems is diesel. The many microorganisms that have been isolated from hydrocarbon fuel systems are listed. The conditions required for microbial growth and the methods used to monitor and to control this activity are discussed. The effects of various fuel additives, including biocides, are considered.

  8. Radical recombination in a hydrocarbon-fueled scramjet nozzle

    Directory of Open Access Journals (Sweden)

    Zhang Xiaoyuan

    2014-12-01

    Full Text Available To reveal the radical recombination process in the scramjet nozzle flow and study the effects of various factors of the recombination, weighted essentially non-oscillatory (WENO schemes are applied to solve the decoupled two-dimensional Euler equations with chemical reactions to simulate the hydrocarbon-fueled scramjet nozzle flow. The accuracy of the numerical method is verified with the measurements obtained by a shock tunnel experiment. The overall model length is nearly 0.5 m, with inlet static temperatures ranging from 2000 K to 3000 K, inlet static pressures ranging from 75 kPa to 175 kPa, and inlet Mach numbers of 2.0 ± 0.4 are involved. The fraction Damkohler number is defined as functions of static temperature and pressure to analyze the radical recombination progresses. Preliminary results indicate that the energy releasing process depends on different chemical reaction processes and species group contributions. In hydrocarbon-fueled scramjet nozzle flow, reactions with H have the greatest contribution during the chemical equilibrium shift. The contrast and analysis of the simulation results show that the radical recombination processes influenced by inflow conditions and nozzle scales are consistent with Damkohler numbers and potential dissociation energy release. The increase of inlet static temperature improves both of them, thus making the chemical non-equilibrium effects on the nozzle performance more significant. While the increase of inlet static pressure improves the former one and reduces the latter, it exerts little influence on the chemical non-equilibrium effects.

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

    Directory of Open Access Journals (Sweden)

    Zhukov Ilya S.

    2016-01-01

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

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

    OpenAIRE

    Zhukov Ilya S.; Borisov Boris V.; Bondarchuk Sergey S.; Zhukov Alexander S.

    2016-01-01

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

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

    Energy Technology Data Exchange (ETDEWEB)

    Friesen, H.N.

    1995-06-01

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

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

    Data.gov (United States)

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

  13. Engine and method for operating an engine

    Science.gov (United States)

    Lauper, Jr., John Christian; Willi, Martin Leo; Thirunavukarasu, Balamurugesh; Gong, Weidong

    2008-12-23

    A method of operating an engine is provided. The method may include supplying a combustible combination of reactants to a combustion chamber of the engine, which may include supplying a first hydrocarbon fuel, hydrogen fuel, and a second hydrocarbon fuel to the combustion chamber. Supplying the second hydrocarbon fuel to the combustion chamber may include at least one of supplying at least a portion of the second hydrocarbon fuel from an outlet port that discharges into an intake system of the engine and supplying at least a portion of the second hydrocarbon fuel from an outlet port that discharges into the combustion chamber. Additionally, the method may include combusting the combustible combination of reactants in the combustion chamber.

  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. Design study of RL10 derivatives. Volume 2: Engine design characteristics. [application of rocket engine to space tug propulsion

    Science.gov (United States)

    Adams, A.

    1973-01-01

    The design characteristics of the RL-10 rocket engine are discussed. The results from critical elements evaluation, baseline engine design, parametric and special study tasks are presented. Critical element evaluation established the feasibility of various engine features such as tank head idle, pumped idle, autogenous tank pressurization, and two-phase pumping. Three baseline engines, derived from the RL-10 were conceptually designed. Parametric life and performance data were generated. Special studies were conducted to establish the impact on the engine design of environment, safety, interchangeability, and maintenance.

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

    OpenAIRE

    Glynne-Jones, Peter; Coletti, Michele; White, Neil M.; Gabriel, Stephen; Bramanti, Cristina

    2010-01-01

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

  17. Rocket Engine Nozzle Side Load Transient Analysis Methodology: A Practical Approach

    Science.gov (United States)

    Shi, John J.

    2005-01-01

    At the sea level, a phenomenon common with all rocket engines, especially for a highly over-expanded nozzle, during ignition and shutdown is that of flow separation as the plume fills and empties the nozzle, Since the flow will be separated randomly. it will generate side loads, i.e. non-axial forces. Since rocket engines are designed to produce axial thrust to power the vehicles, it is not desirable to be excited by non-axial input forcing functions, In the past, several engine failures were attributed to side loads. During the development stage, in order to design/size the rocket engine components and to reduce the risks, the local dynamic environments as well as dynamic interface loads have to be defined. The methodology developed here is the way to determine the peak loads and shock environments for new engine components. In the past it is not feasible to predict the shock environments, e.g. shock response spectra, from one engine to the other, because it is not scaleable. Therefore, the problem has been resolved and the shock environments can be defined in the early stage of new engine development. Additional information is included in the original extended abstract.

  18. 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.

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

    NARCIS (Netherlands)

    Sliphorst, M.

    2011-01-01

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

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

    NARCIS (Netherlands)

    Sliphorst, M.

    2011-01-01

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

  1. State-space analysis of the dynamic characteristics of a variable thrust liquid propellant rocket engine

    Science.gov (United States)

    Zhang, Yu-Lin

    This paper states the application of state-space method to the analysis of the dynamic characteristics of a variable thrust liquid propellant rocket engine and presents a set of state equations for describing the dynamic process of the engine. An efficient numerical method for solving these system equations is developed. The theoretical solutions agree well with the experimental data. The analysis leads to the following conclusion: the set coefficient of the pulse width, the working frequency of the solenoid valves and the deviation of the critical working points of these valves are important parameters for determining the dynamic response time and the control precision of this engine. The methods developed in this paper may be used effectively in the analysis of dynamic characteristics of variable thrust liquid propellant rocket engines.

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

    OpenAIRE

    Sliphorst, M.

    2011-01-01

    Ever since the early stages of space transportation in the 1940’s, and the related liquid propellant rocket engine development, combustion instability has been a major issue. High frequency combustion instability (HFCI) is the interaction between combustion and the acoustic field in the combustion chamber. It destroys the thermal boundary layer wall increasing heat transfer and could lead to compromised performance, and ultimately to destruction of the engine and mission loss. The main object...

  3. Endothermic Heat-Sink of Hydrocarbon Fuels for Scramjet Cooling AIAA 2002-3871

    National Research Council Canada - National Science Library

    Huang, H

    2002-01-01

    Storable liquid hydrocarbon fuels, such as JP-7, JP-8+ 100, and JP-10, that can undergo endothermic reactions may provide sufficient heat sink to enable hypersonic flight without having to resort to cryogenic fuels...

  4. Lightweight Exit Cone for Liquid Rocket Engines Project

    Data.gov (United States)

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

  5. Hydrocarbon Fuel Thermal Performance Modeling based on Systematic Measurement and Comprehensive Chromatographic Analysis

    Science.gov (United States)

    2016-07-31

    distribution unlimited Hydrocarbon Fuel Thermal Performance Modeling based on Systematic Measurement and Comprehensive Chromatographic Analysis Matthew...Technical Note 3. DATES COVERED (From - To) 04 January 2016 - 31 July 2016 4. TITLE AND SUBTITLE Hydrocarbon Fuel Thermal Performance Modeling based on...Systematic Measurement and Comprehensive Chromatographic Analysis 5a. CONTRACT NUMBER 5b. GRANT NUMBER 5c. PROGRAM ELEMENT NUMBER 6. AUTHOR(S

  6. Potential Applications of the Ceramic Thrust Chamber Technology for Future Transpiration Cooled Rocket Engines

    Science.gov (United States)

    Herbertz, Armin; Ortelt, Markus; Müller, Ilja; Hald, Hermann

    The long-term development of ceramic rocket engine thrust chambers at the German Aerospace Center(DLR) currently leads to designs of self-sustaining, transpiration-cooled, fiber-reinforced ceramic rocket engine chamber structures.This paper discusses characteristic issues and potential benefits introduced by this technology. Achievable benefits are the reduction of weight and manufacturing cost, as well as an increased reliability and higher lifetime due to thermal cycle stability.Experiments with porous Ceramic Matrix Composite(CMC) materials for rocket engine chamber walls have been conducted at the DLR since the end of the 1990s.This paper discusses the current status of DLR's ceramic thrust chamber technology and potential applications for high thrust engines.The manufacturing process and the design concept are explained.The impact of variations of engine parameters(chamber pressure and diam-eter)on the required coolant mass flow are discussed.Due to favorable scaling effects a high thrust application utilizes all benefits of the discussed technology, while avoiding the most significant performance drawbacks.

  7. Linear Static and Dynamic Analysis of Rocket Engine Testing Bench Structure using the Finite Element Method

    Directory of Open Access Journals (Sweden)

    Luiza Fabrino Favato

    2015-04-01

    Full Text Available This article presents a study of a testing bench structure for Rocket Engines, which is under development by the PUC-Minas Aerospace Research Group. The Bench is being built for civilian’s liquid bipropellant rocket engines up to 5 kN of thrust. The purpose of this article is to evaluate the bench structure using the Finite Element Method (FEM, by structural linear static and dynamic analysis. Performed to predict the behavior of the structure to the requests of the tests. The virtual simulations were performed using a CAE software with the Nastran solver. The structure is 979 x 1638 mm by 2629 mm, consisting of folded-plates (¼ "x 3¼" x 8" and plates of 1/4" and 1/2 ", both SAE 1020 Steel .The rocket engine is fixed on the structure through a set called engine mount. It was included in the analysis clearances or misalignments that may occur during tests. As well as, the load applied was evaluated with components in varying orientations and directions. It was considered the maximum size of the engine mount and the maximum inclination angle of load. At the end of this article it was observed that the worst stress and displacement values obtained were for the hypothesis with the inclination of five-degrees with load components in the positive directions of the axes defined and it was also obtained the first twenty frequency modes of the structure.

  8. Methods of reforming hydrocarbon fuels using hexaaluminate catalysts

    Science.gov (United States)

    Gardner, Todd H [Morgantown, WV; Berry, David A [Morgantown, WV; Shekhawat, Dushyant [Morgantown, WV

    2012-03-27

    A metal substituted hexaaluminate catalyst for reforming hydrocarbon fuels to synthesis gas of the general formula AB.sub.yAl.sub.12-yO.sub.19-.delta., A being selected from alkali metals, alkaline earth metals and lanthanide metals or mixtures thereof. A dopant or surface modifier selected from a transitions metal, a spinel of an oxygen-ion conductor is incorporated. The dopant may be Ca, Cs, K, La, Sr, Ba, Li, Mg, Ce, Co, Fe, Ir, Rh, Ni, Ru, Cu, Pe, Os, Pd, Cr, Mn, W, Re, Sn, Gd, V, Ti, Ag, Au, and mixtures thereof. The oxygen-ion conductor may be a perovskite selected from M'RhO.sub.3, M'PtO.sub.3, M'PdO.sub.3, M'IrO.sub.3, M'RuO.sub.3 wherein M'=Mg, Sr, Ba, La, Ca; a spinel selected from MRh.sub.2O.sub.4, MPt.sub.2O.sub.4, MPd.sub.2O.sub.4, MIr.sub.2O.sub.4, MRu.sub.2O.sub.4 wherein M=Mg, Sr, Ba, La, Ca and mixtures thereof; a florite is selected from M''O.sub.2.

  9. Preliminary Economics for Hydrocarbon Fuel Production from Cellulosic Sugars

    Energy Technology Data Exchange (ETDEWEB)

    Collett, James R.; Meyer, Pimphan A.; Jones, Susanne B.

    2014-05-18

    Biorefinery process and economic models built in CHEMCAD and a preliminary, genome-scale metabolic model for the oleaginous yeast Lipomyces starkeyi were used to simulate the bioconversion of corn stover to lipids, and the upgrading of these hydrocarbon precursors to diesel and jet fuel. The metabolic model was based on the recently released genome sequence for L. starkeyi and on metabolic pathway information from the literature. The process model was based on bioconversion, lipid extraction, and lipid oil upgrading data found in literature, on new laboratory experimental data, and on yield predictions from the preliminary L. starkeyi metabolic model. The current plant gate production cost for a distillate-range hydrocarbon fuel was estimated by the process model Base Case to be $9.5/gallon ($9.0 /gallon of gasoline equivalent) with assumptions of 2011$, 10% internal return on investment, and 2205 ton/day dry feed rate. Opportunities for reducing the cost to below $5.0/gallon, such as improving bioconversion lipid yield and hydrogenation catalyst selectivity, are presented in a Target Case. The process and economic models developed for this work will be updated in 2014 with new experimental data and predictions from a refined metabolic network model for L. starkeyi. Attaining a production cost of $3.0/gallon will require finding higher value uses for lignin other than power generation, such as conversion to additional fuel or to a co-product.

  10. Thermocatalytic CO2-Free Production of Hydrogen from Hydrocarbon Fuels

    Energy Technology Data Exchange (ETDEWEB)

    University of Central Florida

    2004-01-30

    The main objective of this project is the development of an economically viable thermocatalytic process for production of hydrogen and carbon from natural gas or other hydrocarbon fuels with minimal environmental impact. The three major technical goals of this project are: (1) to accomplish efficient production of hydrogen and carbon via sustainable catalytic decomposition of methane or other hydrocarbons using inexpensive and durable carbon catalysts, (2) to obviate the concurrent production of CO/CO{sub 2} byproducts and drastically reduce CO{sub 2} emissions from the process, and (3) to produce valuable carbon products in order to reduce the cost of hydrogen production The important feature of the process is that the reaction is catalyzed by carbon particulates produced in the process, so no external catalyst is required (except for the start-up operation). This results in the following advantages: (1) no CO/CO{sub 2} byproducts are generated during hydrocarbon decomposition stage, (2) no expensive catalysts are used in the process, (3) several valuable forms of carbon can be produced in the process depending on the process conditions (e.g., turbostratic carbon, pyrolytic graphite, spherical carbon particles, carbon filaments etc.), and (4) CO{sub 2} emissions could be drastically reduced (compared to conventional processes).

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

    Science.gov (United States)

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

    2003-01-01

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

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

    Science.gov (United States)

    Ali, Moonis; Gupta, U. K.

    1990-01-01

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

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

    Science.gov (United States)

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

    2012-01-01

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

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

    Science.gov (United States)

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

    2012-01-01

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

  15. Effects of supercritical environment on hydrocarbon-fuel injection

    Science.gov (United States)

    Shin, Bongchul; Kim, Dohun; Son, Min; Koo, Jaye

    2017-04-01

    In this study, the effects of environment conditions on decane were investigated. Decane was injected in subcritical and supercritical ambient conditions. The visualization chamber was pressurized to 1.68 MPa by using nitrogen gas at a temperature of 653 K for subcritical ambient conditions. For supercritical ambient conditions, the visualization chamber was pressurized to 2.52 MPa by using helium at a temperature of 653 K. The decane injection in the pressurized chamber was visualized via a shadowgraph technique and gradient images were obtained by a post processing method. A large variation in density gradient was observed at jet interface in the case of subcritical injection in subcritical ambient conditions. Conversely, for supercritical injection in supercritical ambient conditions, a small density gradient was observed at the jet interface. In a manner similar to that observed in other cases, supercritical injection in subcritical ambient conditions differed from supercritical ambient conditions such as sphere shape liquid. Additionally, there were changes in the interface, and the supercritical injection core width was thicker than that in the subcritical injection. Furthermore, in cases with the same injection conditions, the change in the supercritical ambient normalized core width was smaller than the change in the subcritical ambient normalized core width owing to high specific heat at the supercritical injection and small phase change at the interface. Therefore, the interface was affected by the changing ambient condition. Given that the effect of changing the thermodynamic properties of propellants could be essential for a variable thrust rocket engine, the effects of the ambient conditions were investigated experimentally.

  16. Health monitoring of rocket engines using image processing

    Science.gov (United States)

    Disimile, Peter J.; Shoe, Bridget; Toy, Norman

    1991-07-01

    Analysis of spectral and video data for anomalous events occurring in the exhaust plume of the Space Shuttle Main Engine (SSME) has shown that the improved time resolution of video tape increases the detection rate of anomalies in the visual region. Preliminary developments and applications of image processing techniques are used to extract information from video data of the SSME exhaust plume. Images have been enhanced to show the exhaust plume shock structure and for the isolation of an anomalous event.

  17. 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.

  18. Design Considerations for Human Rating of Liquid Rocket Engines

    Science.gov (United States)

    Parkinson, Douglas

    2010-01-01

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

  19. Damage-mitigating control of a reusable rocket engine for high performance and extended life

    Science.gov (United States)

    Ray, Asok; Dai, Xiaowen

    1995-01-01

    The goal of damage mitigating control in reusable rocket engines is to achieve high performance with increased durability of mechanical structures such that functional lives of the critical components are increased. The major benefit is an increase in structural durability with no significant loss of performance. This report investigates the feasibility of damage mitigating control of reusable rocket engines. Phenomenological models of creep and thermo-mechanical fatigue damage have been formulated in the state-variable setting such that these models can be combined with the plant model of a reusable rocket engine, such as the Space Shuttle Main Engine (SSME), for synthesizing an optimal control policy. Specifically, a creep damage model of the main thrust chamber wall is analytically derived based on the theories of sandwich beam and viscoplasticity. This model characterizes progressive bulging-out and incremental thinning of the coolant channel ligament leading to its eventual failure by tensile rupture. The objective is to generate a closed form solution of the wall thin-out phenomenon in real time where the ligament geometry is continuously updated to account for the resulting deformation. The results are in agreement with those obtained from the finite element analyses and experimental observation for both Oxygen Free High Conductivity (OFHC) copper and a copper-zerconium-silver alloy called NARloy-Z. Due to its computational efficiency, this damage model is suitable for on-line applications of life prediction and damage mitigating control, and also permits parametric studies for off-line synthesis of damage mitigating control systems. The results are presented to demonstrate the potential of life extension of reusable rocket engines via damage mitigating control. The control system has also been simulated on a testbed to observe how the damage at different critical points can be traded off without any significant loss of engine performance. The research work

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

    Science.gov (United States)

    Clark, John S.; Walton, James T.; Mcguire, Melissa L.

    1992-01-01

    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.

  1. Stennis Space Center's approach to liquid rocket engine health monitoring using exhaust plume diagnostics

    Science.gov (United States)

    Gardner, D. G.; Tejwani, G. D.; Bircher, F. E.; Loboda, J. A.; Van Dyke, D. B.; Chenevert, D. J.

    1991-01-01

    Details are presented of the approach used in a comprehensive program to utilize exhaust plume diagnostics for rocket engine health-and-condition monitoring and assessing SSME component wear and degradation. This approach incorporates both spectral and video monitoring of the exhaust plume. Video monitoring provides qualitative data for certain types of component wear while spectral monitoring allows both quantitative and qualitative information. Consideration is given to spectral identification of SSME materials and baseline plume emissions.

  2. Application of advanced coating techniques to rocket engine components

    Science.gov (United States)

    Verma, S. K.

    1988-01-01

    The materials problem in the space shuttle main engine (SSME) is reviewed. Potential coatings and the method of their application for improved life of SSME components are discussed. A number of advanced coatings for turbine blade components and disks are being developed and tested in a multispecimen thermal fatigue fluidized bed facility at IIT Research Institute. This facility is capable of producing severe strains of the degree present in blades and disk components of the SSME. The potential coating systems and current efforts at IITRI being taken for life extension of the SSME components are summarized.

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

    Data.gov (United States)

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

  4. Fault Detection and Diagnosis Techniques for Liquid-Propellant Rocket Propellant Engines

    Science.gov (United States)

    Wua, Jianjun; Tanb, Songlin

    2002-01-01

    Fault detection and diagnosis plays a pivotal role in the health-monitoring techniques for liquid- propellant rocket engines. This paper firstly gives a brief summary on the techniques of fault detection and diagnosis utilized in liquid-propellant rocket engines. Then, the applications of fault detection and diagnosis algorithms studied and developed to the Long March Main Engine System(LMME) are introduced. For fault detection, an analytical model-based detection algorithm, a time-series-analysis algorithm and a startup- transient detection algorithm based on nonlinear identification developed and evaluated through ground-test data of the LMME are given. For fault diagnosis, neural-network approaches, nonlinear-static-models based methods, and knowledge-based intelligent approaches are presented. Keywords: Fault detection; Fault diagnosis; Health monitoring; Neural networks; Fuzzy logic; Expert system; Long March main engines Contact author and full address: Dr. Jianjun Wu Department of Astronautical Engineering School of Aerospace and Material Engineering National University of Defense Technology Changsha, Hunan 410073 P.R.China Tel:86-731-4556611(O), 4573175(O), 2219923(H) Fax:86-731-4512301 E-mail:jjwu@nudt.edu.cn

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

    Science.gov (United States)

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

    2011-01-01

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

  6. Theoretical prediction of regression rates in swirl-injection hybrid rocket engines

    Science.gov (United States)

    Ozawa, K.; Shimada, T.

    2016-07-01

    The authors theoretically and analytically predict what times regression rates of swirl injection hybrid rocket engines increase higher than the axial injection ones by estimating heat flux from boundary layer combustion to the fuel port. The schematic of engines is assumed as ones whose oxidizer is injected from the opposite side of the nozzle such as ones of Yuasa et al. propose. To simplify the estimation, we assume some hypotheses such as three-dimensional (3D) axisymmetric flows have been assumed. The results of this prediction method are largely consistent with Yuasa's experiments data in the range of high swirl numbers.

  7. Reliability Analysis of a Rocket Engine Using Design for Six Sigma

    Science.gov (United States)

    Kobayashi, Hiroaki; Sato, Tetsuya; Tanatsugu, Nobuhiro

    Six Sigma is the management strategy developed by Motorola to reduce defects in products. Design for Six Sigma (DFSS) is a methodology for determining the values of the design parameters, which maximize the performance of some system without tightening the material, manufacturing or environmental tolerances. This paper presents the implementation of DFSS for redesign of the LE-7 engine. Uncertainties with design parameters and operational conditions are considered in evaluating thrust performance, thrust chamber life, turbo-pump cavitation, and combustion stability. Traditional deterministic optimization results and probabilistic optimization results are compared. It is found that robustness of rocket engine is not always consistent with the extension of thrust chamber life.

  8. With and Without Post-Burning Solar Thermal Rocket Engines: Three New Chances for Space Propulsion

    Science.gov (United States)

    Ruiz Haro, Mercedes; Navarro Vásquez, Ricardo M.

    2002-01-01

    This report studies and compares Solar Thermal Rocket Engines (STRE) with and without post-burning. In a STRE hydrogen is expelled at very high speeds after been heated up to 3000 K thanks to the concentrator-receiver system. In Solar Rocket Engines with Post-Burning (STREPB), this hydrogen is burnt inside a especial combustion chamber where the oxygen is introduced. In this paper the addition of another fuel, LiH, will be also studied. The simple STRE gives higher values for specific impulse than the other two cases. While these values for this configuration go to more than 1000 s, the STREPB reaches around 650 s for hydrogen temperatures of 1500 K. The solution using H2-LiH- O2 gives around 520 s at only 800 K. The consecution of a high temperature is linked to an increase of concentrator's accuracy and mass. For the expedient value of oxidizer-to-fuel ratio the difference of more than 500 K is enough to enable a reduction higher than 50% of the concentrator's area and mass. The calculations for obtained thrust can be approach by means of several thermodynamic equations. It will be less for the STRE, so the use of Post-Burning will be better for missions requiring higher thrust. These figures locate STRE and STREPB between Liquid Rocket Engines' high thrust, which reduce trip time, and the Ion Accelerating Rockets' high specific impulse, which increase the admitted payload's mass. This paper will also compare this kind of propulsion with existing ones by means of Tsiolkovsky equation, V = I spLn M 0 / M p to estimate its possibilities for different manoeuvres as orbit transfers and interplanetary missions.

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

    Science.gov (United States)

    Mazzetti, Alessandro; Merotto, Laura; Pinarello, Giordano

    2016-09-01

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

  10. Development of the platelet micro-orifice injector. [for liquid propellant rocket engines

    Science.gov (United States)

    La Botz, R. J.

    1984-01-01

    For some time to come, liquid rocket engines will continue to provide the primary means of propulsion for space transportation. The injector represents a key to the optimization of engine and system performance. The present investigation is concerned with a unique injector design and fabrication process which has demonstrated performance capabilities beyond that achieved with more conventional approaches. This process, which is called the 'platelet process', makes it feasible to fabricate injectors with a pattern an order of magnitude finer than that obtainable by drilling. The fine pattern leads to an achievement of high combustion efficiencies. Platelet injectors have been identified as one of the significant technology advances contributing to the feasibility of advanced dual-fuel booster engines. Platelet injectors are employed in the Space Shuttle Orbit Maneuvering System (OMS) engines. Attention is given to injector design theory as it relates to pattern fineness, a description of platelet injectors, and test data obtained with three different platelet injectors.

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

    Science.gov (United States)

    Liou, Larry C.

    2008-01-01

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

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

    Science.gov (United States)

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

    2010-07-01

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

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

    Directory of Open Access Journals (Sweden)

    Qiang WEI

    2017-08-01

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

  14. 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

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

    Energy Technology Data Exchange (ETDEWEB)

    Koehlinger, M.W.; Bennett, R.G.; Motloch, C.G. [eds.; 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.

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

    Directory of Open Access Journals (Sweden)

    Zolotorev Nikolay

    2017-01-01

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

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

    Science.gov (United States)

    Gradl, Paul

    2016-01-01

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

  18. Liquid propellant rocket engine combustion simulation with a time-accurate CFD method

    Science.gov (United States)

    Chen, Y. S.; Shang, H. M.; Liaw, Paul; Hutt, J.

    1993-01-01

    Time-accurate computational fluid dynamics (CFD) algorithms are among the basic requirements as an engineering or research tool for realistic simulations of transient combustion phenomena, such as combustion instability, transient start-up, etc., inside the rocket engine combustion chamber. A time-accurate pressure based method is employed in the FDNS code for combustion model development. This is in connection with other program development activities such as spray combustion model development and efficient finite-rate chemistry solution method implementation. In the present study, a second-order time-accurate time-marching scheme is employed. For better spatial resolutions near discontinuities (e.g., shocks, contact discontinuities), a 3rd-order accurate TVD scheme for modeling the convection terms is implemented in the FDNS code. Necessary modification to the predictor/multi-corrector solution algorithm in order to maintain time-accurate wave propagation is also investigated. Benchmark 1-D and multidimensional test cases, which include the classical shock tube wave propagation problems, resonant pipe test case, unsteady flow development of a blast tube test case, and H2/O2 rocket engine chamber combustion start-up transient simulation, etc., are investigated to validate and demonstrate the accuracy and robustness of the present numerical scheme and solution algorithm.

  19. The use of low power dual mode nuclear thermal rocket engines to support space exploration missions

    Science.gov (United States)

    Zubrin, Robert M.

    1991-01-01

    The evolution of dual mode concepts is presented, focusing on advantages and problems associated with both low and high temperature dual mode conversion systems. It is concluded that dual mode nuclear thermal rocket (NTR) systems using high temperature Brayton cycle conversion technology offer a high payoff enhancement of conventional NTR, with a comparatively minor increase of technological challenge. It is recommended that NTR engines be designed so that dual mode conversion systems can be attached to them in a modular way, thus enabling the production of electric power on all missions where it is needed.

  20. Computer Modeling of a Rotating Detonation Engine in a Rocket Configuration

    Science.gov (United States)

    2015-03-01

    coefficient CP Specific heat capacity at constant pressure ( J kg−K ) CS Nozzle stream thrust coefficient D Detonation wave speed in laboratory frame-of...greater than the detonation fuel-to-air ratio, the ratio of specific heats and gas constant at station c3.4 are calculated using Eq. 75 and Eq. 76...COMPUTER MODELING OF A ROTATING DETONATION ENGINE IN A ROCKET CONFIGURATION THESIS Nihar N. Shah, 1st Lt, USAF AFIT-ENY-MS-15-M-230 DEPARTMENT OF THE

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

    CERN Document Server

    Kitsche, Wolfgang

    2010-01-01

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

  2. Raman Spectroscopy for Instantaneous Multipoint, Multispecies Gas Concentration and Temperature Measurements in Rocket Engine Propellant Injector Flows

    Science.gov (United States)

    Wehrmeyer, Joseph A.; Trinh, Huu Phuoc

    2001-01-01

    Propellant injector development at MSFC includes experimental analysis using optical techniques, such as Raman, fluorescence, or Mie scattering. For the application of spontaneous Raman scattering to hydrocarbon-fueled flows a technique needs to be developed to remove the interfering polycyclic aromatic hydrocarbon fluorescence from the relatively weak Raman signals. A current application of such a technique is to the analysis of the mixing and combustion performance of multijet, impinging-jet candidate fuel injectors for the baseline Mars ascent engine, which will burn methane and liquid oxygen produced in-situ on Mars to reduce the propellant mass transported to Mars for future manned Mars missions. The present technique takes advantage of the strongly polarized nature of Raman scattering. It is shown to be discernable from unpolarized fluorescence interference by subtracting one polarized image from another. Both of these polarized images are obtained from a single laser pulse by using a polarization-separating calcite rhomb mounted in the imaging spectrograph. A demonstration in a propane-air flame is presented.

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

    Directory of Open Access Journals (Sweden)

    Moritz Schulze

    2016-09-01

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

  4. Thermal radiation of heterogeneous combustion products in the model rocket engine plume

    Science.gov (United States)

    Kuzmin, V. A.; Maratkanova, E. I.; Zagray, I. A.; Rukavishnikova, R. V.

    2015-05-01

    The work presents a method of complex investigation of thermal radiation emitted by heterogeneous combustion products in the model rocket engine plume. Realization of the method has allowed us to obtain full information on the results in all stages of calculations. Dependence of the optical properties (complex refractive index), the radiation characteristics (coefficients and cross sections) and emission characteristics (flux densities, emissivity factors) of the main determining factors and parameters was analyzed. It was found by the method of computational experiment that the presence of the gaseous phase in the combustion products causes a strongly marked selectivity of emission, due to which the use of gray approximation in the calculation of thermal radiation is unnecessary. The influence of the optical properties, mass fraction, the function of particle size distribution, and the temperature of combustion products on thermal radiation in the model rocket engine plume was investigated. The role of "spotlight" effect-increasing the amount of energy of emission exhaust combustion products due to scattering by condensate particles radiation from the combustion chamber-was established quantitatively.

  5. Thermo-mechanical concepts applied to modeling liquid propellant rocket engine stability

    Science.gov (United States)

    Kassoy, David R.; Norris, Adam

    2016-11-01

    The response of a gas to transient, spatially distributed energy addition can be quantified mathematically using thermo-mechanical concepts available in the literature. The modeling demonstrates that the ratio of the energy addition time scale to the acoustic time scale of the affected volume, and the quantity of energy added to that volume during the former determine the whether the responses to heating can be described as occurring at nearly constant volume, fully compressible or nearly constant pressure. Each of these categories is characterized by significantly different mechanical responses. Application to idealized configurations of liquid propellant rocket engines provides an opportunity to identify physical conditions compatible with gasdynamic disturbances that are sources of engine instability. Air Force Office of Scientific Research.

  6. Heat transfer in rocket engine combustion chambers and regeneratively cooled nozzles

    Science.gov (United States)

    1993-01-01

    A conjugate heat transfer computational fluid dynamics (CFD) model to describe regenerative cooling in the main combustion chamber and nozzle and in the injector faceplate region for a launch vehicle class liquid rocket engine was developed. An injector model for sprays which treats the fluid as a variable density, single-phase media was formulated, incorporated into a version of the FDNS code, and used to simulate the injector flow typical of that in the Space Shuttle Main Engine (SSME). Various chamber related heat transfer analyses were made to verify the predictive capability of the conjugate heat transfer analysis provided by the FDNS code. The density based version of the FDNS code with the real fluid property models developed was successful in predicting the streamtube combustion of individual injector elements.

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

    Science.gov (United States)

    Gradl, Paul R.

    2016-01-01

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

  8. 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.

  9. Residual Fuel Expulsion from a Simulated 50,000 Pound Thrust Liquid-Propellant Rocket Engine Having a Continuous Rocket-Type Igniter

    Science.gov (United States)

    Messing, Wesley E.

    1959-01-01

    Tests have been conducted to determine the starting characteristics of a 50,000-pound-thrust rocket engine with the conditions of a quantity of fuel lying dormant in the simulated main thrust chamber. Ignition was provided by a smaller rocket firing rearwardly along the center line. Both alcohol-water and anhydrous ammonia were used as the residual fuel. The igniter successfully expelled the maximum amount of residual fuel (3 1/2 gal) in 2.9 seconds when the igniter.was equipped with a sonic discharge nozzle operating at propellant flow rates of 3 pounds per second. Lesser amounts of residual fuel required correspondingly lower expulsion times. When the igniter was equipped with a supersonic exhaust nozzle operating at a flow of 4 pounds per second, a slightly less effective expulsion rate was encountered.

  10. 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

  11. Primary atomization of liquid jets issuing from rocket engine coaxial injectors

    Science.gov (United States)

    Woodward, Roger D.

    1993-01-01

    The investigation of liquid jet breakup and spray development is critical to the understanding of combustion phenomena in liquid-propellant rocket engines. Much work has been done to characterize low-speed liquid jet breakup and dilute sprays, but atomizing jets and dense sprays have yielded few quantitative measurements due to their optical opacity. This work focuses on a characteristic of the primary breakup process of round liquid jets, namely the length of the intact liquid core. The specific application considered is that of shear-coaxial type rocket engine injectors. Real-time x-ray radiography, capable of imaging through the dense two-phase region surrounding the liquid core, has been used to make the measurements. Nitrogen and helium were employed as the fuel simulants while an x-ray absorbing potassium iodide aqueous solution was used as the liquid oxygen (LOX) simulant. The intact-liquid-core length data have been obtained and interpreted to illustrate the effects of chamber pressure (gas density), injected-gas and liquid velocities, and cavitation. The results clearly show that the effect of cavitation must be considered at low chamber pressures since it can be the dominant breakup mechanism. A correlation of intact core length in terms of gas-to-liquid density ratio, liquid jet Reynolds number, and Weber number is suggested. The gas-to-liquid density ratio appears to be the key parameter for aerodynamic shear breakup in this study. A small number of hot-fire, LOX/hydrogen tests were also conducted to attempt intact-LOX-core measurements under realistic conditions in a single-coaxial-element rocket engine. The tests were not successful in terms of measuring the intact core, but instantaneous imaging of LOX jets suggests that LOX jet breakup is qualitatively similar to that of cold-flow, propellant-simulant jets. The liquid oxygen jets survived in the hot-fire environment much longer than expected, and LOX was even visualized exiting the chamber nozzle

  12. Towards Flange-to-Flange Turbopump Simulations for Liquid Rocket Engines

    Science.gov (United States)

    Kiris, Cetin; Williams, Robert

    2000-01-01

    The primary objective of this research is to support the design of liquid rocket systems for the Advanced Space Transportation System. Since the space launch systems in the near future are likely to rely on liquid rocket engines, increasing the efficiency and reliability of the engine components is an important task. One of the major problems in the liquid rocket engine is to understand fluid dynamics of fuel and oxidizer flows from the fuel tank to plume. Understanding the flow through the entire turbopump geometry through numerical simulation will be of significant value toward design. This will help to improve safety of future space missions. One of the milestones of this effort is to develop, apply and demonstrate the capability and accuracy of 3D CFD methods as efficient design analysis tools on high performance computer platforms. The development of the MPI and MLP versions of the INS3D code is currently underway. The serial version of INS3D code is a multidimensional incompressible Navier-Stokes solver based on overset grid technology. INS3D-MPI is based on the explicit massage-passing interface across processors and is primarily suited for distributed memory systems. INS3D-MLP is based on multi-level parallel method and is suitable for distributed-shared memory systems. For the entire turbopump simulations, moving boundary capability and an efficient time-accurate integration methods are build in the flow solver. To handle the geometric complexity and moving boundary problems, overset grid scheme is incorporated with the solver that new connectivity data will be obtained at each time step. The Chimera overlapped grid scheme allows subdomains move relative to each other, and provides a great flexibility when the boundary movement creates large displacements. The performance of the two time integration schemes for time-accurate computations is investigated. For an unsteady flow which requires small physical time step, the pressure projection method was found

  13. An experimental investigation of reacting and nonreacting coaxial jet mixing in a laboratory rocket engine

    Science.gov (United States)

    Schumaker, Stephen Alexander

    Coaxial jets are commonly used as injectors in propulsion and combustion devices due to both the simplicity of their geometry and the rapid mixing they provide. In liquid rocket engines it is common to use coaxial jets in the context of airblast atomization. However, interest exists in developing rocket engines using a full flow staged combustion cycle. In such a configuration both propellants are injected in the gaseous phase. In addition, gaseous coaxial jets have been identified as an ideal test case for the validation of the next generation of injector modeling tools. For these reasons an understanding of the fundamental phenomena which govern mixing in gaseous coaxial jets and the effect of combustion on these phenomena in coaxial jet diffusion flames is needed. A study was performed to better understand the scaling of the stoichiometric mixing length in reacting and nonreacting coaxial jets with velocity ratios greater than one and density ratios less than one. A facility was developed that incorporates a single shear coaxial injector in a laboratory rocket engine capable of ten atmospheres. Optical access allows the use of flame luminosity and laser diagnostic techniques such as Planar Laser Induced Fluorescence (PLIF). Stoichiometric mixing lengths (LS), which are defined as the distance along the centerline where the stoichiometric condition occurs, were measured using PLIF. Acetone was seeded into the center jet to provide direct PLIF measurement of the average and instantaneous mixture fraction fields for a range of momentum flux ratios for the nonreacting cases. For the coaxial jet diffusion flames, LS was measured from OH radical contours. For nonreacting cases the use of a nondimensional momentum flux ratio was found to collapse the mixing length data. The flame lengths of coaxial jet diffusion flames were also found to scale with the momentum flux ratio but different scaling constants are required which depended on the chemistry of the reaction. The

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

    Science.gov (United States)

    Casiano, Matthew J.

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

  15. Fuel/oxidizer-rich high-pressure preburners. [staged-combustion rocket engine

    Science.gov (United States)

    Schoenman, L.

    1981-01-01

    The analyses, designs, fabrication, and cold-flow acceptance testing of LOX/RP-1 preburner components required for a high-pressure staged-combustion rocket engine are discussed. Separate designs of injectors, combustion chambers, turbine simulators, and hot-gas mixing devices are provided for fuel-rich and oxidizer-rich operation. The fuel-rich design addresses the problem of non-equilibrium LOX/RP-1 combustion. The development and use of a pseudo-kinetic combustion model for predicting operating efficiency, physical properties of the combustion products, and the potential for generating solid carbon is presented. The oxygen-rich design addresses the design criteria for the prevention of metal ignition. This is accomplished by the selection of materials and the generation of well-mixed gases. The combining of unique propellant injector element designs with secondary mixing devices is predicted to be the best approach.

  16. Operation of a Rotary-valved Pulse Detonation Rocket Engine Utilizing Liquid-kerosene and Oxygen

    Institute of Scientific and Technical Information of China (English)

    WANG Ke; FAN Wei; YAN Yu; ZHU Xudong; YAN Chuanjun

    2011-01-01

    The pulse detonation rocket engine (PDRE) requires periodic supply of oxidizer,fuel and purge gas.A rotary-valve assembly is fabricated to control the periodic supply in this research.Oxygen and liquid aviation kerosene are used as oxidizer and fuel respectively.An ordinary automobile spark plug,with ignition energy as low as 50 mJ,is used to initiate combustion.Steady operation of the PDRE is achieved with operating frequency ranging from 1 Hz to 10 Hz.Experimentally measured pressure is lower than theoretical value by 13% at 1 Hz and 37% at 10 Hz,and there also exists a velocity deficit at different operating frequencies.Both of these two phenomena are believed mainly due to droplet size which depends on atomization and vaporization of liquid fuel.

  17. Analysis of Flame Deflector Spray Nozzles in Rocket Engine Test Stands

    Science.gov (United States)

    Sachdev, Jai S.; Ahuja, Vineet; Hosangadi, Ashvin; Allgood, Daniel C.

    2010-01-01

    The development of a unified tightly coupled multi-phase computational framework is described for the analysis and design of cooling spray nozzle configurations on the flame deflector in rocket engine test stands. An Eulerian formulation is used to model the disperse phase and is coupled to the gas-phase equations through momentum and heat transfer as well as phase change. The phase change formulation is modeled according to a modified form of the Hertz-Knudsen equation. Various simple test cases are presented to verify the validity of the numerical framework. The ability of the methodology to accurately predict the temperature load on the flame deflector is demonstrated though application to an actual sub-scale test facility. The CFD simulation was able to reproduce the result of the test-firing, showing that the spray nozzle configuration provided insufficient amount of cooling.

  18. Characterization of typical platelet injector flow configurations. [liquid propellant rocket engines

    Science.gov (United States)

    Hickox, C. E.

    1975-01-01

    A study to investigate the hydraulic atomization characteristics of several novel injector designs for use in liquid propellant rocket engines is presented. The injectors were manufactured from a series of thin stainless steel platelets through which orifices were very accurately formed by a photoetching process. These individual platelets were stacked together and the orifices aligned so as to produce flow passages of prescribed geometry. After alignment, the platelets were bonded into a single, 'platelet injector', unit by a diffusion bonding process. Because of the complex nature of the flow associated with platelet injectors, it was necessary to use experimental techniques, exclusively, throughout the study. Large scale models of the injectors were constructed from aluminum plates and the appropriate fluids were modeled using a glycerol-water solution. Stop-action photographs of test configurations, using spark-shadowgraph or stroboscopic back-lighting, are shown.

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

    Science.gov (United States)

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

    2003-01-01

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

  20. A retrospective on early cryogenic primary rocket subsystem designs as integrated into rocket-based combined-cycle (RBCC) engines

    Science.gov (United States)

    Escher, William J. D.; Schnurstein, Robert E.

    1993-06-01

    A study (Escher and Flornes, 1966) of aerospace propulsion systems for a fully reusable earth-to-orbit space transport application that was performed in 1965-67 is reviewed. The present review provides a detailed, subject-focused technical retrospective on a key subsystem element of the rocket-based combined-cycle (RBCC) class of aerospace propulsion systems. The RBCC concept is considered to be a leading candidate propulsion approach for either SSTO or two-stage-to-orbit space transportaion applications.

  1. Experimental testing of a liquid bipropellant rocket engine using nitrous oxide and ethanol diluted with water

    Science.gov (United States)

    Phillip, Jeff; Morales, Rudy; Youngblood, Stewart; Saul, W. Venner; Grubelich, Mark; Hargather, Michael

    2016-11-01

    A research scale liquid bipropellant rocket engine testing facility was constructed at New Mexico Tech to perform research with various propellants. The facility uses a modular engine design that allows for variation of nozzle geometry and injector configurations. Initial testing focused on pure nitrous oxide and ethanol propellants, operating in the range of 5.5-6.9 MPa (800-1000 psi) chamber pressure with approximately 667 N (150 lbf) thrust. The system is instrumented with sensors for temperature, pressure, and thrust. Experimentally found values for specific impulse are in the range of 250-260 s which match computational predictions. Exhaust flow visualization is performed using high speed schlieren imaging. The engine startup and steady state exhaust flow features are studied through these videos. Computational and experimental data are presented for a study of dilution of the ethanol-nitrous oxide propellants with water. The study has shown a significant drop in chamber temperature compared to a small drop in specific impulse with increasing water dilution.

  2. Structural Analyses of the Support Trusses for the Nuclear Thermal Rocket Engines and Drop Tanks

    Science.gov (United States)

    Myers, David E.; Kosareo, Daniel N.

    2006-01-01

    Finite element structural analyses were performed on the support trusses of the Nuclear Thermal Rocket (NTR) engines and drop tanks to verify that the proper amount of mass was allocated for these components in the vehicle sizing model. The verification included a static stress analysis, a modal analysis, and a buckling analysis using the MSC/NASTRAN™ structural analysis software package. In addition, a crippling stress analysis was performed on the truss beams using a handbook equation. Two truss configurations were examined as possible candidates for the drop tanks truss while a baseline was examined for the engine support thrust structure. For the drop tanks trusses, results showed that both truss configurations produced similar results although one performed slightly better in buckling. In addition, it was shown that the mass allocated in the vehicle sizing model was adequate although the engine thrust structure may need to be modified slightly to increase its lateral natural frequency above the minimum requirement of 8 Hz that is specified in the Delta IV Payload Planners Guide.

  3. Integrated control and health management. Orbit transfer rocket engine technology program

    Science.gov (United States)

    Holzmann, Wilfried A.; Hayden, Warren R.

    1988-01-01

    To insure controllability of the baseline design for a 7500 pound thrust, 10:1 throttleable, dual expanded cycle, Hydrogen-Oxygen, orbit transfer rocket engine, an Integrated Controls and Health Monitoring concept was developed. This included: (1) Dynamic engine simulations using a TUTSIM derived computer code; (2) analysis of various control methods; (3) Failure Modes Analysis to identify critical sensors; (4) Survey of applicable sensors technology; and, (5) Study of Health Monitoring philosophies. The engine design was found to be controllable over the full throttling range by using 13 valves, including an oxygen turbine bypass valve to control mixture ratio, and a hydrogen turbine bypass valve, used in conjunction with the oxygen bypass to control thrust. Classic feedback control methods are proposed along with specific requirements for valves, sensors, and the controller. Expanding on the control system, a Health Monitoring system is proposed including suggested computing methods and the following recommended sensors: (1) Fiber optic and silicon bearing deflectometers; (2) Capacitive shaft displacement sensors; and (3) Hot spot thermocouple arrays. Further work is needed to refine and verify the dynamic simulations and control algorithms, to advance sensor capabilities, and to develop the Health Monitoring computational methods.

  4. 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.

  5. Flow visualization study in high aspect ratio cooling channels for rocket engines

    Science.gov (United States)

    Meyer, Michael L.; Giuliani, James E.

    1993-11-01

    The structural integrity of high pressure liquid propellant rocket engine thrust chambers is typically maintained through regenerative cooling. The coolant flows through passages formed either by constructing the chamber liner from tubes or by milling channels in a solid liner. Recently, Carlile and Quentmeyer showed life extending advantages (by lowering hot gas wall temperatures) of milling channels with larger height to width aspect ratios (AR is greater than 4) than the traditional, approximately square cross section, passages. Further, the total coolant pressure drop in the thrust chamber could also be reduced, resulting in lower turbomachinery power requirements. High aspect ratio cooling channels could offer many benefits to designers developing new high performance engines, such as the European Vulcain engine (which uses an aspect ratio up to 9). With platelet manufacturing technology, channel aspect ratios up to 15 could be formed offering potentially greater benefits. Some issues still exist with the high aspect ratio coolant channels. In a coolant passage of circular or square cross section, strong secondary vortices develop as the fluid passes through the curved throat region. These vortices mix the fluid and bring lower temperature coolant to the hot wall. Typically, the circulation enhances the heat transfer at the hot gas wall by about 40 percent over a straight channel. The effect that increasing channel aspect ratio has on the curvature heat transfer enhancement has not been sufficiently studied. If the increase in aspect ratio degrades the secondary flow, the fluid mixing will be reduced. Analysis has shown that reduced coolant mixing will result in significantly higher wall temperatures, due to thermal stratification in the coolant, thus decreasing the benefits of the high aspect ratio geometry. A better understanding of the fundamental flow phenomena in high aspect ratio channels with curvature is needed to fully evaluate the benefits of this

  6. Deposit formation in hydrocarbon rocket fuels

    Science.gov (United States)

    Roback, R.; Szetela, E. J.; Spadaccini, L. J.

    1981-01-01

    An experimental program was conducted to study deposit formation in hydrocarbon fuels under flow conditions that exist in high-pressure, rocket engine cooling systems. A high pressure fuel coking test apparatus was designed and developed and was used to evaluate thermal decomposition (coking) limits and carbon deposition rates in heated copper tubes for two hydrocarbon rocket fuels, RP-1 and commercial-grade propane. Tests were also conducted using JP-7 and chemically-pure propane as being representative of more refined cuts of the baseline fuels. A parametric evaluation of fuel thermal stability was performed at pressures of 136 atm to 340 atm, bulk fuel velocities in the range 6 to 30 m/sec, and tube wall temperatures in the range 422 to 811 K. Results indicated that substantial deposit formation occurs with RP-1 fuel at wall temperatures between 600 and 800 K, with peak deposit formation occurring near 700 K. No improvements were obtained when deoxygenated JP-7 fuel was substituted for RP-1. The carbon deposition rates for the propane fuels were generally higher than those obtained for either of the kerosene fuels at any given wall temperature. There appeared to be little difference between commercial-grade and chemically-pure propane with regard to type and quantity of deposit. Results of tests conducted with RP-1 indicated that the rate of deposit formation increased slightly with pressure over the range 136 atm to 340 atm. Finally, lating the inside wall of the tubes with nickel was found to significantly reduce carbon deposition rates for RP-1 fuel.

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

    Directory of Open Access Journals (Sweden)

    Jun Qian

    2014-04-01

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

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

    Science.gov (United States)

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

    2010-01-01

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

  9. Analysis of Flowfields over Four-Engine DC-X Rockets

    Science.gov (United States)

    Wang, Ten-See; Cornelison, Joni

    1996-01-01

    The objective of this study is to validate a computational methodology for the aerodynamic performance of an advanced conical launch vehicle configuration. The computational methodology is based on a three-dimensional, viscous flow, pressure-based computational fluid dynamics formulation. Both wind-tunnel and ascent flight-test data are used for validation. Emphasis is placed on multiple-engine power-on effects. Computational characterization of the base drag in the critical subsonic regime is the focus of the validation effort; until recently, almost no multiple-engine data existed for a conical launch vehicle configuration. Parametric studies using high-order difference schemes are performed for the cold-flow tests, whereas grid studies are conducted for the flight tests. The computed vehicle axial force coefficients, forebody, aftbody, and base surface pressures compare favorably with those of tests. The results demonstrate that with adequate grid density and proper distribution, a high-order difference scheme, finite rate afterburning kinetics to model the plume chemistry, and a suitable turbulence model to describe separated flows, plume/air mixing, and boundary layers, computational fluid dynamics is a tool that can be used to predict the low-speed aerodynamic performance for rocket design and operations.

  10. Liquid Rocket Engine Testing - Historical Lecture: Simulated Altitude Testing at AEDC

    Science.gov (United States)

    Dougherty, N. S.

    2010-01-01

    The span of history covered is from 1958 to the present. The outline of this lecture draws from historical examples of liquid propulsion testing done at AEDC primarily for NASA's Marshall Space Flight Center (NASA/MSFC) in the Saturn/Apollo Program and for USAF Space and Missile Systems dual-use customers. NASA has made dual use of Air Force launch vehicles, Test Ranges and Tracking Systems, and liquid rocket altitude test chambers / facilities. Examples are drawn from the Apollo/ Saturn vehicles and the testing of their liquid propulsion systems. Other examples are given to extend to the family of the current ELVs and Evolved ELVs (EELVs), in this case, primarily to their Upper Stages. The outline begins with tests of the XLR 99 Engine for the X-15 aircraft, tests for vehicle / engine induced environments during flight in the atmosphere and in Space, and vehicle staging at high altitude. The discussion is from the author's perspective and background in developmental testing.

  11. Microchennel development for autothermal reforming of hydrocarbon fuels

    Science.gov (United States)

    Bae, J.-M.; Ahmed, S.; Kumar, R.; Doss, E.

    Fuel-processing is a bridging technology to assist the commercialization of fuel cell systems in the absence of a hydrogen infrastructure. The Argonne National Laboratory has been developing fuel-processing technologies for fuel cells, and has reported the development of novel catalysts that are active and selective for hydrocarbon-reforming reactions. It has been realized, however, that with pellets or conventional honeycomb catalysts, the reforming process is mass-transport limited. This study addresses the development of catalysts structures with microchannels that are able to reduce the diffusion resistance and, thereby, achieve the same production rate within a smaller reactor bed. The microchannel reforming catalysts are prepared and tested with natural gas and gasoline-type fuels in a microreactor (diameter: 1 cm) at space velocities of up to 250 000 h -1. The catalysts have also been used in engineering-scale reactors (10 kWe; diameter: 7 cm) with similar product qualities. Compared with pellet catalysts, the microchannel catalysts offer a nearly five-fold reduction in catalyst weight and volume.

  12. Ceramic Microchannel Development for Compact Fuel Processors of Hydrocarbon Fuels

    Science.gov (United States)

    Bae, J.-M.; Ahmed, S.; Kumar, R.; Doss, E.

    Fuel processing is a bridging technology for faster commercialization of fuel cell system under lack of hydrogen infrastructures. Argonne national laboratory has been developing fuel processing technologies for fuel cell based electric power. We have reported the development of novel catalysts that are active and selective for hydrocarbon reforming reactions. It has been realized, however, that with pellet or conventional honeycomb catalysts, the reforming process is mass transport limited. This paper reports the development of catalyst structures with microchannels that are able to reduce the diffusion resistance and thereby achieve the same production rate within a smaller reactor bed. These microchannel reforming catalysts were prepared and tested with natural gas and gasoline-type fuels in a microreactor (1-cm dia.) at space velocities of up to 250,000 per hour. These catalysts have also been used in engineering-scale reactors (10 kWe, 7-cm dia.) with similar product qualities. Compared to pellet catalysts, the microchannel catalysts enable a nearly 5-fold reduction in catalyst weight and volume.

  13. On-board Optical Spectrometry for Detection of Mixture Ratio and Eroded Materials in Rocket Engine Exhaust Plume

    Science.gov (United States)

    Barkhoudarian, Sarkis; Kittinger, Scott

    2006-01-01

    Optical spectrometry can provide means to characterize rocket engine exhaust plume impurities due to eroded materials, as well as combustion mixture ratio without any interference with plume. Fiberoptic probes and cables were designed, fabricated and installed on Space Shuttle Main Engines (SSME), allowing monitoring of the plume spectra in real time with a Commercial of the Shelf (COTS) fiberoptic spectrometer, located in a test-stand control room. The probes and the cables survived the harsh engine environments for numerous hot-fire tests. When the plume was seeded with a nickel alloy powder, the spectrometer was able to successfully detect all the metallic and OH radical spectra from 300 to 800 nanometers.

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

    Energy Technology Data Exchange (ETDEWEB)

    Immich, H.; Alting, J.; Kretschmer, J.; Preclik, D. [Astrium GmbH, Space Infrastructure Div. Advanced Programs and System Engineering, Munich (Germany)

    2003-11-01

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

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

    Science.gov (United States)

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

    2000-01-01

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

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

    Science.gov (United States)

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

    2003-08-01

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

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

    Science.gov (United States)

    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.

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

    Science.gov (United States)

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

    2017-02-01

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

  19. A detailed numerical simulation of a liquid-propellant rocket engine ground test experiment

    Science.gov (United States)

    Lankford, D. W.; Simmons, M. A.; Heikkinen, B. D.

    1992-07-01

    A computational simulation of a Liquid Rocket Engine (LRE) ground test experiment was performed using two modeling approaches. The results of the models were compared with selected data to assess the validity of state-of-the-art computational tools for predicting the flowfield and radiative transfer in complex flow environments. The data used for comparison consisted of in-band station radiation measurements obtained in the near-field portion of the plume exhaust. The test article was a subscale LRE with an afterbody, resulting in a large base region. The flight conditions were such that afterburning regions were observed in the plume flowfield. A conventional standard modeling approach underpredicted the extent of afterburning and the associated radiation levels. These results were attributed to the absence of the base flow region which is not accounted for in this model. To assess the effects of the base region a Navier-Stokes model was applied. The results of this calculation indicate that the base recirculation effects are dominant features in the immediate expansion region and resulted in a much improved comparison. However, the downstream in-band station radiation data remained underpredicted by this model.

  20. Experimental study of a valveless pulse detonation rocket engine using nontoxic hypergolic propellants

    Science.gov (United States)

    Kan, Brandon K.

    A pulsed detonation rocket engine concept was explored through the use of hypergolic propellants in a fuel-centered pintle injector combustor. The combustor design yielded a simple open ended chamber with a pintle type injection element and pressure instrumentation. High-frequency pressure measurements from the first test series showed the presence of large pressure oscillations in excess of 2000 psia at frequencies between 400-600 hz during operation. High-speed video confirmed the high-frequency pulsed behavior and large amounts of after burning. Damaged hardware and instrumentation failure limited the amount of data gathered in the first test series, but the experiments met original test objectives of producing large over-pressures in an open chamber. A second test series proceeded by replacing hardware and instrumentation, and new data showed that pulsed events produced under expanded exhaust prior to pulsing, peak pressures around 8000 psi, and operating frequencies between 400-800 hz. Later hot-fires produced no pulsed behavior despite undamaged hardware. The research succeeded in producing pulsed combustion behavior using hypergolic fuels in a pintle injector setup and provided insights into design concepts that would assist future injector designs and experimental test setups.

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

    Science.gov (United States)

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

    2016-02-01

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

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

    Science.gov (United States)

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

    2017-07-01

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

  3. An Eight-Parameter Function for Simulating Model Rocket Engine Thrust Curves

    Science.gov (United States)

    Dooling, Thomas A.

    2007-01-01

    The toy model rocket is used extensively as an example of a realistic physical system. Teachers from grade school to the university level use them. Many teachers and students write computer programs to investigate rocket physics since the problem involves nonlinear functions related to air resistance and mass loss. This paper describes a nonlinear…

  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. Low-Cost High-Performance Non-Toxic Self-Pressurizing Storable Liquid Bi-Propellant Pressure-Fed Rocket Engine Project

    Data.gov (United States)

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

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

    Science.gov (United States)

    1984-01-01

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

  7. Improving Catalyst Efficiency in Bio-Based Hydrocarbon Fuels; NREL (National Renewable Energy Laboratory)

    Energy Technology Data Exchange (ETDEWEB)

    None

    2015-06-01

    This article investigates upgrading biomass pyrolysis vapors to form hydrocarbon fuels and chemicals using catalysts with different concentrations of acid sites. It shows that greater separation of acid sites makes catalysts more efficient at producing hydrocarbon fuels and chemicals. The conversion of biomass into liquid transportation fuels has attracted significant attention because of depleting fossil fuel reserves and environmental concerns resulting from the use of fossil fuels. Biomass is a renewable resource, which is abundant worldwide and can potentially be exploited to produce transportation fuels that are less damaging to the environment. This renewable resource consists of cellulose (40–50%), hemicellulose (25–35%), and lignin (16–33%) biopolymers in addition to smaller quantities of inorganic materials such as silica and alkali and alkaline earth metals (calcium and potassium). Fast pyrolysis is an attractive thermochemical technology for converting biomass into precursors for hydrocarbon fuels because it produces up to 75 wt% bio-oil,1 which can be upgraded to feedstocks and/or blendstocks for further refining to finished fuels. Bio-oil that has not been upgraded has limited applications because of the presence of oxygen-containing functional groups, derived from cellulose, hemicellulose and lignin, which gives rise to high acidity, high viscosity, low heating value, immiscibility with hydrocarbons and aging during storage. Ex situ catalytic vapor phase upgrading is a promising approach for improving the properties of bio-oil. The goal of this process is to reject oxygen and produce a bio-oil with improved properties for subsequent downstream conversion to hydrocarbons.

  8. Accounting for water formation from hydrocarbon fuel combustion in life cycle analyses

    Science.gov (United States)

    Belmont, E. L.; Davidson, F. T.; Glazer, Y. R.; Beagle, E. A.; Webber, M. E.

    2017-09-01

    Hydrocarbon fuel production and utilization are considered water intensive processes due to the high volumes of water used in source development and fuel processing. At the same time, there is significant water formed during combustion. However, this water is not currently widely harvested at the site of production. Instead, it is added to the hydrologic cycle, often in a different location from the fuel production site. This study quantifies the water formed from combustion of these fuels and analyzes the magnitudes of formation in the context of other hydrologic sources and sinks in order to facilitate future assessments of water harvesting technology and/or atmospheric impacts of combustion. Annual water formation from stoichiometric combustion of hydrocarbon fuels, including natural gas, oil- and natural gas liquid-derived products, and coal, in the United States and worldwide are presented and compared with quantities of water sequestered, evaporated, and stored in the atmosphere. Water production factors in terms of mass and energy of fuel consumed, WPFm and WPFe, respectively, are defined for the comparison of fuels and incorporation into future life cycle analyses (LCAs). Results show that water formation from combustion has increased worldwide from 2005 to 2015, with the largest increase coming from growth in combustion of natural gas. Water formation from combustion of hydrocarbon fuels equals or exceeds water sequestered from the hydrologic cycle through deep well injection in the US annually. Overall, water formation is deemed significant enough to warrant consideration by LCAs of water intensity in fuel production and use, and should be included in future analyses.

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

    Science.gov (United States)

    Brown, T. M.; Smith, Norm

    1999-01-01

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

  10. Nondestructive testing of rocket engine injector panel using ultrasonic burst phase thermography

    Science.gov (United States)

    Chen, Dapeng; Zhang, Cunlin; Wu, Naiming; Zeng, Zhi; Xing, Chunfei; Li, Yue; Zhao, Shibin; Ning, Tao

    2010-10-01

    As the key parts of the liquid rocket oxyhydrogen engine, the injector panel is a kind of transpiration material, which is braided and Sintered with stainless steel wire. If some hidden delaminition defects that are difficult to detect appear in the process of Sintering and rolling, a significant safety problem would occur. In this paper, we use the Ultrasonic Burst Phase Thermography (UBP) to detect the delamination defects in the injector panel, UBP is a rapid and reliable nondestructive technique derived from Ultrasonic Lock-in Thermography(ULT). It uses a controllable, adjustable ultrasonic burst as the heat source to stimulate the sample, the defects within the material are revealed through their heat generation caused by friction, clapping and thermoelastic effect, as the resulting surface temperature distribution is observed by an infrared camera. The original thermal images sequence is processed by Fast Fourier Transformation to obtain the phase information of the defects. In the experiments of the delamination sample, the UBP realized the selective heating of delamination defects in the injector panel, and the signal to noise of phase image is higher than the original thermal image because the phase information can not be disturbed by the initial conditions (such as the reflective surface of sample). However, the result of the detection of flat bottom hole transpiration panel sample reflects that UBP is not appropriate for the detection of this kind of defects, because it is difficult to induce frictional heating of flat bottom holes. As contrast, Flash Pulse Thermography is used to detect the flat bottom holes, all of the holes of different depth and sizes can be seen distinctly. The results show that PT is more appropriate for the detection of flat bottom holes defects than UBP, therefore, it is important to select the appropriate excitation method according to different defects.

  11. Effect of turbulence models on predicting convective heat transfer to hydrocarbon fuel at supercritical pressure

    Institute of Scientific and Technical Information of China (English)

    Tao Zhi; Cheng Zeyuan; Zhu Jianqin; Li Haiwang

    2016-01-01

    A variety of turbulence models were used to perform numerical simulations of heat transfer for hydrocarbon fuel flowing upward and downward through uniformly heated vertical pipes at supercritical pressure. Inlet temperatures varied from 373 K to 663 K, with heat flux rang-ing from 300 kW/m2 to 550 kW/m2. Comparative analyses between predicted and experimental results were used to evaluate the ability of turbulence models to respond to variable thermophys-ical properties of hydrocarbon fuel at supercritical pressure. It was found that the prediction per-formance of turbulence models is mainly determined by the damping function, which enables them to respond differently to local flow conditions. Although prediction accuracy for experimental results varied from condition to condition, the shear stress transport (SST) and launder and sharma models performed better than all other models used in the study. For very small buoyancy-influenced runs, the thermal-induced acceleration due to variations in density lead to the impairment of heat transfer occurring in the vicinity of pseudo-critical points, and heat transfer was enhanced at higher temperatures through the combined action of four thermophysical properties: density, viscosity, thermal conductivity and specific heat. For very large buoyancy-influenced runs, the thermal-induced acceleration effect was over predicted by the LS and AB models.

  12. Effect of turbulence models on predicting convective heat transfer to hydrocarbon fuel at supercritical pressure

    Directory of Open Access Journals (Sweden)

    Tao Zhi

    2016-10-01

    Full Text Available A variety of turbulence models were used to perform numerical simulations of heat transfer for hydrocarbon fuel flowing upward and downward through uniformly heated vertical pipes at supercritical pressure. Inlet temperatures varied from 373 K to 663 K, with heat flux ranging from 300 kW/m2 to 550 kW/m2. Comparative analyses between predicted and experimental results were used to evaluate the ability of turbulence models to respond to variable thermophysical properties of hydrocarbon fuel at supercritical pressure. It was found that the prediction performance of turbulence models is mainly determined by the damping function, which enables them to respond differently to local flow conditions. Although prediction accuracy for experimental results varied from condition to condition, the shear stress transport (SST and launder and sharma models performed better than all other models used in the study. For very small buoyancy-influenced runs, the thermal-induced acceleration due to variations in density lead to the impairment of heat transfer occurring in the vicinity of pseudo-critical points, and heat transfer was enhanced at higher temperatures through the combined action of four thermophysical properties: density, viscosity, thermal conductivity and specific heat. For very large buoyancy-influenced runs, the thermal-induced acceleration effect was over predicted by the LS and AB models.

  13. Energy and climate impacts of producing synthetic hydrocarbon fuels from CO(2).

    Science.gov (United States)

    van der Giesen, Coen; Kleijn, René; Kramer, Gert Jan

    2014-06-17

    Within the context of carbon dioxide (CO2) utilization there is an increasing interest in using CO2 as a resource to produce sustainable liquid hydrocarbon fuels. When these fuels are produced by solely using solar energy they are labeled as solar fuels. In the recent discourse on solar fuels intuitive arguments are used to support the prospects of these fuels. This paper takes a quantitative approach to investigate some of the claims made in this discussion. We analyze the life cycle performance of various classes of solar fuel processes using different primary energy and CO2 sources. We compare their efficacy with respect to carbon mitigation with ubiquitous fossil-based fuels and conclude that producing liquid hydrocarbon fuels starting from CO2 by using existing technologies requires much more energy than existing fuels. An improvement in life cycle CO2 emissions is only found when solar energy and atmospheric CO2 are used. Producing fuels from CO2 is a very long-term niche at best, not the panacea suggested in the recent public discourse.

  14. Biological Production of a Hydrocarbon Fuel Intermediate Polyhydroxybutyrate (Phb) from a Process Relevant Lignocellulosic Derived Sugar

    Energy Technology Data Exchange (ETDEWEB)

    Wang, Wei; Mohagheghi, Ali; Mittal, Ashutosh; Pilath, Heidi; Johnson, David K.

    2015-03-22

    PHAs are synthesized by many microorganisms to serve as intracellular carbon storage molecules. In some bacterial strains, PHB can account for up to 80% of cell mass. In addition to its application in the packaging sector, PHB also has great potential as an intermediate in the production of hydrocarbon fuels. PHB can be thermally depolymerized and decarboxylated to propene which can be upgraded to hydrocarbon fuels via commercial oligomerization technologies. In recent years a great effort has been made in bacterial production of PHB, yet the production cost of the polymer is still much higher than conventional petrochemical plastics. The high cost of PHB is because the cost of the substrates can account for as much as half of the total product cost in large scale fermentation. Thus searching for cheaper and better substrates is very necessary for PHB production. In this study, we demonstrate production of PHB by Cupriavidus necator from a process relevant lignocellulosic derived sugar stream, i.e., saccharified hydrolysate slurry from pretreated corn stover. Good cell growth was observed on slurry saccharified with advanced enzymes and 40~60% of PHB was accumulated in the cells. The mechanism of inhibition in the toxic hydrolysate generated by pretreatment and saccharification of biomass, will be discussed.

  15. Cu-Cr-Nb-Zr Alloy for Rocket Engines and Other High-Heat- Flux Applications

    Science.gov (United States)

    Ellis, David L.

    2013-01-01

    Rocket-engine main combustion chamber liners are used to contain the burning of fuel and oxidizer and provide a stream of high-velocity gas for propulsion. The liners in engines such as the Space Shuttle Main Engine are regeneratively cooled by flowing fuel, e.g., cryogenic hydrogen, through cooling channels in the back side of the liner. The heat gained by the liner from the flame and compression of the gas in the throat section is transferred to the fuel by the liner. As a result, the liner must either have a very high thermal conductivity or a very high operating temperature. In addition to the large heat flux (>10 MW/sq m), the liners experience a very large thermal gradient, typically more than 500 C over 1 mm. The gradient produces thermally induced stresses and strains that cause low cycle fatigue (LCF). Typically, a liner will experience a strain differential in excess of 1% between the cooling channel and the hot wall. Each time the engine is fired, the liner undergoes an LCF cycle. The number of cycles can be as few as one for an expendable booster engine, to as many as several thousand for a reusable launch vehicle or reaction control system. Finally, the liners undergo creep and a form of mechanical degradation called thermal ratcheting that results in the bowing out of the cooling channel into the combustion chamber, and eventual failure of the liner. GRCop-84, a Cu-Cr-Nb alloy, is generally recognized as the best liner material available at the time of this reporting. The alloy consists of 14% Cr2Nb precipitates in a pure copper matrix. Through experimental work, it has been established that the Zr will not participate in the formation of Laves phase precipitates with Cr and Nb, but will instead react with Cu to form the desired Cu-Zr compounds. It is believed that significant improvements in the mechanical properties of GRCop-84 will be realized by adding Zr. The innovation is a Cu-Cr-Nb-Zr alloy covering the composition range of 0.8 to 8.1 weight

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

    Science.gov (United States)

    Trushlyakov, V.; Shatrov, Ya.

    2017-09-01

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

  17. Deposit formation and heat transfer in hydrocarbon rocket fuels

    Science.gov (United States)

    Giovanetti, A. J.; Spadaccini, L. J.; Szetela, E. J.

    1983-01-01

    An experimental research program was undertaken to investigate the thermal stability and heat transfer characteristics of several hydrocarbon fuels under conditions that simulate high-pressure, rocket engine cooling systems. The rates of carbon deposition in heated copper and nickel-plated copper tubes were determined for RP-1, propane, and natural gas using a continuous flow test apparatus which permitted independent variation and evaluation of the effect on deposit formation of wall temperature, fuel pressure, and fuel velocity. In addition, the effects of fuel additives and contaminants, cryogenic fuel temperatures, and extended duration testing with intermittent operation were examined. Parametric tests to map the thermal stability characteristics of RP-1, commercial-grade propane, and natural gas were conducted at pressures of 6.9 to 13.8 MPa, bulk fuel velocities of 30 to 90 m/s, and tube wall temperatures in the range of 230 to 810 K. Also, tests were run in which propane and natural gas fuels were chilled to 230 and 160 K, respectively. Corrosion of the copper tube surface was detected for all fuels tested. Plating the inside of the copper tubes with nickel reduced deposit formation and eliminated tube corrosion in most cases. The lowest rates of carbon deposition were obtained for natural gas, and the highest rates were obtained for propane. For all fuels tested, the forced-convection heat transfer film coefficients were satisfactorily correlated using a Nusselt-Reynolds-Prandtl number equation.

  18. 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.

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

    Science.gov (United States)

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

    2001-01-01

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

  20. Techno-Economic Basis for Coproduct Manufacturing To Enable Hydrocarbon Fuel Production from Lignocellulosic Biomass

    Energy Technology Data Exchange (ETDEWEB)

    Biddy, Mary J.; Davis, Ryan; Humbird, David; Tao, Ling; Dowe, Nancy; Guarnieri, Michael T.; Linger, Jeffrey G.; Karp, Eric M.; Salvachua, Davinia; Vardon, Derek R.; Beckham, Gregg T.

    2016-06-06

    Biorefinery process development relies on techno-economic analysis (TEA) to identify primary cost drivers, prioritize research directions, and mitigate technical risk for scale-up through development of detailed process designs. Here, we conduct TEA of a model 2000 dry metric ton-per-day lignocellulosic biorefinery that employs a two-step pretreatment and enzymatic hydrolysis to produce biomass-derived sugars, followed by biological lipid production, lipid recovery, and catalytic hydrotreating to produce renewable diesel blendstock (RDB). On the basis of projected near-term technical feasibility of these steps, we predict that RDB could be produced at a minimum fuel selling price (MFSP) of USD $9.55/gasoline-gallon-equivalent (GGE), predicated on the need for improvements in the lipid productivity and yield beyond current benchmark performance. This cost is significant given the limitations in scale and high costs for aerobic cultivation of oleaginous microbes and subsequent lipid extraction/recovery. In light of this predicted cost, we developed an alternative pathway which demonstrates that RDB costs could be substantially reduced in the near term if upgradeable fractions of biomass, in this case hemicellulose-derived sugars, are diverted to coproducts of sufficient value and market size; here, we use succinic acid as an example coproduct. The coproduction model predicts an MFSP of USD $5.28/GGE when leaving conversion and yield parameters unchanged for the fuel production pathway, leading to a change in biorefinery RDB capacity from 24 to 15 MM GGE/year and 0.13 MM tons of succinic acid per year. Additional analysis demonstrates that beyond the near-term projections assumed in the models here, further reductions in the MFSP toward $2-3/GGE (which would be competitive with fossil-based hydrocarbon fuels) are possible with additional transformational improvements in the fuel and coproduct trains, especially in terms of carbon efficiency to both fuels and

  1. 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

  2. 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

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

    Directory of Open Access Journals (Sweden)

    N. S. Dorosh

    2014-01-01

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

  4. Starting of rocket engine at conditions of simulated altitude using crude monoethylaniline and other fuels with mixed acid

    Science.gov (United States)

    Ladanyi, Dezso J; Sloop, John L; Humphrey, Jack C; Morrell, Gerald

    1950-01-01

    Experiments were conducted at sea level and pressure altitude of about 55,000 feet at various temperatures to determine starting characteristics of a commercial rocket engine using crude monoethylaniline and other fuels with mixed acid. With crude monoethylaniline, ignition difficulties were encountered at temperatures below about 20 degrees F. With mixed butyl mercaptans, water-white turpentine, and x-pinene, no starting difficulties were experienced at temperatures as low as minus 74 degrees F. Turpentine and x-pinene, however, sometimes left deposits on the injector face. With blends containing furfuryl alcohol and with other blends, difficulties were experienced either from appreciable deposits or from starting.

  5. Rotating and positive-displacement pumps for low-thrust rocket engines. Volume 2: Fabrication and testing

    Science.gov (United States)

    Csomor, A.

    1974-01-01

    Rotating and positive displacement pumps of various types were studied for pumping liquid fluorine for low thrust high performance rocket engines. Included in the analysis were: centrifugal, pitot, Barske, Tesla, drag, gear, vane, axial piston, radial piston, diaphragm and helirotor pump concepts. The centrifugal and gear pumps were carried through detail design and fabrication. After preliminary testing in Freon 12, the centrifugal pump was selected for further testing and development. It was tested in Freon 12 to obtain the hydrodynamic performance. Tests were also conducted in liquid fluorine to demonstrate chemical compatibility.

  6. Sustainable hydrocarbon fuels by recycling CO2 and H2O with renewable or nuclear energy

    DEFF Research Database (Denmark)

    Graves, Christopher R.; Ebbesen, Sune; Mogensen, Mogens Bjerg

    2011-01-01

    . The dominant costs of the process are the electricity cost and the capital cost of the electrolyzer, and this capital cost is significantly increased when operating intermittently (on renewable power sources such as solar and wind). The potential of this CO2 recycling process is assessed, in terms of what......) and biofuels have received the most attention, similar hydrocarbons can be produced without using fossil fuels or biomass. Using renewable and/or nuclear energy, carbon dioxide and water can be recycled into liquid hydrocarbon fuels in non-biological processes which remove oxygen from CO2 and H2O (the reverse...... dissociation, and fuel synthesis. Dissociation methods include thermolysis, thermochemical cycles, electrolysis, and photoelectrolysis of CO2 and/or H2O. High temperature co-electrolysis of H2O and CO2 makes very efficient use of electricity and heat (near-100% electricity-to-syngas efficiency), provides high...

  7. Hydrocarbon fuel processing of micro solid oxide fuel cell systems[Dissertation 17455

    Energy Technology Data Exchange (ETDEWEB)

    Stutz, M. J.

    2007-07-01

    The scope of this thesis is the numerical and experimental investigation of the fuel processing of a micro solid oxide fuel cell (SOFC) running on hydrocarbon fuel. The goal is to enhance the overall system efficiency by optimization of the reforming process in the steady state and the improvement of the start-up process. Micro SOFC are a potential alternative to the currently used batteries in portable devices. Liquid butane in a cartridge could be the energy source. This dissertation is focused on the fuel processing of the system, namely the reforming and post-combusting processes. The reformer converts the hydrocarbon fuel to a hydrogen rich gas that can be utilized by the SOFC. The post-combustor depletes the toxic and/or explosive gases before leaving the exhaust. Chapter One presents a short introduction to the field of hydrocarbon fuel processing in micro solid oxide fuel cell systems, the next three chapters deal with computational modeling of the transport phenomena inside a micro-reformer, which leads to a better understanding of the chemistry and the physics therein, hence progress in the design and operation parameters. The experimental part (i.e. Chapter Five) of this thesis focuses on the feasibility of a novel hybrid start-up method of a fuel cell system that employs existing components as an additional heat source. In Chapter Two the effect of wall heat conduction on the syngas (hydrogen and carbon monoxide) production of a micro-reformer, representing micro-fabricated channels or monoliths, is investigated. Methane is used as a model hydrocarbon fuel since its heterogeneous reaction path on rhodium is known and validated. The simulations demonstrate that the axial wall conduction strongly influences the performance of the micro-reformer and should not be neglected without a careful a priori investigation of its impact. Methane conversion and hydrogen yield are strongly dependent of the wall inner surface temperature, which is influenced by the

  8. Investigation of bubble-point vapor pressures for mixtures of an endothermic hydrocarbon fuel with ethanol

    Energy Technology Data Exchange (ETDEWEB)

    Haiyun Sun; Wenjun Fang; Yongsheng Guo; Ruisen Lin [Zhejiang University, Hangzhou (China). Department of Chemistry, Molecular Thermodynamics

    2005-05-01

    Bubble-point vapor pressures and equilibrium temperatures for several mixtures with different mass fractions of a kerosene based endothermic hydrocarbon fuel (EHF) and ethanol were measured by comparative ebulliometry with inclined ebulliometers. Correlation between vapor pressures and equilibrium temperatures by the Antoine equation was given with satisfactory precision. The bubble-point lines of pressure versus composition at different temperatures and temperature versus composition at different pressures were obtained. The pseudo binary systems of EHF+ethanol appear with very large positive deviations from Raoult's law. It follows that the addition of ethanol had a critical effect on the vapor pressure of fuels. Ethanol may be an effective oxygenated hydrocarbon additive to adjust the volatility of EHF. 17 refs., 8 figs., 4 tabs.

  9. Laser Spectrometric Measurement System for Local Express Diagnostics of Flame at Combustion of Liquid Hydrocarbon Fuels

    Science.gov (United States)

    Kobtsev, V. D.; Kozlov, D. N.; Kostritsa, S. A.; Smirnov, V. V.; Stel'makh, O. M.; Tumanov, A. A.

    2016-03-01

    A laboratory laser spectrometric measurement system for investigation of spatial distributions of local temperatures in a flame at combustion of vapors of various liquid hydrocarbon fuels in oxygen or air at atmospheric pressure is presented. The system incorporates a coherent anti-Stokes Raman spectrometer with high spatial resolution for local thermometry of nitrogen-containing gas mixtures in a single laser shot and a continuous operation burner with a laminar diffusion flame. The system test results are presented for measurements of spatial distributions of local temperatures in various flame zones at combustion of vapor—gas n-decane/nitrogen mixtures in air. Its applicability for accomplishing practical tasks in comparative laboratory investigation of characteristics of various fuels and for research on combustion in turbulent flames is discussed.

  10. Hydrocarbon fuel processing of micro solid oxide fuel cell systems[Dissertation 17455

    Energy Technology Data Exchange (ETDEWEB)

    Stutz, M. J.

    2007-07-01

    The scope of this thesis is the numerical and experimental investigation of the fuel processing of a micro solid oxide fuel cell (SOFC) running on hydrocarbon fuel. The goal is to enhance the overall system efficiency by optimization of the reforming process in the steady state and the improvement of the start-up process. Micro SOFC are a potential alternative to the currently used batteries in portable devices. Liquid butane in a cartridge could be the energy source. This dissertation is focused on the fuel processing of the system, namely the reforming and post-combusting processes. The reformer converts the hydrocarbon fuel to a hydrogen rich gas that can be utilized by the SOFC. The post-combustor depletes the toxic and/or explosive gases before leaving the exhaust. Chapter One presents a short introduction to the field of hydrocarbon fuel processing in micro solid oxide fuel cell systems, the next three chapters deal with computational modeling of the transport phenomena inside a micro-reformer, which leads to a better understanding of the chemistry and the physics therein, hence progress in the design and operation parameters. The experimental part (i.e. Chapter Five) of this thesis focuses on the feasibility of a novel hybrid start-up method of a fuel cell system that employs existing components as an additional heat source. In Chapter Two the effect of wall heat conduction on the syngas (hydrogen and carbon monoxide) production of a micro-reformer, representing micro-fabricated channels or monoliths, is investigated. Methane is used as a model hydrocarbon fuel since its heterogeneous reaction path on rhodium is known and validated. The simulations demonstrate that the axial wall conduction strongly influences the performance of the micro-reformer and should not be neglected without a careful a priori investigation of its impact. Methane conversion and hydrogen yield are strongly dependent of the wall inner surface temperature, which is influenced by the

  11. Nano-structured noble metal catalysts based on hexametallate architecture for the reforming of hydrocarbon fuels

    Energy Technology Data Exchange (ETDEWEB)

    Gardner, Todd H.

    2015-09-15

    Nano-structured noble metal catalysts based on hexametallate lattices, of a spinel block type, and which are resistant to carbon deposition and metal sulfide formation are provided. The catalysts are designed for the reforming of hydrocarbon fuels to synthesis gas. The hexametallate lattices are doped with noble metals (Au, Pt, Rh, Ru) which are atomically dispersed as isolated sites throughout the lattice and take the place of hexametallate metal ions such as Cr, Ga, In, and/or Nb. Mirror cations in the crystal lattice are selected from alkali metals, alkaline earth metals, and the lanthanide metals, so as to reduce the acidity of the catalyst crystal lattice and enhance the desorption of carbon deposit forming moieties such as aromatics. The catalysts can be used at temperatures as high as 1000.degree. C. and pressures up to 30 atmospheres. A method for producing these catalysts and applications of their use also is provided.

  12. On-Line Measurement of Heat of Combustion of Gaseous Hydrocarbon Fuel Mixtures

    Science.gov (United States)

    Sprinkle, Danny R.; Chaturvedi, Sushil K.; Kheireddine, Ali

    1996-01-01

    A method for the on-line measurement of the heat of combustion of gaseous hydrocarbon fuel mixtures has been developed and tested. The method involves combustion of a test gas with a measured quantity of air to achieve a preset concentration of oxygen in the combustion products. This method involves using a controller which maintains the fuel (gas) volumetric flow rate at a level consistent with the desired oxygen concentration in the combustion products. The heat of combustion is determined form a known correlation with the fuel flow rate. An on-line computer accesses the fuel flow data and displays the heat of combustion measurement at desired time intervals. This technique appears to be especially applicable for measuring heats of combustion of hydrocarbon mixtures of unknown composition such as natural gas.

  13. 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

  14. Enhanced Large Solid Rocket Motor Understanding Through Performance Margin Testing: RSRM Five-Segment Engineering Test Motor (ETM-3)

    Science.gov (United States)

    Huppi, Hal; Tobias, Mark; Seiler, James

    2003-01-01

    The Five-Segment Engineering Test Motor (ETM-3) is an extended length reusable solid rocket motor (RSRM) intended to increase motor performance and internal environments above the current four-segment RSRM flight motor. The principal purpose of ETM-3 is to provide a test article for RSRM component margin testing. As the RSRM and Space Shuttle in general continue to age, replacing obsolete materials becomes an ever-increasing issue. Having a five-segment motor that provides environments in excess of normal opera- tion allows a mechanism to subject replacement materials to a more severe environment than experienced in flight. Additionally, ETM-3 offers a second design data point from which to develop and/or validate analytical models that currently have some level of empiricism associated with them. These enhanced models have the potential to further the understanding of RSRM motor performance and solid rocket motor (SRM) propulsion in general. Furthermore, these data could be leveraged to support a five-segment booster (FSB) development program should the Space Shuttle program choose to pursue this option for abort mode enhancements during the ascent phase. A tertiary goal of ETM-3 is to challenge both the ATK Thiokol Propulsion and NASA MSFC technical personnel through the design and analysis of a large solid rocket motor without the benefit of a well-established performance database such as the RSRM. The end result of this undertaking will be a more competent and experienced workforce for both organizations. Of particular interest are the motor design characteristics and the systems engineering approach used to conduct a complex yet successful large motor static test. These aspects of ETM-3 and more will be summarized.

  15. Deposit formation in hydrocarbon rocket fuels: Executive summary

    Science.gov (United States)

    Roback, R.; Szetela, E. J.; Spadaccini, L. J.

    1981-01-01

    An experimental program was conducted to study deposit formation in hydrocarbon fuels under flow conditions that exist in high-pressure, rocket engine cooling systems. A high pressure fuel coking test apparatus was designed and developed and was used to evaluate thermal decomposition (coking) limits and carbon deposition rates in heated copper tubes for two hydrocarbon rocket fuels, RP-1 and commercial-grade propane. Tests were also conducted using JP-7 and chemically-pure propane as being representative of more refined cuts of the baseline fuels. A parametric evaluation of fuel thermal stability was performed at pressures of 136 atm to 340 atm, bulk fuel velocities in the range 6 to 30 m/sec, and tube wall temperatures in the range 422 to 811K. In addition, the effect of the inside wall material on deposit formation was evaluated in selected tests which were conducted using nickel-plated tubes. The results of the tests indicated that substantial deposit formation occurs with RP-1 fuel at wall temperatures between 600 and 800K, with peak deposit formation occurring near 700K. No improvements were obtained when de-oxygenated JP-7 fuel was substituted for RP-1. The carbon deposition rates for the propane fuels were generally higher than those obtained for either of the kerosene fuels at any given wall temperature. There appeared to be little difference between commercial-grade and chemically-pure propane with regard to type and quantity of deposit. The results of tests conducted with RP-1 indicated that the rate of deposit formation increased slightly with pressure over the range 136 atm to 340 atm. Finally, plating the inside wall of the tubes with nickel was found to significantly reduce carbon deposition rates for RP-1 fuel.

  16. 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.

  17. Concept of a self-pressurized feed system for liquid rocket engines and its fundamental experiment results

    Science.gov (United States)

    Matsumoto, Jun; Okaya, Shunichi; Igoh, Hiroshi; Kawaguchi, Junichiro

    2017-04-01

    A new propellant feed system referred to as a self-pressurized feed system is proposed for liquid rocket engines. The self-pressurized feed system is a type of gas-pressure feed system; however, the pressurization source is retained in the liquid state to reduce tank volume. The liquid pressurization source is heated and gasified using heat exchange from the hot propellant using a regenerative cooling strategy. The liquid pressurization source is raised to critical pressure by a pressure booster referred to as a charger in order to avoid boiling and improve the heat exchange efficiency. The charger is driven by a part of the generated pressurization gas using a closed-loop self-pressurized feed system. The purpose of this study is to propose a propellant feed system that is lighter and simpler than traditional gas pressure feed systems. The proposed system can be applied to all liquid rocket engines that use the regenerative cooling strategy. The concept and mathematical models of the self-pressurized feed system are presented first. Experiment results for verification are then shown and compared with the mathematical models.

  18. 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

    NARloy-Z alloy (Cu-3 percent, Ag-0.5 percent, Zr) is a state of the art alloy currently used for fabricating rocket engine combustion chamber liners. Research conducted at NASA-MSFC and Penn State – Applied Research Laboratory has shown that thermal conductivity of NARloy-Z can be increased significantly by adding diamonds to form a composite (NARloy-Z-D). NARloy-Z-D is also lighter than NARloy-Z. These attributes make this advanced composite material an ideal candidate for fabricating combustion chamber liner for an advanced rocket engine. Increased thermal conductivity will directly translate into increased turbopump power and increased chamber pressure for improved thrust and specific impulse. This paper describes the process development for fabricating a subscale high thermal conductivity NARloy-Z-D combustion chamber liner using Field Assisted Sintering Technology (FAST). The FAST process uses a mixture of NARloy-Z and diamond powders which is sintered under pressure at elevated temperatures. Several challenges were encountered, i.e., segregation of diamonds, machining the super hard NARloy-Z-D composite, net shape fabrication and nondestructive examination. The paper describes how these challenges were addressed. Diamonds coated with copper (CuD) appear to give the best results. A near net shape subscale combustion chamber liner is being fabricated by diffusion bonding cylindrical rings of NARloy-Z-CuD using the FAST process.

  19. Large-eddy simulations of real-fluid effects in rocket engine combustors

    Science.gov (United States)

    Ma, Peter C.; Hickey, Jean-Pierre; Ihme, Matthias

    2013-11-01

    This study is concerned with the LES-modeling of real-fluid effects in rocket combustors. The non-ideal fluid behavior is modeled using the Peng-Robinson equation of state, and high-pressure effects on the thermo-viscous transport properties are also considered. An efficient and robust algorithm is developed to evaluate the thermodynamic state-vector. The highly non-linear coupling of the primitive thermodynamic variables in regions near the critical point requires special consideration to avoid spurious numerical oscillations. To avoid these non-physical oscillations, a second-order essentially non-oscillatory (ENO) scheme is applied in regions that are identified by a density-based sensor. The resulting algorithm is applied in LES to a coaxial rocket-injector, and super- and transcritical operating conditions are considered. Simulation results and comparisons with experimental data will be presented, and the influence of boundary conditions on the mixing characteristics will be discussed.

  20. Solar thermal rocket engine (STRE) thrust characteristics at the change of engine operation mode and of the flight vehicle attitude in the solar system

    Science.gov (United States)

    Kudrin, O. I.

    1993-10-01

    Relationships are presented which describe changes in the thrust and specific impulse of a solar thermal rocket engine due to a change in the flow rate of the working fluid (hydrogen). Expressions are also presented which describe the variation of the STRE thrust and specific impulse with the distance between the flight vehicle and the sun. Results of calculations are presented for an STRE with afterburning of the working fluid (hydrogen + oxygen) using hydrogen heating by solar energy to a temperature of 2360 K.

  1. Recycling Carbon Dioxide into Sustainable Hydrocarbon Fuels: Electrolysis of Carbon Dioxide and Water

    Science.gov (United States)

    Graves, Christopher Ronald

    Great quantities of hydrocarbon fuels will be needed for the foreseeable future, even if electricity based energy carriers begin to partially replace liquid hydrocarbons in the transportation sector. Fossil fuels and biomass are the most common feedstocks for production of hydrocarbon fuels. However, using renewable or nuclear energy, carbon dioxide and water can be recycled into sustainable hydrocarbon fuels in non-biological processes which remove oxygen from CO2 and H2O (the reverse of fuel combustion). Capture of CO2 from the atmosphere would enable a closed-loop carbon-neutral fuel cycle. The purpose of this work was to develop critical components of a system that recycles CO2 into liquid hydrocarbon fuels. The concept is examined at several scales, beginning with a broad scope analysis of large-scale sustainable energy systems and ultimately studying electrolysis of CO 2 and H2O in high temperature solid oxide cells as the heart of the energy conversion, in the form of three experimental studies. The contributions of these studies include discoveries about electrochemistry and materials that could significantly improve the overall energy use and economics of the CO2-to-fuels system. The broad scale study begins by assessing the sustainability and practicality of the various energy carriers that could replace petroleum-derived hydrocarbon fuels, including other hydrocarbons, hydrogen, and storage of electricity on-board vehicles in batteries, ultracapacitors, and flywheels. Any energy carrier can store the energy of any energy source. This sets the context for CO2 recycling -- sustainable energy sources like solar and wind power can be used to provide the most energy-dense, convenient fuels which can be readily used in the existing infrastructure. The many ways to recycle CO2 into hydrocarbons, based on thermolysis, thermochemical loops, electrolysis, and photoelectrolysis of CO2 and/or H 2O, are critically reviewed. A process based on high temperature co

  2. Rocket University at KSC

    Science.gov (United States)

    Sullivan, Steven J.

    2014-01-01

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

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

    Directory of Open Access Journals (Sweden)

    A. V. Aliev

    2015-01-01

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

  4. A Flash Vaporization System for Detonation of Hydrocarbon Fuels in a Pulse Detonation Engine

    Science.gov (United States)

    2005-06-01

    Gordon and Breach Science Publishers, Inc. Canada, 1975. 78. Material Safety Data Sheet,” Knovel Solvents - A Properties Database, ChemTec...Publishing, 2000. 79. Knovel Chemical Properties Handbook, McGraw-Hill, 1999. 80. Lin, K.C., Cox-Stouffer, S., Kennedy, P. J., and Jackson, T

  5. Ground and Space-Based Measurement of Rocket Engine Burns in the Ionosphere

    Science.gov (United States)

    Bernhardt, P. A.; Ballenthin, J. O.; Baumgardner, J. L.; Bhatt, A.; Boyd, I. D.; Burt, J. M.; Caton, R. G.; Coster, A.; Erickson, P. J.; Huba, J. D.; Earle, G. D.; Kaplan, C. R.; Foster, J. C.; Groves, K. M.; Haaser, R. A.; Heelis, R. A.; Hunton, D. E.; Hysell, D. L.; Klenzing, J. H.; Larsen, M. F.; Lind, F. D.; Pedersen, T. R.; Pfaff, R. F.; Stoneback, R. A.; Roddy, P. A.; Rodriguez, S. P.; San Antonio, G. S.; Schuck, P. W.; Siefring, C. L.; Selcher, C. A.; Smith, S. M.; Talaat, E. R.; Thomason, J. F.; Tsunoda, R. T.; Varney, R. H.

    2013-01-01

    On-orbit firings of both liquid and solid rocket motors provide localized disturbances to the plasma in the upper atmosphere. Large amounts of energy are deposited to ionosphere in the form of expanding exhaust vapors which change the composition and flow velocity. Charge exchange between the neutral exhaust molecules and the background ions (mainly O+) yields energetic ion beams. The rapidly moving pickup ions excite plasma instabilities and yield optical emissions after dissociative recombination with ambient electrons. Line-of-sight techniques for remote measurements rocket burn effects include direct observation of plume optical emissions with ground and satellite cameras, and plume scatter with UHF and higher frequency radars. Long range detection with HF radars is possible if the burns occur in the dense part of the ionosphere. The exhaust vapors initiate plasma turbulence in the ionosphere that can scatter HF radar waves launched from ground transmitters. Solid rocket motors provide particulates that become charged in the ionosphere and may excite dusty plasma instabilities. Hypersonic exhaust flow impacting the ionospheric plasma launches a low-frequency, electromagnetic pulse that is detectable using satellites with electric field booms. If the exhaust cloud itself passes over a satellite, in situ detectors measure increased ion-acoustic wave turbulence, enhanced neutral and plasma densities, elevated ion temperatures, and magnetic field perturbations. All of these techniques can be used for long range observations of plumes in the ionosphere. To demonstrate such long range measurements, several experiments were conducted by the Naval Research Laboratory including the Charged Aerosol Release Experiment, the Shuttle Ionospheric Modification with Pulsed Localized Exhaust experiments, and the Shuttle Exhaust Ionospheric Turbulence Experiments.

  6. Influence of hydrogen temperature on the stability of a rocket engine combustor operated with hydrogen and oxygen

    Science.gov (United States)

    Gröning, Stefan; Hardi, Justin; Suslov, Dmitry; Oschwald, Michael

    2017-03-01

    Since the late 1960s, low hydrogen injection temperature is known to have a destabilising effect on rocket engines with the propellant combination hydrogen/oxygen. Self-excited combustion instabilities of the first tangential mode have been found recently in a research rocket combustor operated with the propellant combination hydrogen/oxygen with a hydrogen temperature of 95 K. A hydrogen temperature ramping experiment has been performed with this research combustor to analyse the impact of hydrogen temperature on the self-excited combustion instabilities. The temperature was varied between 40 and 135 K. Contrary to past results found in literature, the combustor was found to be stable at low hydrogen temperatures while increased oscillation amplitudes of the first tangential mode were found at higher temperatures of around 100 K and above, which is consistent with previous observations of instabilities in this combustor. Further analysis shows that hydrogen temperature has a strong impact on the combustion chamber resonance frequencies. By varying the hydrogen injection temperature, the frequency of the first tangential mode is shifted to coincide with the second longitudinal resonance frequency of the liquid oxygen injector. Excitation of combustion chamber pressure oscillations was observed during such events.

  7. Design and evaluation of an oxidant-fuel-ratio-zoned rocket injector for high performance and ablative engine compatibility

    Science.gov (United States)

    Winter, J. M.; Pavli, A. J.; Shinn, A. M., Jr.

    1972-01-01

    A method for temperature control of the combustion gases in the peripheral zone of a rocket combustor which would reduce ablative throat erosion, prevent melting of zirconia throat inserts, and maintain high combustion performance is discussed. Included are techniques for analyzing and predicting zoned injector performance, as well as the philosophy and method for accomplishing an optimum compromise between high performance and reduced effective gas temperature. The experimental work was done with a 1000-lbf rocket engine which used as propellants N2O4 and a blend of 50-percent N2H4 and 50-percent UDMH at 100-psia chamber pressure and an overall O/F of 2.0. The method selected to provide temperature control was to use 30 percent of the propellant to form a peripheral zone of combustion gases at an O/F of 1.31 and 2700 K. The remaining 70 percent of the propellant in the core was at an O/F of 2.45 to keep the overall O/F at 2.0.

  8. Mixing characteristics of injector elements in liquid rocket engines - A computational study

    Science.gov (United States)

    Lohr, Jonathan C.; Trinh, Huu P.

    1992-01-01

    A computational study has been performed to better understand the mixing characteristics of liquid rocket injector elements. Variations in injector geometry as well as differences in injector element inlet flow conditions are among the areas examined in the study. Most results involve the nonreactive mixing of gaseous fuel with gaseous oxidizer but preliminary results are included that involve the spray combustion of oxidizer droplets. The purpose of the study is to numerically predict flowfield behavior in individual injector elements to a high degree of accuracy and in doing so to determine how various injector element properties affect the flow.

  9. An injector design model for predicting rocket engine performance and heat transfer

    Science.gov (United States)

    Calhoon, D. F.; Kors, D. L.; Gordon, L. H.

    1973-01-01

    A model is formulated for estimating the performance and chamber heat transfer in rocket injectors/chambers operating with gaseous H2-O2 propellants. The model quantifies the combustion performance and chamber heat flux for variables such as chamber length, element type, element area ratio, impingement angle, thrust/element, mixture ratio, moment ratio, element spacing, and physical size. Design equations are given and curves are plotted for evaluation of combustion performance in injectors comprised of F-O-F triplet, premix, coaxial and swirl coaxial element types. Curve plots and equations are also included for estimation of the chamber wall heat fluxes generated by these element types.

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

    Directory of Open Access Journals (Sweden)

    Trushlyakov Valeriy I.

    2017-01-01

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

  11. Physiological tolerance and stoichiometric potential of cyanobacteria for hydrocarbon fuel production.

    Science.gov (United States)

    Kämäräinen, Jari; Knoop, Henning; Stanford, Natalie J; Guerrero, Fernando; Akhtar, M Kalim; Aro, Eva-Mari; Steuer, Ralf; Jones, Patrik R

    2012-11-30

    Cyanobacteria are capable of directly converting sunlight, carbon dioxide and water into hydrocarbon fuel or precursors thereof. Many biological and non-biological factors will influence the ability of such a production system to become economically sustainable. We evaluated two factors in engineerable cyanobacteria which could potentially limit economic sustainability: (i) tolerance of the host to the intended end-product, and (ii) stoichiometric potential for production. Alcohols, when externally added, inhibited growth the most, followed by aldehydes and acids, whilst alkanes were the least inhibitory. The growth inhibition became progressively greater with increasing chain-length for alcohols, whilst the intermediate C6 alkane caused more inhibition than both C3 and C11 alkane. Synechocystis sp. PCC 6803 was more tolerant to some of the tested chemicals than Synechococcus elongatus PCC 7942, particularly ethanol and undecane. Stoichiometric evaluation of the potential yields suggested that there is no difference in the potential productivity of harvestable energy between any of the studied fuels, with the exception of ethylene, for which maximal stoichiometric yield is considerably lower. In summary, it was concluded that alkanes would constitute the best choice metabolic end-product for fuel production using cyanobacteria if high-yielding strains can be developed. Copyright © 2012 Elsevier B.V. All rights reserved.

  12. Uncertainty analysis of steady state incident heat flux measurements in hydrocarbon fuel fires.

    Energy Technology Data Exchange (ETDEWEB)

    Nakos, James Thomas

    2005-12-01

    The objective of this report is to develop uncertainty estimates for three heat flux measurement techniques used for the measurement of incident heat flux in a combined radiative and convective environment. This is related to the measurement of heat flux to objects placed inside hydrocarbon fuel (diesel, JP-8 jet fuel) fires, which is very difficult to make accurately (e.g., less than 10%). Three methods will be discussed: a Schmidt-Boelter heat flux gage; a calorimeter and inverse heat conduction method; and a thin plate and energy balance method. Steady state uncertainties were estimated for two types of fires (i.e., calm wind and high winds) at three times (early in the fire, late in the fire, and at an intermediate time). Results showed a large uncertainty for all three methods. Typical uncertainties for a Schmidt-Boelter gage ranged from {+-}23% for high wind fires to {+-}39% for low wind fires. For the calorimeter/inverse method the uncertainties were {+-}25% to {+-}40%. The thin plate/energy balance method the uncertainties ranged from {+-}21% to {+-}42%. The 23-39% uncertainties for the Schmidt-Boelter gage are much larger than the quoted uncertainty for a radiative only environment (i.e ., {+-}3%). This large difference is due to the convective contribution and because the gage sensitivities to radiative and convective environments are not equal. All these values are larger than desired, which suggests the need for improvements in heat flux measurements in fires.

  13. Microwave-Assisted Decarboxylation of Sodium Oleate and Renewable Hydrocarbon Fuel Production

    Institute of Scientific and Technical Information of China (English)

    Wang Yunpu; Liu Yuhuan; Ruan Rongsheng; Wen Pingwei; Wan Yiqin; Zhang Jinsheng

    2013-01-01

    The carboxyl terminal of sodium oleate has a stronger polarity than that of oleic acid;this terminal is more likely to be dipole polarized and ionically conductive in a microwave ifeld. Sodium oleate was used as the model compound to study the decarboxylation of oleic acid leading to hydrocarbon formation via microwave-assisted pyrolysis technology. The pyrolysis gas, liquid, and solid products were precisely analyzed to deduce the mechanism for decarboxylation of sodium oleate. Microwave energy was able to selectively heat the carboxyl terminal of sodium oleate. During decarboxylation, the double bond in the long hydrocarbon chain formed a p-πconjugated system with the carbanion intermediate. The resulting p-πconjugated system was more stable and beneifcial to the pyrolysis reaction (decarboxylation, terminal allylation, isomeriza-tion, and aromatization). The physical properties of pyrolysis liquid were generally similar to those of diesel fuel, thereby demonstrating the possible use of microwaves for controlling the decarboxylation of sodium oleate in order to manufacture renewable hydrocarbon fuels.

  14. Integration of direct carbon and hydrogen fuel cells for highly efficient power generation from hydrocarbon fuels

    Energy Technology Data Exchange (ETDEWEB)

    Muradov, Nazim; Choi, Pyoungho; Smith, Franklyn; Bokerman, Gary [Florida Solar Energy Center, University of Central Florida, 1679 Clearlake Road, Cocoa, FL 32922-5703 (United States)

    2010-02-15

    In view of impending depletion of hydrocarbon fuel resources and their negative environmental impact, it is imperative to significantly increase the energy conversion efficiency of hydrocarbon-based power generation systems. The combination of a hydrocarbon decomposition reactor with a direct carbon and hydrogen fuel cells (FC) as a means for a significant increase in chemical-to-electrical energy conversion efficiency is discussed in this paper. The data on development and operation of a thermocatalytic hydrocarbon decomposition reactor and its coupling with a proton exchange membrane FC are presented. The analysis of the integrated power generating system including a hydrocarbon decomposition reactor, direct carbon and hydrogen FC using natural gas and propane as fuels is conducted. It was estimated that overall chemical-to-electrical energy conversion efficiency of the integrated system varied in the range of 49.4-82.5%, depending on the type of fuel and FC used, and CO{sub 2} emission per kW{sub el}h produced is less than half of that from conventional power generation sources. (author)

  15. Chemical storage of renewable electricity in hydrocarbon fuels via H{sub 2}

    Energy Technology Data Exchange (ETDEWEB)

    Eilers, H.; Iglesias Gonzalez, M.; Schaub, G. [Karlsruhe Institute of Technology (KIT), Karlsruhe (Germany). Engler-Bunte-Institute I

    2012-07-01

    The increased generation of renewable electricity leads to an increasing demand for storage due to its fluctuating production. Electrical energy can be stored as chemical energy carriers e.g. in form of H{sub 2} that can be further processed to hydrocarbons. Storage in form of hydrocarbons is advantageous compared to H{sub 2} storage since (i) a higher volumetric energy density in the product can be achieved and (ii) the infrastructure for hydrocarbon distribution, storage and utilization already exists. The present contribution introduces the potential of H{sub 2} integration in upgrading/production processes to hydrocarbon fuels, based on stoichiometry and kind of carbon feedstock. Processes include petroleum refining, vegetable oil hydrogenation, production of synfuel from lignocellulosic biomass and substitute natural gas from H{sub 2}/CO{sub 2}. In the case of fossil raw materials, yields per feedstock can be increased and fossil CO{sub 2} emissions decreased since fossil resources for H{sub 2} production can be avoided. In the case of biomass conversion to synfuels, product yields per biomass/hectare can be increased. If CO{sub 2} is hydrogenated to fuels, no gasification step is needed, however lower hydrocarbon product yields per H{sub 2} are achieved since CO{sub 2} has the highest oxygen content. (orig.)

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

    Directory of Open Access Journals (Sweden)

    Tarakanova S.Y.

    2014-12-01

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

  17. A methodology to study the possible occurrence of chugging in liquid rocket engines during transient start-up

    Science.gov (United States)

    Leonardi, Marco; Nasuti, Francesco; Di Matteo, Francesco; Steelant, Johan

    2017-10-01

    An investigation on the low frequency combustion instabilities due to the interaction of combustion chamber and feed line dynamics in a liquid rocket engine is carried out implementing a specific module in the system analysis software EcosimPro. The properties of the selected double time lag model are identified according to the two classical assumptions of constant and variable time lag. Module capabilities are evaluated on a literature experimental set up consisting of a combustion chamber decoupled from the upstream feed lines. The computed stability map results to be in good agreement with both experimental data and analytical models. Moreover, the first characteristic frequency of the engine is correctly predicted, giving confidence on the use of the module for the analysis of chugging instabilities. As an example of application, a study is carried out on the influence of the feed lines on the system stability, correctly capturing that the lines extend the stable regime of the combustion chamber and that the propellant domes play a key role in coupling the dynamics of combustion chamber and feed lines. A further example is presented to discuss on the role of pressure growth rate and of the combustion chamber properties on the possible occurrence of chug instability during engine start-up and on the conditions that lead to its damping or growth.

  18. Condensation of water vapor and carbon dioxide in the jet exhausts of rocket engines: 1. Model calculation of the physical conditions in a jet exhaust

    Science.gov (United States)

    Platov, Yu. V.; Alpatov, V. V.; Klyushnikov, V. Yu.

    2014-01-01

    Model calculations have been performed for the temperature and pressure of combustion products in the jet exhaust of rocket engines of last stages of Proton, Molniya, and Start launchers operating in the upper atmosphere at altitudes above 120 km. It has been shown that the condensation of water vapor and carbon dioxide can begin at distances of 100-150 and 450-650 m away from the engine nozzle, respectively.

  19. Development of the Functional Flow Block Diagram for the J-2X Rocket Engine System

    Science.gov (United States)

    White, Thomas; Stoller, Sandra L.; Greene, WIlliam D.; Christenson, Rick L.; Bowen, Barry C.

    2007-01-01

    The J-2X program calls for the upgrade of the Apollo-era Rocketdyne J-2 engine to higher power levels, using new materials and manufacturing techniques, and with more restrictive safety and reliability requirements than prior human-rated engines in NASA history. Such requirements demand a comprehensive systems engineering effort to ensure success. Pratt & Whitney Rocketdyne system engineers performed a functional analysis of the engine to establish the functional architecture. J-2X functions were captured in six major operational blocks. Each block was divided into sub-blocks or states. In each sub-block, functions necessary to perform each state were determined. A functional engine schematic consistent with the fidelity of the system model was defined for this analysis. The blocks, sub-blocks, and functions were sequentially numbered to differentiate the states in which the function were performed and to indicate the sequence of events. The Engine System was functionally partitioned, to provide separate and unique functional operators. Establishing unique functional operators as work output of the System Architecture process is novel in Liquid Propulsion Engine design. Each functional operator was described such that its unique functionality was identified. The decomposed functions were then allocated to the functional operators both of which were the inputs to the subsystem or component performance specifications. PWR also used a novel approach to identify and map the engine functional requirements to customer-specified functions. The final result was a comprehensive Functional Flow Block Diagram (FFBD) for the J-2X Engine System, decomposed to the component level and mapped to all functional requirements. This FFBD greatly facilitates component specification development, providing a well-defined trade space for functional trades at the subsystem and component level. It also provides a framework for function-based failure modes and effects analysis (FMEA), and a

  20. Development of the Functional Flow Block Diagram for the J-2X Rocket Engine System

    Science.gov (United States)

    White, Thomas; Stoller, Sandra L.; Greene, WIlliam D.; Christenson, Rick L.; Bowen, Barry C.

    2007-01-01

    The J-2X program calls for the upgrade of the Apollo-era Rocketdyne J-2 engine to higher power levels, using new materials and manufacturing techniques, and with more restrictive safety and reliability requirements than prior human-rated engines in NASA history. Such requirements demand a comprehensive systems engineering effort to ensure success. Pratt & Whitney Rocketdyne system engineers performed a functional analysis of the engine to establish the functional architecture. J-2X functions were captured in six major operational blocks. Each block was divided into sub-blocks or states. In each sub-block, functions necessary to perform each state were determined. A functional engine schematic consistent with the fidelity of the system model was defined for this analysis. The blocks, sub-blocks, and functions were sequentially numbered to differentiate the states in which the function were performed and to indicate the sequence of events. The Engine System was functionally partitioned, to provide separate and unique functional operators. Establishing unique functional operators as work output of the System Architecture process is novel in Liquid Propulsion Engine design. Each functional operator was described such that its unique functionality was identified. The decomposed functions were then allocated to the functional operators both of which were the inputs to the subsystem or component performance specifications. PWR also used a novel approach to identify and map the engine functional requirements to customer-specified functions. The final result was a comprehensive Functional Flow Block Diagram (FFBD) for the J-2X Engine System, decomposed to the component level and mapped to all functional requirements. This FFBD greatly facilitates component specification development, providing a well-defined trade space for functional trades at the subsystem and component level. It also provides a framework for function-based failure modes and effects analysis (FMEA), and a

  1. A unified Navier-Stokes flowfield and performance analysis of liquid rocket engines

    Science.gov (United States)

    Wang, Ten-See; Chen, Yen-Sen

    1990-07-01

    To improve the current composite solutions in the design and analysis of liquid propulsive engines, a computational fluid dynamics model capable of calculating the nonreacting and reacting flows from the combustion chamber, through the nozzle to the external plume, was developed. The Space Shuttle Main Engine (SSME) fired at sea level, along with the flowfields of several other nozzles were investigated. The bell-shaped SSME nozzle was run at 100 percent power level at various flow conditions, the computed flow results and performance compared well with those of other standard codes and engine hot fire test data.

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

    Science.gov (United States)

    1972-01-01

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

  3. MMH/NTO火箭发动机燃烧动态稳定性数值评定%Numerical assessment of MMH/NTO rocket engine combustion instability

    Institute of Scientific and Technical Information of China (English)

    庄逢辰; 聂万胜; 邹勤; 张中光

    2001-01-01

    应用脉冲枪不稳定燃烧模型对有/无声腔的MMH/NTO火箭发动机的燃烧稳定性进行了数值模拟,比较了3台MMH/NTO发动机的燃烧动态稳定性,计算与发动机的试车结果一致。%A given MMH/NTO rocket engine combustion stability with/without acoustic cavities was numerically simulated by pulse gun combustion instability model.Three MMH/NTO rocket engines combustion dynamic stabilities were compared and assessed. Numerical simulation and assessment results are agreeable with the engine hot test data.

  4. Condensation of water vapor and carbon dioxide in the jet exhausts of rocket engines: 1. Heterogeneous condensation of combustion products

    Science.gov (United States)

    Platov, Yu. V.; Semenov, A. I.; Filippov, B. V.

    2014-01-01

    Condensation of water vapor and carbon dioxide in the jet exhausts of rocket engines during last stages of Proton, Molniya, and Start launchers operating in the upper atmospheric with different types of fuels is considered. Particle heating is taken into account with emission of latent heat of condensation and energy loss due to radiation and heat exchange with combustion products. Using the solution of the heat balance and condensed particle mass equations, the temporal change in the temperature and thickness of the condensate layer is obtained. Practically, no condensation of water vapor and carbon dioxide in the jet exhaust of a Start launcher occurs. In plumes of Proton and Molniya launchers, the condensation of water vapor and carbon dioxide can start at distances of 120-170 m and 450-650 m from the engine nozzle, respectively. In the course of condensation, the thickness of the "water" layer on particles can exceed 100 Å, and the thickness of carbon dioxide can exceed 60 Å.

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

    Science.gov (United States)

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

    2014-01-01

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

  6. OPERATION OF SOLID OXIDE FUEL CELL ANODES WITH PRACTICAL HYDROCARBON FUELS

    Energy Technology Data Exchange (ETDEWEB)

    Scott A. Barnett; Jiang Liu; Yuanbo Lin

    2004-07-30

    This work was carried out to achieve a better understanding of how SOFC anodes work with real fuels. The motivation was to improve the fuel flexibility of SOFC anodes, thereby allowing simplification and cost reduction of SOFC power plants. The work was based on prior results indicating that Ni-YSZ anode-supported SOFCs can be operated directly on methane and natural gas, while SOFCs with novel anode compositions can work with higher hydrocarbons. While these results were promising, more work was clearly needed to establish the feasibility of these direct-hydrocarbon SOFCs. Basic information on hydrocarbon-anode reactions should be broadly useful because reformate fuel gas can contain residual hydrocarbons, especially methane. In the Phase I project, we have studied the reaction mechanisms of various hydrocarbons--including methane, natural gas, and higher hydrocarbons--on two kinds of Ni-containing anodes: conventional Ni-YSZ anodes and a novel ceramic-based anode composition that avoid problems with coking. The effect of sulfur impurities was also studied. The program was aimed both at achieving an understanding of the interactions between real fuels and SOFC anodes, and providing enough information to establish the feasibility of operating SOFC stacks directly on hydrocarbon fuels. A combination of techniques was used to provide insight into the hydrocarbon reactions at these anodes during SOFC operation. Differentially-pumped mass spectrometry was be used for product-gas analysis both with and without cell operation. Impedance spectroscopy was used in order to understand electrochemical rate-limiting steps. Open-circuit voltages measurements under a range of conditions was used to help determine anode electrochemical reactions. Life tests over a wide range of conditions were used to establish the conditions for stable operation of anode-supported SOFC stacks directly on methane. Redox cycling was carried out on ceramic-based anodes. Tests on sulfur tolerance of

  7. The development of a solid-state hydrogen sensor for rocket engine leakage detection

    Science.gov (United States)

    Liu, Chung-Chiun

    Hydrogen propellant leakage poses significant operational problems in the rocket propulsion industry as well as for space exploratory applications. Vigorous efforts have been devoted to minimizing hydrogen leakage in assembly, test, and launch operations related to hydrogen propellant. The objective has been to reduce the operational cost of assembling and maintaining hydrogen delivery systems. Specifically, efforts have been made to develop a hydrogen leak detection system for point-contact measurement. Under the auspices of Lewis Research Center, the Electronics Design Center at Case Western Reserve University, Cleveland, Ohio, has undertaken the development of a point-contact hydrogen gas sensor with potential applications to the hydrogen propellant industry. We envision a sensor array consisting of numbers of discrete hydrogen sensors that can be located in potential leak sites. Silicon-based microfabrication and micromachining techniques are used in the fabrication of these sensor prototypes. Evaluations of the sensor are carried out in-house at Case Western Reserve University as well as at Lewis Research Center and GenCorp Aerojet, Sacramento, California. The hydrogen gas sensor is not only applicable in a hydrogen propulsion system, but also usable in many other civilian and industrial settings. This includes vehicles or facility use, or in the production of hydrogen gas. Dual space and commercial uses of these point-contacted hydrogen sensors are feasible and will directly meet the needs and objectives of NASA as well as various industrial segments.

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

    Science.gov (United States)

    Ross, M. N.; Toohey, D.

    2008-12-01

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

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

    Science.gov (United States)

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

    2008-01-01

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

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

    Science.gov (United States)

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

    2008-01-01

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

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

    Science.gov (United States)

    Valenti, Elizabeth L.

    1992-10-01

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

  12. Thermodynamic study of characteristics of the converter with separated supply of hydrocarbon fuel for thermo-oxidative and steam reforming

    Science.gov (United States)

    Bassina, I. A.; Malkov, Yu. P.; Molchanov, O. N.; Stepanov, S. G.; Troshchinenko, G. A.; Zasypkin, I. M.

    2014-04-01

    Thermodynamic studies of the converter characteristics were performed to produce hydrogen-containing syngas from hydrocarbon fuel (kerosene) with its separated supply for thermo-oxidative and steam reforming. It is demonstrated that the optimal conditions of the converter performance correlate with the oxidant ratio of α > 0.5 at the heattransfer wall temperature of 1200 K. Hydrogen content in the final syngas reaches 60 % by volume, free carbon (soot) deposition in reforming products is excluded, and there is no need to apply walls water cooling in the converter.

  13. Efficiency Analysis of Technological Methods for Reduction of NOx Emissions while Burning Hydrocarbon Fuels in Heat and Power Plants

    Directory of Open Access Journals (Sweden)

    S. Kabishov

    2013-01-01

    Full Text Available The paper contains a comparative efficiency analysis pertaining to application of existing technological methods for suppression of nitric oxide formation in heating boilers of heat generators. A special attention has been given to investigation of NOx  emission reduction while burning hydrocarbon fuel with the help of oxygen-enriched air. The calculations have demonstrated that while enriching oxidizer with the help of oxygen up to 50 % (by volume it is possible to reduce volume of NOx formation (while burning fuel unit by 21 %.

  14. 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.

  15. Large eddy simulation of combustion characteristics in a kerosene fueled rocket-based combined-cycle engine combustor

    Science.gov (United States)

    Huang, Zhi-wei; He, Guo-qiang; Qin, Fei; Cao, Dong-gang; Wei, Xiang-geng; Shi, Lei

    2016-10-01

    This study reports combustion characteristics of a rocket-based combined-cycle engine combustor operating at ramjet mode numerically. Compressible large eddy simulation with liquid kerosene sprayed and vaporized is used to study the intrinsic unsteadiness of combustion in such a propulsion system. Results for the pressure oscillation amplitude and frequency in the combustor as well as the wall pressure distribution along the flow-path, are validated using experimental data, and they show acceptable agreement. Coupled with reduced chemical kinetics of kerosene, results are compared with the simultaneously obtained Reynolds-Averaged Navier-Stokes results, and show significant differences. A flow field analysis is also carried out for further study of the turbulent flame structures. Mixture fraction is used to determine the most probable flame location in the combustor at stoichiometric condition. Spatial distributions of the Takeno flame index, scalar dissipation rate, and heat release rate reveal that different combustion modes, such as premixed and non-premixed modes, coexisted at different sections of the combustor. The RBCC combustor is divided into different regions characterized by their non-uniform features. Flame stabilization mechanism, i.e., flame propagation or fuel auto-ignition, and their relative importance, is also determined at different regions in the combustor.

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

    Science.gov (United States)

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

    2015-06-01

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

  17. On the use of a three-dimensional Navier-Stokes solver for rocket engine pump impeller design

    Science.gov (United States)

    Chen, Wei-Chung; Prueger, George H.; Chan, Daniel C.; Eastland, Anthony H.

    1992-07-01

    A 3D Reynolds-averaged Navier-Stokes Solver and a Fast Grid Generator (FGG), developed specially for centrifugal impeller design, were incorporated into the pump impeller design process. The impeller performance from the CFD analysis was compared to one-dimensional prediction. Both analyses showed good agreement of the impeller hydraulic efficiency, 94.5 percent, but with an 8 percent discrepancy of Euler head prediction. The impeller blade angle, discharge hub to shroud width, axial length and blade stacking were systematically changed to achieve an optimum impeller design. Impeller overall efficiency, loss distribution, hub-to-tip flow angle distortion and blade-to-blade flow angle change are among those criteria used to evaluate impeller performance. Two grid sizes, one with 10 K grid points and one with 80 K grid points were used to evaluate grid dependency issues. The effects of grid resolution on the accuracy and turnaround time are discussed. In conclusion, it is demonstrated that CFD can be effectively used for design and optimization of rocket engine pump components.

  18. 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)

  19. Reynolds-averaged Navier-Stokes analysis of the flow through a model rocket-based combined-cycle engine with an independently-fueled ramjet stream

    Science.gov (United States)

    Bond, Ryan Bomar

    A new concept for the low speed propulsion mode in rocket based combined cycle (RBCC) engines has been developed as part of the NASA GTX program. This concept, called the independent ramjet stream (IRS) cycle, is a variation of the traditional ejector ramjet (ER) design and involves the injection of hydrogen fuel directly into the air stream, where it is ignited by the rocket plume. Experiments and computational fluid dynamics (CFD) are currently being used to evaluate the feasibility of the new design. In this work, a Navier-Stokes code valid for general reactive flows is applied to the model engine under cold flow, ejector ramjet, and IRS cycle operation. Pressure distributions corresponding to cold-flow and ejector ramjet operation are compared with experimental data. The engine response under independent ramjet stream cycle operation is examined for different reaction models and grid sizes. The engine response to variations in fuel injection is also examined. Mode transition simulations are also analyzed both with and without a nitrogen purge of the rocket. The solutions exhibit a high sensitivity to both grid resolution and reaction mechanism, but they do indicate that thermal throat ramjet operation is possible through the injection and burning of additional fuel into the air stream. The solutions also indicate that variations in fuel injection location can affect the position of the thermal throat. The numerical simulations predicted successful mode transition both with and without a nitrogen purge of the rocket; however, the reliability of the mode transition results cannot be established without experimental data to validate the reaction mechanism.

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

    Science.gov (United States)

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

    2017-03-01

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

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

    Science.gov (United States)

    Brown, Andrew M.

    2014-01-01

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

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

    Science.gov (United States)

    Bennewitz, John William

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

  3. Numerical analyses of a rocket engine turbine and comparison with air test data

    Science.gov (United States)

    Tran, Ken; Chan, Daniel C.; Hudson, Susan T.; Gaddis, Stephen W.

    1992-01-01

    The study presents cold air test data on the Space Shuttle Main Engine High Pressure Fuel Turbopump turbine recently collected at the NASA Marshall Space Flight Center. Overall performance data, static pressures on the first- and second-stage nozzles, and static pressures along with the gas path at the hub and tip are gathered and compared with various (1D, quasi-3D, and 3D viscous) analysis procedures. The results of each level of analysis are compared to test data to demonstrate the range of applicability for each step in the design process of a turbine. One-dimensional performance prediction, quasi-3D loading prediction, 3D wall pressure distribution prediction, and 3D viscous wall pressure distribution prediction are illustrated.

  4. Some typical solid propellant rocket motors

    NARCIS (Netherlands)

    Zandbergen, B.T.C.

    2013-01-01

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

  5. Some typical solid propellant rocket motors

    NARCIS (Netherlands)

    Zandbergen, B.T.C.

    2013-01-01

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

  6. Experimental investigations on active cooling thermal protection structure of hydrocarbon-fueled scramjet combustor in arc heated facility

    Science.gov (United States)

    Jianqiang, Tu; Jinlong, Peng; Xianning, Yang; Lianzhong, Chen

    2016-10-01

    The active cooling thermal protection technology is the efficient method to resolve the long-duration work and reusable problems of hydrocarbon-fueled scramjet combustor, where worst thermo-mechanical loads occur. The fuel is passed through coolant channels adjacent to the heated surfaces to absorb heat from the heating exchanger panels, prior to injection into the combustor. The heating exchanger both cooled down the wall temperature of the combustor wall and heats and cracks the hydrocarbon fuel inside the panel to permit an easier combustion and satisfying combustion efficiency. The subscale active cooling metallic panels, with dimensions of 100×100 mm and different coolant channel sizes, have been tested under typical combustion thermal environment produced by arc heated Turbulent Flow Duct (TFD). The heat exchange ability of different coolant channel sizes has been obtained. The big-scale active cooling metallic panel, with dimensions of 100 × 750 mm and the coolant channel sizes of better heating exchange performance, has been made and tested in the big-scale arc heated TFD facility. The test results show that the local superheated ablation is easy to happen for the cooling fuel assigned asymmetrically in the bigscale active cooling metallic panel, and the cooling fuel rate can reduce 8%˜10% after spraying the Thermal Barrier Coating (TBC) in the heating surface.

  7. Biological Production of a Hydrocarbon Fuel Intermediate Polyhydroxybutyrate (PHB) from a Process Relevant Lignocellulosic Derived Sugar (Poster)

    Energy Technology Data Exchange (ETDEWEB)

    Wang, W.; Mittal, A.; Mohagheghi, A.; Johnson, D. K.

    2014-04-01

    PHAs are synthesized by many microorganisms to serve as intracellular carbon storage molecules. In some bacterial strains, PHB can account for up to 80% of cell mass. In addition to its application in the packaging sector, PHB also has great potential as an intermediate in the production of hydrocarbon fuels. PHB can be thermally depolymerized and decarboxylated to propene which can be upgraded to hydrocarbon fuels via commercial oligomerization technologies. Cupriavidus necator is the microorganism that has been most extensively studied and used for PHB production on an industrial scale; However the substrates used for producing PHB are mainly fructose, glucose, sucrose, fatty acids, glycerol, etc., which are expensive. In this study, we demonstrate production of PHB from a process relevant lignocellulosic derived sugar stream, i.e., saccharified slurry from pretreated corn stover. The strain was first investigated in shake flasks for its ability to utilize glucose, xylose and acetate. In addition, the strain was also grown on pretreated lignocellulose hydrolyzate slurry and evaluated in terms of cell growth, sugar utilization, PHB accumulation, etc. The mechanism of inhibition in the toxic hydrolysate generated by the pretreatment and saccharification process of biomass, was also studied.

  8. Analysis of reacting flowfields in low-thrust rocket engines and plumes

    Science.gov (United States)

    Weiss, Jonathan Mitchell

    The mixing and combustion processes in small gaseous hydrogen-oxygen thrusters and plumes are studied by means of a computational model developed as a general purpose analytic procedure for solving low speed, reacting, internal flowfields. The model includes the full Navier-Stokes equations coupled with species diffusion equations for a hydrogen-oxygen reaction kinetics system as well as the option to use either the k-Epsilon or q-Omega low Reynolds number, two-equation turbulence models. Solution of the governing equations is accomplished by a finite-volume formulation with central-difference spatial discretizations and an explicit, four-stage, Runge Kutta time-integration procedure. The Runge-Kutta scheme appears to provide efficient convergence when applied to the calculation of turbulent, reacting flowfields in these small thrusters. Appropriate boundary conditions are developed to properly model propellant mass flowrates and regenerative wall cooling. The computational method is validated against measured engine performance parameters on a global level, as well as experimentally obtained exit plane and plume flowfield properties on a local level. The model does an excellent job of predicting the measured performance trends of an auxiliary thruster as a function of O/F ratio, although the performance levels are consistently underpredicted by approximately 4 percent. These differences arise because the extent to which the wall coolant layer and combustion gases mix and react is underpredicted. Predictions of velocity components, temperature and species number densities in the near-field plume regions of several low-thrust engines show reasonable agreement with experimental data obtained by two separate laser diagnostic techniques. Discrepancies between the predictions and measurements are primarily due to three-dimensional mixing processes which are not accounted for in the analysis. Both comparisons with experiment and the evident reason for errors in absolute

  9. Prediction of engine and near-field plume reacting flows in low-thrust chemical rockets

    Science.gov (United States)

    Weiss, Jonathan M.; Merkle, Charles L.

    1993-01-01

    A computational model is employed to study the reacting flow within the engine and near-field plumes of several small gaseous hydrogen-oxygen thrusters. The model solves the full Navier-Stokes equations coupled with species diffusion equations for a hydrogen-oxygen reaction kinetics system and includes a two-equation q-omega model for turbulence. Predictions of global performance parameters and localized flowfield variables are compared with experimental data in order to assess the accuracy with which these flowfields are modeled and to identify aspects of the model which require improvement. Predicted axial and radial velocities 3 mm downstream of the exit plane show reasonable agreement with the measurements. The predicted peak in axial velocity in the hydrogen film coolant along the nozzle wall shows the best agreement; however, predictions within the core region are roughly 15 percent below measured values, indicating an underprediction of the extent to which the hydrogen diffuses and mixes with the core flow. There is evidence that this is due to three-dimensional mixing processes which are not included in the axisymmetric model.

  10. Application of powder metallurgy techniques to produce improved bearing elements for liquid rocket engines

    Science.gov (United States)

    Moracz, D. J.; Shipley, R. J.; Moxson, V. S.; Killman, R. J.; Munson, H. E.

    1992-01-01

    The objective was to apply powder metallurgy techniques for the production of improved bearing elements, specifically balls and races, for advanced cryogenic turbopump bearings. The materials and fabrication techniques evaluated were judged on the basis of their ability to improve fatigue life, wear resistance, and corrosion resistance of Space Shuttle Main Engine (SSME) propellant bearings over the currently used 440C. An extensive list of candidate bearing alloys in five different categories was considered: tool/die steels, through hardened stainless steels, cobalt-base alloys, and gear steels. Testing of alloys for final consideration included hardness, rolling contact fatigue, cross cylinder wear, elevated temperature wear, room and cryogenic fracture toughness, stress corrosion cracking, and five-ball (rolling-sliding element) testing. Results of the program indicated two alloys that showed promise for improved bearing elements. These alloys were MRC-2001 and X-405. 57mm bearings were fabricated from the MRC-2001 alloy for further actual hardware rig testing by NASA-MSFC.

  11. Labyrinth Seal Flutter Analysis and Test Validation in Support of Robust Rocket Engine Design

    Science.gov (United States)

    El-Aini, Yehia; Park, John; Frady, Greg; Nesman, Tom

    2010-01-01

    High energy-density turbomachines, like the SSME turbopumps, utilize labyrinth seals, also referred to as knife-edge seals, to control leakage flow. The pressure drop for such seals is order of magnitude higher than comparable jet engine seals. This is aggravated by the requirement of tight clearances resulting in possible unfavorable fluid-structure interaction of the seal system (seal flutter). To demonstrate these characteristics, a benchmark case of a High Pressure Oxygen Turbopump (HPOTP) outlet Labyrinth seal was studied in detail. First, an analytical assessment of the seal stability was conducted using a Pratt & Whitney legacy seal flutter code. Sensitivity parameters including pressure drop, rotor-to-stator running clearances and cavity volumes were examined and modeling strategies established. Second, a concurrent experimental investigation was undertaken to validate the stability of the seal at the equivalent operating conditions of the pump. Actual pump hardware was used to construct the test rig, also referred to as the (Flutter Rig). The flutter rig did not include rotational effects or temperature. However, the use of Hydrogen gas at high inlet pressure provided good representation of the critical parameters affecting flutter especially the speed of sound. The flutter code predictions showed consistent trends in good agreement with the experimental data. The rig test program produced a stability threshold empirical parameter that separated operation with and without flutter. This empirical parameter was used to establish the seal build clearances to avoid flutter while providing the required cooling flow metering. The calibrated flutter code along with the empirical flutter parameter was used to redesign the baseline seal resulting in a flutter-free robust configuration. Provisions for incorporation of mechanical damping devices were introduced in the redesigned seal to ensure added robustness

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

    Science.gov (United States)

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

    2017-01-01

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

  13. Analysis of Rocket, Ram-Jet, and Turbojet Engines for Supersonic Propulsion of Long-Range Missles. II - Rocket Missile Performance

    Science.gov (United States)

    Huff, Vearl N.; Kerrebrock, Jack

    1954-01-01

    The theoretical performance of a two-stage ballistic rocket mis having a centerbody and two parallel boosters was investigated for J oxygen and ammonia-fluorine propellants. Both power-plant and missi parameters were optimized to give minimum cost on-the basis of the analysis for a range of 5500 nautical miles. After optimum values were found, each parameter was varied independently to determine its effect on performance of the missile. The missile using the ammonia-fluorine propellant weighs about one half as much as a missile using JP4-oxygen. Based on an expected unit cost of fluorine in quantity production, the ammonia-fluorine missile has a substantially lower relative cost than a JP4-oxygen missile. Optimum chamber pressures for both propellant systems and for both the centerbody and boosters were between 450 and 600 pounds per square inch. High design altitudes for the exhaust nozzle are desirable for both the centerbody and boosters. For the centerbody, the design altitude should be between 45,000 and 60,000 feet, with the value for ammonia-fluorine lower than that for JP4-oxygen. For the boosters, the design altitude should be 20,000 to 30,000 feet, with the value for the ammonia-fluorine. missile higher.

  14. Understanding of Electrochemical Mechanisms for CO2 Capture and Conversion into Hydrocarbon Fuels in Transition-Metal Carbides (MXenes).

    Science.gov (United States)

    Li, Neng; Chen, Xingzhu; Ong, Wee-Jun; MacFarlane, Douglas R; Zhao, Xiujian; Cheetham, Anthony K; Sun, Chenghua

    2017-09-13

    Two-dimensional (2D) transition-metal (groups IV, V, VI) carbides (MXenes) with formulas M3C2 have been investigated as CO2 conversion catalysts with well-resolved density functional theory calculations. While MXenes from the group IV to VI series have demonstrated an active behavior for the capture of CO2, the Cr3C2 and Mo3C2 MXenes exhibit the most promising CO2 to CH4 selective conversion capabilities. Our results predicted the formation of OCHO(•) and HOCO(•) radical species in the early hydrogenation steps through spontaneous reactions. This provides atomic level insights into the computer-aided screening for high-performance catalysts and the understanding of electrochemical mechanisms for CO2 reduction to energy-rich hydrocarbon fuels, which is of fundamental significance to elucidate the elementary steps for CO2 fixation.

  15. Pressure Drop and Experiment of Liquid Rocket Engine Filter%液体火箭发动机用过滤器流阻特性及试验

    Institute of Scientific and Technical Information of China (English)

    窦唯

    2011-01-01

    Liquid rocket engine filter is an important element which ensures clean working liquid and reliable operation of the test equipment and engine. The reasonable filer design is a key factor to ensure liquid rocket engine success. Therefore, in order to prevent extra material appearance in liquid pipeline, gas pipeline etc. which would affect engine's normal work, a filter was designed according to the requirements of the engine system. The characteristics of the pressure drop are studied theoretically. And the liquid flow experiment is carried out. The analysis of theory and experimental results shows that the designed filter meets the requirements of the engine system.%液体火箭发动机用过滤器是保持工质清洁,保证试验设备和发动机可靠工作的重要设备,合理的过滤器设计是保证液体火箭发动机发射成败与否的关键因素.因此,为了防止液体火箭发动机液路、气路等管路出现多余物影响发动机正常工作,根据发动机系统的要求设计了某过滤器,从理论上研究了其流阻特性,并开展了液流试验研究.通过对比分析,理论和试验结果表明,所设计的过滤器满足发动机系统的要求.

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

    Directory of Open Access Journals (Sweden)

    V. S. Zarubin

    2016-01-01

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

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

    Science.gov (United States)

    Macgregor, C.; Csomor, A.

    1974-01-01

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

  18. Solid propellant rocket motor

    Science.gov (United States)

    Dowler, W. L.; Shafer, J. I.; Behm, J. W.; Strand, L. D. (Inventor)

    1973-01-01

    The characteristics of a solid propellant rocket engine with a controlled rate of thrust buildup to a desired thrust level are discussed. The engine uses a regressive burning controlled flow solid propellant igniter and a progressive burning main solid propellant charge. The igniter is capable of operating in a vacuum and sustains the burning of the propellant below its normal combustion limit until the burning propellant surface and combustion chamber pressure have increased sufficiently to provide a stable chamber pressure.

  19. 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

  20. Parametric Trends in the Combustion Stability Characteristics of a Single-Element Gas-Gas Rocket Engine

    Science.gov (United States)

    2013-12-01

    toroidal recirculation zone which promotes flame stabilization.16 For rocket applications swirl injectors are less sensitive to manufacturing defects and...were there are geometric changes. The mesh contains 5 of 22 American Institute of Aeronautics and Astronautics Figure 2: Geometric details of the

  1. 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...

  2. Numerical simulation analysis of water-hammer pressure of rocket engine%火箭发动机水击压力数值模拟分析

    Institute of Scientific and Technical Information of China (English)

    徐峰; 刘英元; 陈海峰

    2012-01-01

    依据水击理论,以某液体火箭发动机为例,采用特征线法建立推进剂供应系统一维流动数学模型,实现了关机水击数值模拟。实例计算分析了测压导管对水击过程的影响。%Taking one liquid rocket engine as example, the characteristic method is adopted to es- tablish a one-dimension mathematical model of the propellant feeding system on the basis of water-hammer pressure theory. The numerical simulation of water-hammer pressure as the engine is shut down. The effect of the manometry conduit on the water-hammer process is computed and analyzed with the example.

  3. Propellant Vaporization as a Criterion for Rocket-Engine Design; Experimental Performance, Vaporization and Heat-Transfer Rates with Various Propellant Combinations

    Science.gov (United States)

    Clark, Bruce J.; Hersch, Martin; Priem, Richard J.

    1959-01-01

    Experimental combustion efficiencies of eleven propellant combinations were determined as a function of chamber length. Efficiencies were measured in terms of characteristic exhaust velocities at three chamber lengths and in terms of gas velocities. The data were obtained in a nominal 200-pound-thrust rocket engine. Injector and engine configurations were kept essentially the same to allow comparison of the performance. The data, except for those on hydrazine and ammonia-fluorine, agreed with predicted results based on the assumption that vaporization of the propellants determines the rate of combustion. Decomposition in the liquid phase may be.responsible for the anomalous behavior of hydrazine. Over-all heat-transfer rates were also measured for each combination. These rates were close to the values predicted by standard heat-transfer calculations except for the combinations using ammonia.

  4. Technology Method Design of Assembly and Testing for Solid Propellant Rocket Engine of Aviation Seat%航空座椅固体火箭发动机装配及检测工艺技术设计

    Institute of Scientific and Technical Information of China (English)

    段祥军

    2014-01-01

    本文对航空座椅某型固体火箭发动机部装、总装及检测、试验、包装技术难点等进行了工艺分析;介绍了固体火箭发动机装配全过程工艺流程、检测、试验方法及注意事项等,对于同类及新型火箭发动机的装配制造过程具有良好的借鉴、推广应用意义。%Aiming at the difficulty of solid propellant rocket engine of aviation seat to assembly, testing and packaging technology, the assembly, testing process and method for solid propellant rocket engine were introduced. It can be regarded as reference with application for solid propellant rocket engine assembly process.

  5. Understanding of catalyst deactivation caused by sulfur poisoning and carbon deposition in steam reforming of liquid hydrocarbon fuels

    Science.gov (United States)

    Xie, Chao

    2011-12-01

    The present work was conducted to develop a better understanding on the catalyst deactivation in steam reforming of sulfur-containing liquid hydrocarbon fuels for hydrogen production. Steam reforming of Norpar13 (a liquid hydrocarbon fuel from Exxon Mobile) without and with sulfur was performed on various metal catalysts (Rh, Ru, Pt, Pd, and Ni) supported on different materials (Al2O3, CeO2, SiO2, MgO, and CeO2- Al2O3). A number of characterization techniques were applied to study the physicochemical properties of these catalysts before and after the reactions. Especially, X-ray absorption near edge structure (XANES) spectroscopy was intensively used to investigate the nature of sulfur and carbon species in the used catalysts to reveal the catalyst deactivation mechanism. Among the tested noble metal catalysts (Rh, Ru, Pt, and Pd), Rh catalyst is the most sulfur tolerant. Al2O3 and CeO2 are much better than SiO2 and MgO as the supports for the Rh catalyst to reform sulfur-containing hydrocarbons. The good sulfur tolerance of Rh/Al2O3 can be attributed to the acidic nature of the Al2O3 support and its small Rh crystallites (1-3 nm) as these characteristics facilitate the formation of electron-deficient Rh particles with high sulfur tolerance. The good catalytic performance of Rh/CeO2 in the presence of sulfur can be ascribed to the promotion effect of CeO2 on carbon gasification, which significantly reduced the carbon deposition on the Rh/CeO2catalyst. Steam reforming of Norpar13 in the absence and presence of sulfur was further carried out over CeO2-Al2O3 supported monometallic Ni and Rh and bimetallic Rh-Ni catalysts at 550 and 800 °C. Both monometallic catalysts rapidly deactivated at 550 °C, iv and showed poor sulfur tolerance. Although ineffective for the Ni catalyst, increasing the temperature to 800 °C dramatically improved the sulfur tolerance of the Rh catalyst. Sulfur K-edge XANES revealed that metal sulfide and organic sulfide are the dominant sulfur

  6. Landing screw-rockets array on asteroids, digging soil and fueling engines in phase, to overcome the spin and to fly in space

    CERN Document Server

    Fargion, D

    2007-01-01

    To deflect impact-trajectory of massive km^3 and spinning asteroid by a few terrestrial radiuses one need a large momentum exchange. The dragging of huge spinning bodies in space by external engine seems difficult or impossible. Our solution is based on the landing of multi screw-rockets, powered by mini-nuclear engines, on the body, that dig a small fraction of the soil surface, to use as an exhaust propeller, ejecting it vertically in phase among themselves. Such a mass ejection increases the momentum exchange, their number redundancy guarantes the stability of the system. The soft landing of engine-unity may be easely achieved at low asteroid gravity. The engine array tuned activity, overcomes the asteroid angular velocity. Coherent turning of the jet heads increases the deflection efficiency. A procession along its surface may compensate at best the asteroid spin. A small skin-mass (about 2 10^4 tons) may be ejected by mini nuclear engines. Such prototypes may build first save galleries for humans on the ...

  7. Rocket noise - A review

    Science.gov (United States)

    McInerny, S. A.

    1990-10-01

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

  8. Improvement of lean combustion characteristics of heavy-hydrocarbon fuels with hydrogen addition; Suiso tenka ni yoru kokyu tanka suisokei nenryo no kihaku nensho no kaizen

    Energy Technology Data Exchange (ETDEWEB)

    Sakai, Y. [Saitama Institute of Technology, Saitama (Japan); Ishizuka, S. [Hiroshima University, Hiroshima (Japan). Faculty of Engineering

    1999-09-25

    The Lewis numbers of lean heavy-hydrocarbon fuels are larger than unity, and hence, their flames are prone to extinction in a shear flow, which occurs in a turbulent combustion. Here, propane is used as a representative fuel of heavy-hydrocarbon fuels because the Lewis number of lean propane/air mixtures is larger than unity, and an attempt to improve its combustion characteristics by hydrogen addition has been made. A tubular flame burner is used to evaluate its improvement, since a rotating, stretched vortex flow is established in the burner. The results show that with' hydrogen addition, the fuel concentration, the flame diameter and the flame temperature at extinction are reduced and its combustion characteristics are improved. However, it is found that the effective equivalence ration at extinction cannot become so small as that of lean methane/air mixture, which has a Lewis number less than unity. (author)

  9. Self-doped Ti(3+)-TiO2 as a photocatalyst for the reduction of CO2 into a hydrocarbon fuel under visible light irradiation.

    Science.gov (United States)

    Sasan, Koroush; Zuo, Fan; Wang, Yuan; Feng, Pingyun

    2015-08-28

    Self-doped TiO2 shows visible light photocatalytic activity, while commercial TiO2 (P25) is only UV responsive. The incorporation of Ti(3+) into TiO2 structures narrows the band gap (2.90 eV), leading to significantly increased photocatalytic activity for the reduction of CO2 into a renewable hydrocarbon fuel (CH4) in the presence of water vapour under visible light irradiation.

  10. Ethanol dehydration via azeotropic distillation with gasoline fractions as entrainers: A pilot-scale study of the manufacture of an ethanol–hydrocarbon fuel blend

    OpenAIRE

    Gomis Yagües, Vicente; Pedraza Berenguer, Ricardo; Saquete Ferrándiz, María Dolores; Font, Alicia; García Cano, Jorge

    2015-01-01

    We establish experimentally and through simulations the economic and technical viability of dehydrating ethanol by means of azeotropic distillation, using a hydrocarbon as entrainer. The purpose of this is to manufacture a ready-to-use ethanol–hydrocarbon fuel blend. In order to demonstrate the feasibility of this proposition, we have tested an azeotropic water–ethanol feed mixture, using a hydrocarbon as entrainer, in a semi pilot-plant scale distillation column. Four different hydrocarbons ...

  11. A Dual-Line Detection Rayleigh Scattering Diagnostic Technique for the Combustion of Hydrocarbon Fuels and Filtered UV Rayleigh Scattering for Gas Velocity Measurements

    Science.gov (United States)

    Otugen, M. Volkan

    1997-01-01

    Non-intrusive techniques for the dynamic measurement of gas flow properties such as density, temperature and velocity, are needed in the research leading to the development of new generation high-speed aircraft. Accurate velocity, temperature and density data obtained in ground testing and in-flight measurements can help understand the flow physics leading to transition and turbulence in supersonic, high-altitude flight. Such non-intrusive measurement techniques can also be used to study combustion processes of hydrocarbon fuels in aircraft engines. Reliable, time and space resolved temperature measurements in various combustor configurations can lead to a better understanding of high temperature chemical reaction dynamics thus leading to improved modeling and better prediction of such flows. In view of this, a research program was initiated at Polytechnic University's Aerodynamics Laboratory with support from NASA Lewis Research Center through grants NAG3-1301 and NAG3-1690. The overall objective of this program has been to develop laser-based, non-contact, space- and time-resolved temperature and velocity measurement techniques. In the initial phase of the program a ND:YAG laser-based dual-line Rayleigh scattering technique was developed and tested for the accurate measurement of gas temperature in the presence of background laser glare. Effort was next directed towards the development of a filtered, spectrally-resolved Rayleigh/Mie scattering technique with the objective of developing an interferometric method for time-frozen velocity measurements in high-speed flows utilizing the uv line of an ND:YAG laser and an appropriate molecular absorption filter. This effort included both a search for an appropriate filter material for the 266 nm laser line and the development and testing of several image processing techniques for the fast processing of Fabry-Perot images for velocity and temperature information. Finally, work was also carried out for the development of

  12. Diagnostic development for determining the joint temperature/soot statistics in hydrocarbon-fueled pool fires : LDRD final report.

    Energy Technology Data Exchange (ETDEWEB)

    Casteneda, Jaime N.; Frederickson, Kraig; Grasser, Thomas W.; Hewson, John C.; Kearney, Sean Patrick; Luketa, Anay Josephine

    2009-09-01

    A joint temperature/soot laser-based optical diagnostic was developed for the determination of the joint temperature/soot probability density function (PDF) for hydrocarbon-fueled meter-scale turbulent pool fires. This Laboratory Directed Research and Development (LDRD) effort was in support of the Advanced Simulation and Computing (ASC) program which seeks to produce computational models for the simulation of fire environments for risk assessment and analysis. The development of this laser-based optical diagnostic is motivated by the need for highly-resolved spatio-temporal information for which traditional diagnostic probes, such as thermocouples, are ill-suited. The in-flame gas temperature is determined from the shape of the nitrogen Coherent Anti-Stokes Raman Scattering (CARS) signature and the soot volume fraction is extracted from the intensity of the Laser-Induced Incandescence (LII) image of the CARS probed region. The current state of the diagnostic will be discussed including the uncertainty and physical limits of the measurements as well as the future applications of this probe.

  13. Leaf-architectured 3D Hierarchical Artificial Photosynthetic System of Perovskite Titanates Towards CO2 Photoreduction Into Hydrocarbon Fuels

    Science.gov (United States)

    Zhou, Han; Guo, Jianjun; Li, Peng; Fan, Tongxiang; Zhang, Di; Ye, Jinhua

    2013-04-01

    The development of an ``artificial photosynthetic system'' (APS) having both the analogous important structural elements and reaction features of photosynthesis to achieve solar-driven water splitting and CO2 reduction is highly challenging. Here, we demonstrate a design strategy for a promising 3D APS architecture as an efficient mass flow/light harvesting network relying on the morphological replacement of a concept prototype-leaf's 3D architecture into perovskite titanates for CO2 photoreduction into hydrocarbon fuels (CO and CH4). The process uses artificial sunlight as the energy source, water as an electron donor and CO2 as the carbon source, mimicking what real leaves do. To our knowledge this is the first example utilizing biological systems as ``architecture-directing agents'' for APS towards CO2 photoreduction, which hints at a more general principle for APS architectures with a great variety of optimized biological geometries. This research would have great significance for the potential realization of global carbon neutral cycle.

  14. Catalytic conversion of palm oil over mesoporous aluminosilicate MCM-41 for the production of liquid hydrocarbon fuels

    Energy Technology Data Exchange (ETDEWEB)

    Twaiq, Farouq A.; Mohamed, Abdul Rahman; Bhatia, Subhash [School of Chemical Engineering, Engineering Campus, Universiti Sains Malaysia, 14300 Nibong Tebal, SPS, Pinang (Malaysia); Zabidi, Noor Asmawati M. [Universiti Teknologi Petronas, Sri Iskandar, 31750 Tronoh, Perak (Malaysia)

    2003-11-15

    The catalytic cracking of palm oil to liquid hydrocarbon fuels was studied in a fixed bed micro-reactor operated at atmospheric pressure, reaction temperature of 723 K and weight hourly space velocity (WHSV) of 2.5 h{sup -1} over the synthesized mesoporous molecular sieve MCM-41 materials. Mesoporous aluminosilicate with Si/Al ratio of 50 was synthesized using the hydrothermal method. Different pore sizes were obtained by changing the type of template and organic directing agent (ODA) used. The synthesized materials were characterized using various analytical methods such as X-ray powder diffraction (XRD), BET surface area, inductive coupled plasma (ICP), MAS NMR, FTIR and temperature-programmed desorption (TPD). The materials exhibit a crystalline structure of MCM-41 mesoporous molecular sieves with surface area varying from 550 to 1200 m{sup 2}/g and an average pore size (APS) ranging from 1.8 to 2.8 nm. The synthesized MCM-41 catalysts show high activity for palm oil cracking. The conversion of palm kernel oil, lower-molecular-weight oil, was higher as compared to higher-molecular-weight, palm olein oil. MCM-41 materials were selective for the formation of linear hydrocarbons, particularly, C{sub 13} when palm kernel oil was used and C{sub 17} when palm olein oil was fed. The yield of liquid product decreased with the increase of surface area of the catalyst. The gasoline selectivity increased whereas diesel selectivity decreased with the conversion of palm oil.

  15. Valorization of Waste Lipids through Hydrothermal Catalytic Conversion to Liquid Hydrocarbon Fuels with in Situ Hydrogen Production

    Energy Technology Data Exchange (ETDEWEB)

    Kim, Dongwook; Vardon, Derek R.; Murali, Dheeptha; Sharma, Brajendra K.; Strathmann, Timothy J.

    2016-03-07

    We demonstrate hydrothermal (300 degrees C, 10 MPa) catalytic conversion of real waste lipids (e.g., waste vegetable oil, sewer trap grease) to liquid hydrocarbon fuels without net need for external chemical inputs (e.g., H2 gas, methanol). A supported bimetallic catalyst (Pt-Re/C; 5 wt % of each metal) previously shown to catalyze both aqueous phase reforming of glycerol (a triacylglyceride lipid hydrolysis coproduct) to H2 gas and conversion of oleic and stearic acid, model unsaturated and saturated fatty acids, to linear alkanes was applied to process real waste lipid feedstocks in water. For reactions conducted with an initially inert headspace gas (N2), waste vegetable oil (WVO) was fully converted into linear hydrocarbons (C15-C17) and other hydrolyzed byproducts within 4.5 h, and H2 gas production was observed. Addition of H2 to the initial reactor headspace accelerated conversion, but net H2 production was still observed, in agreement with results obtained for aqueous mixtures containing model fatty acids and glycerol. Conversion to liquid hydrocarbons with net H2 production was also observed for a range of other waste lipid feedstocks (animal fat residuals, sewer trap grease, dry distiller's grain oil, coffee oil residual). These findings demonstrate potential for valorization of waste lipids through conversion to hydrocarbons that are more compatible with current petroleum-based liquid fuels than the biodiesel and biogas products of conventional waste lipid processing technologies.

  16. A quantum chemistry study on thermochemical properties of high energy-density endothermic hydrocarbon fuel JP-10.

    Science.gov (United States)

    Qin, Xiao-Mei; Xie, Hu-Jun; Yue, Lei; Lu, Xiao-Xing; Fang, Wen-Jun

    2014-04-01

    The density functional theory (DFT) calculations at the M06-2X/6-31++G(d,p) level have been performed to explore the molecular structure, electronic structure, C-H bond dissociation enthalpy, and reaction enthalpies for five isodesmic reactions of a high energy-density endothermic hydrocarbon fuel JP-10. On the basis of the calculations, it is found that the carbonium ion C-6 isomer formed from the catalytic cracking at the C₆ site of JP-10 has the lowest energy, and the R-5 radical generated from the thermal cracking at the C₅ site of JP-10 is the most stable isomer. Furthermore, a series of hypothetical and isodesmic work reactions containing similar bond environments are used to calculate the reaction enthalpies for target compounds. For the same isodesmic reaction, the reaction enthalpy of each carbon site radical has also been calculated. The present work is of fundamental significance and strategic importance to provide some valuable insights into the component design and energy utilization of advanced endothermic fuels.

  17. General simulation on modularization of liquid rocket engine%液体火箭发动机模块化通用仿真

    Institute of Scientific and Technical Information of China (English)

    张黎辉; 李伟; 段娜

    2011-01-01

    以大型液体火箭发动机研制为背景,根据模块化建模思想和通用仿真要求,提出一种具有良好通用性和系统组织能力的仿真方法,数学模型在模块内的存储形式是代码文本,操作者只需从界面即可添加模型,扩展模块库.描述了实现这种方法的软件架构,在Microsoft Visual Studio 6.0平台上,采用C++语言开发出该仿真软件,并在MFC(Microsoft Foundation Classes)文档视图结构的基础上实现了可视化建模,完成了对某液体火箭发动机瞬变过程仿真.通过计算结果与试车曲线的对比,初步验证了所采用仿真方法的可行性和正确性.研究工作可以很方便地实现对其他液体火箭发动机系统动态过程的仿真.%Based on the development of large liquid rocket engines, a universal system organization capability approach was presented according to the thought of modularization modeling and the requirements of general simulation. Mathematical model in the module was stored in the form of code text; and the operator could add models or expand block library from the operator interface. This paper described the approach of realizing such software architecture. In Microsoft Visual Studio 6.0 platform, C+ + language was utilized to develop the simulation software and implement visual modeling based on MFC (Microsoft Foundation Classes) document view architecture, and finally a particular transient simulation of liquid propellant rocket engine was fulfilled. By comparing the results with the test run curve,the feasibility and correctness by adopting the simulation method was verified primarily.The researchers can carry out expediently dynamic process simulation in other liquid rocket engines.

  18. Computational modeling of nuclear thermal rockets

    Science.gov (United States)

    Peery, Steven D.

    1993-01-01

    The topics are presented in viewgraph form and include the following: rocket engine transient simulation (ROCETS) system; ROCETS performance simulations composed of integrated component models; ROCETS system architecture significant features; ROCETS engineering nuclear thermal rocket (NTR) modules; ROCETS system easily adapts Fortran engineering modules; ROCETS NTR reactor module; ROCETS NTR turbomachinery module; detailed reactor analysis; predicted reactor power profiles; turbine bypass impact on system; and ROCETS NTR engine simulation summary.

  19. Laser Transmission Measurements of Soot Extinction Coefficients in the Exhaust Plume of the X-34 60k-lb Thrust Fastrac Rocket Engine

    Science.gov (United States)

    Dobson, C. C.; Eskridge, R. H.; Lee, M. H.

    2000-01-01

    A four-channel laser transmissometer has been used to probe the soot content of the exhaust plume of the X-34 60k-lb thrust Fastrac rocket engine at NASA's Marshall Space Flight Center. The transmission measurements were made at an axial location about equal 1.65 nozzle diameters from the exit plane and are interpreted in terms of homogeneous radial zones to yield extinction coefficients from 0.5-8.4 per meter. The corresponding soot mass density, spatially averaged over the plume cross section, is, for Rayleigh particles, approximately equal to 0.7 micrograms/cubic cm and alternative particle distributions are briefly considered. Absolute plume radiance at the laser wavelength (515 nm) is estimated from the data at approximately equal to 2.200 K equivalent blackbody temperature, and temporal correlations in emission from several spatial locations are noted.

  20. Design study of RL10 derivatives. Volume 2: Engine design characteristics, appendices. [development of rocket engine for application to space tug propulsion system

    Science.gov (United States)

    1973-01-01

    Calculations, curves, and substantiating data which support the engine design characteristics of the RL-10 engines are presented. A description of the RL-10 ignition system is provided. The performance calculations of the RL-10 derivative engines and the performance results obtained are reported. The computer simulations used to establish the control system requirements and to define the engine transient characteristics are included.

  1. Photographic Study of Combustion in a Rocket Engine I : Variation in Combustion of Liquid Oxygen and Gasoline with Seven Methods of Propellant Injection

    Science.gov (United States)

    Bellman, Donald R; Humphrey, Jack C

    1948-01-01

    Motion pictures at camera speeds up to 3000 frames per second were taken of the combustion of liquid oxygen and gasoline in a 100-pound-thrust rocket engine. The engine consisted of thin contour and injection plates clamped between two clear plastic sheets forming a two-dimensional engine with a view of the entire combustion chamber and nozzle. A photographic investigation was made of the effect of seven methods of propellant injection on the uniformity of combustion. From the photographs, it was found that the flame front extended almost to the faces of the injectors with most of the injection methods, all the injection systems resulted in a considerable nonuniformity of combustion, and luminosity rapidly decreased in the divergent part of the nozzle. Pressure vibration records indicated combustion vibrations that approximately corresponded to the resonant frequencies of the length and the thickness of the chamber. The combustion temperature divided by the molecular weight of the combustion gases as determined from the combustion photographs was about 50 to 70 percent of the theoretical value.

  2. Next Generation Heavy-Lift Launch Vehicle: Large Diameter, Hydrocarbon-Fueled Concepts

    Science.gov (United States)

    Holliday, Jon; Monk, Timothy; Adams, Charles; Campbell, Ricky

    2012-01-01

    With the passage of the 2010 NASA Authorization Act, NASA was directed to begin the development of the Space Launch System (SLS) as a follow-on to the Space Shuttle Program. The SLS is envisioned as a heavy lift launch vehicle that will provide the foundation for future large-scale, beyond low Earth orbit (LEO) missions. Supporting the Mission Concept Review (MCR) milestone, several teams were formed to conduct an initial Requirements Analysis Cycle (RAC). These teams identified several vehicle concept candidates capable of meeting the preliminary system requirements. One such team, dubbed RAC Team 2, was tasked with identifying launch vehicles that are based on large stage diameters (up to the Saturn V S-IC and S-II stage diameters of 33 ft) and utilize high-thrust liquid oxygen (LOX)/RP engines as a First Stage propulsion system. While the trade space for this class of LOX/RP vehicles is relatively large, recent NASA activities (namely the Heavy Lift Launch Vehicle Study in late 2009 and the Heavy Lift Propulsion Technology Study of 2010) examined specific families within this trade space. Although the findings from these studies were incorporated in the Team 2 activity, additional branches of the trade space were examined and alternative approaches to vehicle development were considered. Furthermore, Team 2 set out to define a highly functional, flexible, and cost-effective launch vehicle concept. Utilizing this approach, a versatile two-stage launch vehicle concept was chosen as a preferred option. The preferred vehicle option has the capability to fly in several different configurations (e.g. engine arrangements) that gives this concept an inherent operational flexibility which allows the vehicle to meet a wide range of performance requirements without the need for costly block upgrades. Even still, this concept preserves the option for evolvability should the need arise in future mission scenarios. The foundation of this conceptual design is a focus on low

  3. n-Hexadecane Fuel for a Phosphoric Acid Direct Hydrocarbon Fuel Cell

    Directory of Open Access Journals (Sweden)

    Yuanchen Zhu

    2015-01-01

    Full Text Available The objective of this work was to examine fuel cells as a possible alternative to the diesel fuel engines currently used in railway locomotives, thereby decreasing air emissions from the railway transportation sector. We have investigated the performance of a phosphoric acid fuel cell (PAFC reactor, with n-hexadecane, C16H34 (a model compound for diesel fuel, cetane number = 100. This is the first extensive study reported in the literature in which n-hexadecane is used directly as the fuel. Measurements were made to obtain both polarization curves and time-on-stream results. Because deactivation was observed hydrogen polarization curves were measured before and after n-hexadecane experiments, to determine the extent of deactivation of the membrane electrode assembly (MEA. By feeding water-only (no fuel to the fuel cell anode the deactivated MEAs could be regenerated. One set of fuel cell operating conditions that produced a steady-state was identified. Identification of steady-state conditions is significant because it demonstrates that stable fuel cell operation is technically feasible when operating a PAFC with n-hexadecane fuel.

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

    Science.gov (United States)

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

    2010-01-01

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

  5. 碳氢燃料低压裂解特性%Pyrolysis Characteristics of Hydrocarbon Fuel Under Low Pressure

    Institute of Scientific and Technical Information of China (English)

    张其翼; 魏微; 周灏; 文旭; 李象远

    2015-01-01

    针对主动冷却中利用碳氢燃料裂解吸热释放热沉的特性,研究分析了碳氢燃料在低压条件下的裂解反应特征,从产气率、气相组分与结焦等方面进行了对比研究.结果表明,在相同停留时间,压力降低,裂解的产气率更高,裂解气组分中烯烃含量大幅提升,表明低压裂解具有一定的反应定向性.通过吸光光度法半定量分析高温裂解残液,发现裂解压力降低,吸光度值降低,表明低压裂解可以明显抑制结焦反应发生.%This study focuses on the cracking reaction of hydrocarbon fuel under low pressure.The rate of pyrolysis gas yielding,pyrolysis gas composition and coke formation are studied and compared.The results demonstrate that during the same residence time,the rate of gas yielding increases while the pressure is reduced.A substantial rise of the olefin content in the pyrolysis gas indicates that the cracking reaction is directed.The semi-quantitative analysis of pyrolysis residue absorbance shows that absorbance values are reduced with the decrease of cracking pressure,and low pressure pyrolysis can significantly inhibit coking reaction.

  6. Hydrocarbon fuels from gas phase decarboxylation of hydrolyzed free fatty acid

    KAUST Repository

    Wang, Weicheng

    2012-01-01

    Gas phase decarboxylation of hydrolyzed free fatty acid (FFA) from canola oil has beeninvestigated in two fix-bed reactors by changing reaction parameters such as temperatures,FFA feed rates, and H 2-to-FFA molar ratios. FFA, which contains mostly C 18 aswell as a few C 16, C 20, C 22, and C 24 FFA, was fed into the boiling zone, evaporated, carriedby hydrogen flow at the rate of 0.5-20 ml/min, and reacted with the 5% Pd/C catalystin the reactor. Reactions were conducted atmospherically at 380-450 °C and the products,qualified and quantified through gas chromatography-flame ionization detector(GC-FID), showed mostly n-heptadecane and a few portion of n-C 15, n-C 19, n-C 21, n-C 23 as well as some cracking species. Results showed that FFA conversion increased withincreasing reaction temperatures but decreased with increasing FFA feed rates and H 2-to-FFA molar ratios. The reaction rates were found to decrease with higher temperatureand increase with higher H 2 flow rates. Highly selective heptadecane was achieved byapplying higher temperatures and higher H 2-to-FFA molar ratios. From the results, ascatalyst loading and FFA feed rate were fixed, an optimal reaction temperature of 415 °C as well as H 2-to-FFA molar ratio of 4.16 were presented. These results provided goodbasis for studying the kinetics of decarboxylation process. © 2012 American Society of Mechanical Engineers.

  7. Understanding the role of multifunctional nanoengineered particulate additives on supercritical pyrolysis and combustion of hydrocarbon fuels/propellants

    Science.gov (United States)

    Sim, Hyung Sub

    This dissertation aims to understand the fundamental effects of colloidal nanostructured materials on the supercritical pyrolysis, injection, ignition, and combustion of hydrocarbon fuels/propellants. As a fuel additive, functionalized graphene sheets (FGS) without or with the decoration of metal catalysts, such as platinum (Pt) or polyoxometalates (POM) nanoparticles, were examined against conventional materials including nanometer sized fumed silica and aluminum particles. Supercritical pyrolysis experiments were performed as a function of temperature, residence time, and particle type, using a high pressure and temperature flow reactor designed to provide isothermal and isobaric flow conditions. Supercritical pyrolysis results showed that the addition of FGS-based particles at a loading concentration of 50 ppmw increased the conversion rates and reduced apparent activation energies for methylcyclohexane (MCH) and n-dodecane (n-C12H26) fuels. For example, conversion rates, and formations of C1-C5 n-alkanes and C2-C6 1-alkenes were significantly increased by 43.5 %, 59.1 %, and 50.0 % for MCH decomposition using FGS 19 (50 ppmw) at a temperature of 820 K and reduced pressure of 1.36. In addition, FGS decorated with 20 wt % Pt (20wt%Pt FGS) at a loading concentration of 50 ppmw exhibited additional enhancement in the conversion rate of n-C12H26 by up to 24.0 % compared to FGS. Especially, FGS-based particles seem to alter initiation mechanisms, which could result in higher hydrogen formation. Hydrogen selectivities for both MCH and n-C12H26 decompositions were observed to increase by nearly a factor of 2 and 10, respectively. Supercritical injection and combustion experiments were conducted using a high pressure and temperature windowed combustion chamber coupled to the flow reactor through a feed system. Supercritical injection/combustion experiments indicated that the presence of a small amount of particles (100 ppmw) in the fuel affected the injection, ignition

  8. Rocket Tablet,

    Science.gov (United States)

    1984-09-12

    is a vast and desolate world, this is a strip of mir- aculous land! How many struggling dramas full of power and * grandeur were cheered, resisted and...rocket officers and men, a group enormous and powerful , marched into this land soaked with the fresh blood of our ancestors. This place is about to...and tough pestering said he wanted an American aircraft ob- tained on the battlefield to transport goods from Lanzhou, Xian, Beijing, Guangzhou and

  9. 基于Simulink的火箭弹发动机内弹道性能%Simulation of Trajectory Mode of Rocket Engine on Basis of Simulink

    Institute of Scientific and Technical Information of China (English)

    黄延平; 高俊国

    2012-01-01

    研究火箭弹发动机内弹道性能,主要是对燃烧室内压强随时间的变化规律进行研究.一般采用实验法可以获得直观可靠的数据,由于在燃烧室中的压强很大,对实验的设备要求较苛刻,需要具备一定的条件才能进行实验.根据零维内弹道数学模型,应用Simulink仿真模型对某型固体火箭弹发动机内弹道工作过程进行数值仿真,画出燃烧室内压强曲线,进一步分析并得出影响发动机内弹道性能的因素.%The interior ballistics performance of rocket engine is studied,and mainly carry on research with the changing law of pressure in the blast chamber while time changes, at the same time experimentation can achieve the intuitionistic and credible data, as the pressure in the blast chamber is extremely high, which requires the apparatus of equipment very harsh, also need to possess certain condition to carrying on the experiment. According to mathematic model of zero-dimensional interior ballistics, the mode of Simulink is applied to take numerical simulating of the interior ballistics working process of a certain rocket projectile. The pressure-time curve of the blast chamber is drawn, further the factors influencing the interior ballistics performance is analyzed and obtained.

  10. Hybrid Rocket Technology

    Directory of Open Access Journals (Sweden)

    Sankaran Venugopal

    2011-04-01

    Full Text Available With their unique operational characteristics, hybrid rockets can potentially provide safer, lower-cost avenues for spacecraft and missiles than the current solid propellant and liquid propellant systems. Classical hybrids can be throttled for thrust tailoring, perform in-flight motor shutdown and restart. In classical hybrids, the fuel is stored in the form of a solid grain, requiring only half the feed system hardware of liquid bipropellant engines. The commonly used fuels are benign, nontoxic, and not hazardous to store and transport. Solid fuel grains are not highly susceptible to cracks, imperfections, and environmental temperature and are therefore safer to manufacture, store, transport, and use for launch. The status of development based on the experience of the last few decades indicating the maturity of the hybrid rocket technology is given in brief.Defence Science Journal, 2011, 61(3, pp.193-200, DOI:http://dx.doi.org/10.14429/dsj.61.518

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

    Science.gov (United States)

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

    2014-01-01

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

  12. Diesel Engine Convert to Port Injection CNG Engine Using Gaseous Injector Nozzle Multi Holes Geometries Improvement: A Review

    OpenAIRE

    Semin; Abdul R. Ismail; Rosli A. Bakar

    2009-01-01

    The objective of this study was to review the previous research in the development of gaseous fuel injector for port injection CNG engine converted from diesel engine. Problem statement: The regular development of internal combustion engines change direction to answer the two most important problems determining the development trends of engines technology and in particular, their combustion systems. They were environmental protection against emission and noise, shortage of hydrocarbon fuels, ...

  13. 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

  14. Combustion efficiency and altitude operational limits of three liquid hydrocarbon fuels having high volumetric energy content in a J33 single combustor

    Science.gov (United States)

    Stricker, Edward G

    1950-01-01

    Combustion efficiency and altitude operational limits were determined in a J33 single combustor for AN-F-58 fuel and three liquid hydrocarbon fuels having high volumetric energy content (decalin, tetralin, and monomethylnaphthalene) at simulated altitude and combustor inlet-air conditions. At the conditions investigated, the combustion efficiency for the four fuels generally decreased with an increase in volumetric energy content. The altitude operational limits for decalin and tetralin fuels were higher than for AN-F-58 fuel; monomethylnaphthalene fuel gave the lowest altitude operational limit.

  15. Direct hydrocarbon fuel cells

    Science.gov (United States)

    Barnett, Scott A.; Lai, Tammy; Liu, Jiang

    2010-05-04

    The direct electrochemical oxidation of hydrocarbons in solid oxide fuel cells, to generate greater power densities at lower temperatures without carbon deposition. The performance obtained is comparable to that of fuel cells used for hydrogen, and is achieved by using novel anode composites at low operating temperatures. Such solid oxide fuel cells, regardless of fuel source or operation, can be configured advantageously using the structural geometries of this invention.

  16. Low-thrust rocket trajectories

    Energy Technology Data Exchange (ETDEWEB)

    Keaton, P.W.

    1986-01-01

    The development of low-thrust propulsion systems to complement chemical propulsion systems will greatly enhance the evolution of future space programs. Two advantages of low-thrust rockets are stressed: first, in a strong gravitational field, such as occurs near the Earth, freighter missions with low-thrust engines require one-tenth as much propellant as do chemical engines. Second, in a weak gravitational field, such as occurs in the region between Venus and Mars, low-thrust rockets are faster than chemical rockets with comparable propellant mass. The purpose here is to address the physics of low-thrust trajectories and to interpret the results with two simple models. Analytic analyses are used where possible - otherwise, the results of numerical calculations are presented in graphs. The author has attempted to make this a self-contained report. 57 refs., 10 figs.

  17. Low-thrust rocket trajectories

    Energy Technology Data Exchange (ETDEWEB)

    Keaton, P.W.

    1987-03-01

    The development of low-thrust propulsion systems to complement chemical propulsion systems will greatly enhance the evolution of future space programs. Two advantages of low-thrust rockets are stressed: first, in a strong gravitational field, such as occurs near the Earth, freighter missions with low-thrust engines require one-tenth as much propellant as do chemical engines. Second, in a weak gravitational field, such as occurs in the region between Venus and Mars, low-thrust rockets are faster than chemical rockets with comparable propellant mass. The purpose here is to address the physics of low-thrust trajectories and to interpret the results with two simple models. Analytic analyses are used where possible - otherwise, the results of numerical calculations are presented in graphs. The author has attempted to make this a self-contained report.

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

    Science.gov (United States)

    Blotzer, Michael J.; Woods, Jody L.

    2009-01-01

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

  19. 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.

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

    Science.gov (United States)

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

    2013-01-01

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

  1. The Role of the Strutjet Engine in New Global and Space Markets

    Science.gov (United States)

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

    1998-01-01

    The Strutjet, discussed in previous IAF papers, was originally introduced as an enabling propulsion concept for single stage to orbit applications. Recent design considerations indicate that this systems also provides benefits supportive of other commercial non-space applications. This paper describes the technical progress of the Strutjet since 1997 together with a rationale why Rocket Based Combined Cycle Engines in general, and the Strutjet in particular, lend themselves uniquely to systems having the ability to expand current space and open new global 'rapid delivery' markets. During this decade, Strutjet technology has been evaluated in over 1000 tests. Its design maturity has been continuously improved and desired features, like simple variable geometry and low drag flowpath resulting in high performance, have been verified. In addition, data is now available which allows the designer, who is challenged to maximize system operability and economic feasibility, to choose between hydrogen or hydrocarbon fuels for a variety of application. The ability exists now to apply this propulsion system to various vehicles with a multitude of missions. In this paper, storable hydrocarbon and gaseous hydrogen Strutjet RBCC test data as accomplished to date and as planned for the future is presented, and the degree of required technology maturity achieved so far is assessed. Two vehicles, using cryogenic propane fuel Strutjet engines, and specifically designed for rapid point-to-point cargo delivery between Pacific rim locations are introduced, discussed, and compared.

  2. Self-doped Ti3+-TiO2 as a photocatalyst for the reduction of CO2 into a hydrocarbon fuel under visible light irradiation

    Science.gov (United States)

    Sasan, Koroush; Zuo, Fan; Wang, Yuan; Feng, Pingyun

    2015-08-01

    Self-doped TiO2 shows visible light photocatalytic activity, while commercial TiO2 (P25) is only UV responsive. The incorporation of Ti3+ into TiO2 structures narrows the band gap (2.90 eV), leading to significantly increased photocatalytic activity for the reduction of CO2 into a renewable hydrocarbon fuel (CH4) in the presence of water vapour under visible light irradiation.Self-doped TiO2 shows visible light photocatalytic activity, while commercial TiO2 (P25) is only UV responsive. The incorporation of Ti3+ into TiO2 structures narrows the band gap (2.90 eV), leading to significantly increased photocatalytic activity for the reduction of CO2 into a renewable hydrocarbon fuel (CH4) in the presence of water vapour under visible light irradiation. Electronic supplementary information (ESI) available: Experimental details, XPS, XRD and SEM images. See DOI: 10.1039/c5nr02974k

  3. 火箭发动机喷管真空加压钎焊技术与设备%THE VACUUM PRESUURE BRAZING TECHNOLOGY AND EQUIPMENT FOR ROCKET ENGINE SPOUT

    Institute of Scientific and Technical Information of China (English)

    牛小莉

    2012-01-01

    介绍了火箭发动机喷管真空加压钎焊技术工艺原理及特点,真空加压钎焊设备的结构和工艺过程.该技术工艺是对火箭发动机喷管夹层抽空,在达到钎焊温度时.炉膛内充人保护性气体.满足工艺所需0.8 MPa的外压力条件,对火箭发动机喷管形成真空钎焊与真空扩散焊两种焊接方式相结合的综合性工艺方法.%This paper introduces a rocket engine spout vacuum and pressure brazing technology principle and characteristic of vacuum braze welding, introduces the structure and process equipment. The technology is on the rocket engine spout dissection in time, to the temperature within the furnace brazing, filled with protective gas, process to meet the required 0.8Mpa outer pressure condition, the rocket engine spout to form a vacuum brazing and vacuum diffusion welding two welding methods of combining a comprehensive process.

  4. Collaborative Sounding Rocket launch in Alaska and Development of Hybrid Rockets

    Science.gov (United States)

    Ono, Tomohisa; Tsutsumi, Akimasa; Ito, Toshiyuki; Kan, Yuji; Tohyama, Fumio; Nakashino, Kyouichi; Hawkins, Joseph

    Tokai University student rocket project (TSRP) was established in 1995 for a purpose of the space science and engineering hands-on education, consisting of two space programs; the one is sounding rocket experiment collaboration with University of Alaska Fairbanks and the other is development and launch of small hybrid rockets. In January of 2000 and March 2002, two collaborative sounding rockets were successfully launched at Poker Flat Research Range in Alaska. In 2001, the first Tokai hybrid rocket was successfully launched at Alaska. After that, 11 hybrid rockets were launched to the level of 180-1,000 m high at Hokkaido and Akita in Japan. Currently, Tokai students design and build all parts of the rockets. In addition, they are running the organization and development of the project under the tight budget control. This program has proven to be very effective in providing students with practical, real-engineering design experience and this program also allows students to participate in all phases of a sounding rocket mission. Also students learn scientific, engineering subjects, public affairs and system management through experiences of cooperative teamwork. In this report, we summarize the TSRP's hybrid rocket program and discuss the effectiveness of the program in terms of educational aspects.

  5. Development of a numerical tool to study the mixing phenomenon occurring during mode one operation of a multi-mode ejector-augmented pulsed detonation rocket engine

    Science.gov (United States)

    Dawson, Joshua

    A novel multi-mode implementation of a pulsed detonation engine, put forth by Wilson et al., consists of four modes; each specifically designed to capitalize on flow features unique to the various flow regimes. This design enables the propulsion system to generate thrust through the entire flow regime. The Multi-Mode Ejector-Augmented Pulsed Detonation Rocket Engine operates in mode one during take-off conditions through the acceleration to supersonic speeds. Once the mixing chamber internal flow exceeds supersonic speed, the propulsion system transitions to mode two. While operating in mode two, supersonic air is compressed in the mixing chamber by an upstream propagating detonation wave and then exhausted through the convergent-divergent nozzle. Once the velocity of the air flow within the mixing chamber exceeds the Chapman-Jouguet Mach number, the upstream propagating detonation wave no longer has sufficient energy to propagate upstream and consequently the propulsive system shifts to mode three. As a result of the inability of the detonation wave to propagate upstream, a steady oblique shock system is established just upstream of the convergent-divergent nozzle to initiate combustion. And finally, the propulsion system progresses on to mode four operation, consisting purely of a pulsed detonation rocket for high Mach number flight and use in the upper atmosphere as is needed for orbital insertion. Modes three and four appear to be a fairly significant challenge to implement, while the challenge of implementing modes one and two may prove to be a more practical goal in the near future. A vast number of potential applications exist for a propulsion system that would utilize modes one and two, namely a high Mach number hypersonic cruise vehicle. There is particular interest in the dynamics of mode one operation, which is the subject of this research paper. Several advantages can be obtained by use of this technology. Geometrically the propulsion system is fairly

  6. Process Design and Economics for the Conversion of Lignocellulosic Biomass to Hydrocarbon Fuels. Thermochemical Research Pathways with In Situ and Ex Situ Upgrading of Fast Pyrolysis Vapors

    Energy Technology Data Exchange (ETDEWEB)

    Dutta, A.; Sahir, A.; Tan, E.; Humbird, D.; Snowden-Swan, L. J.; Meyer, P.; Ross, J.; Sexton, D.; Yap, R.; Lukas, J.

    2015-03-01

    This report was developed as part of the U.S. Department of Energy’s Bioenergy Technologies Office’s efforts to enable the development of technologies for the production of infrastructurecompatible, cost-competitive liquid hydrocarbon fuels from biomass. Specifically, this report details two conceptual designs based on projected product yields and quality improvements via catalyst development and process integration. It is expected that these research improvements will be made within the 2022 timeframe. The two conversion pathways detailed are (1) in situ and (2) ex situ upgrading of vapors produced from the fast pyrolysis of biomass. While the base case conceptual designs and underlying assumptions outline performance metrics for feasibility, it should be noted that these are only two of many other possibilities in this area of research. Other promising process design options emerging from the research will be considered for future techno-economic analysis.

  7. Dynamic Combustion Stability Rating of LOX/LH2 Rocket Engine%氢氧火箭发动机动态燃烧稳定性评定技术研究

    Institute of Scientific and Technical Information of China (English)

    丁兆波; 许晓勇; 乔桂玉; 陶瑞峰

    2013-01-01

      为了实现氢氧发动机的动态燃烧稳定性试验评定,基于国内外液体火箭发动机动态稳定性评定的相关经验,并结合 CPIA655关于稳定性评定的准则,进行了氢氧发动机动态稳定性评定的方案探讨。分析表明,氢氧发动机有必要在全系统热试车状态下进行动态稳定性评定试验。所选定的扰动装置和传感器在喷注器面安装的方案可实现性最好,结构变动最小,可保持试验在原型燃烧室状态下进行,同时扰动效果较好,传感器敏感性较好。%In order to carry out dynamic combustion stability rating of a LOX/LH2 rocket engine, the schemes of stability rating for the LOX/LH2 rocket engine are investigated based on the stability rating datum of some rocket engines and the basic criteria of CPIA655, including the evaluation standard, testing method, disturbance method, dynamic pressure testing and structure design. Compared to other schemes, the selected scheme that disturbance device and high frequency pressure sensors install on injector surface has better feasibility and less structural changes, which could ensure the rating test to be carried in a prototype engine, thereby leading to better disturbance efficiency and measure sensitivity.

  8. Mars Rocket Propulsion System

    Science.gov (United States)

    Zubrin, Robert; Harber, Dan; Nabors, Sammy

    2008-01-01

    A report discusses the methane and carbon monoxide/LOX (McLOx) rocket for ascent from Mars as well as other critical space propulsion tasks. The system offers a specific impulse over 370 s roughly 50 s higher than existing space-storable bio-propellants. Current Mars in-situ propellant production (ISPP) technologies produce impure methane and carbon monoxide in various combinations. While separation and purification of methane fuel is possible, it adds complexity to the propellant production process and discards an otherwise useful fuel product. The McLOx makes such complex and wasteful processes unnecessary by burning the methane/CO mixtures produced by the Mars ISPP systems without the need for further refinement. Despite the decrease in rocket-specific impulse caused by the CO admixture, the improvement offered by concomitant increased propellant density can provide a net improvement in stage performance. One advantage is the increase of the total amount of propellant produced, but with a decrease in mass and complexity of the required ISPP plant. Methane/CO fuel mixtures also may be produced by reprocessing the organic wastes of a Moon base or a space station, making McLOx engines key for a human Lunar initiative or the International Space Station (ISS) program. Because McLOx propellant components store at a common temperature, very lightweight and compact common bulkhead tanks can be employed, improving overall stage performance further.

  9. Rocket propulsion elements

    CERN Document Server

    Sutton, George P

    2011-01-01

    The definitive text on rocket propulsion-now revised to reflect advancements in the field For sixty years, Sutton's Rocket Propulsion Elements has been regarded as the single most authoritative sourcebook on rocket propulsion technology. As with the previous edition, coauthored with Oscar Biblarz, the Eighth Edition of Rocket Propulsion Elements offers a thorough introduction to basic principles of rocket propulsion for guided missiles, space flight, or satellite flight. It describes the physical mechanisms and designs for various types of rockets' and provides an unders

  10. Replacement of chemical rocket launchers by beamed energy propulsion.

    Science.gov (United States)

    Fukunari, Masafumi; Arnault, Anthony; Yamaguchi, Toshikazu; Komurasaki, Kimiya

    2014-11-01

    Microwave Rocket is a beamed energy propulsion system that is expected to reach space at drastically lower cost. This cost reduction is estimated by replacing the first-stage engine and solid rocket boosters of the Japanese H-IIB rocket with Microwave Rocket, using a recently developed thrust model in which thrust is generated through repetitively pulsed microwave detonation with a reed-valve air-breathing system. Results show that Microwave Rocket trajectory, in terms of velocity versus altitude, can be designed similarly to the current H-IIB first stage trajectory. Moreover, the payload ratio can be increased by 450%, resulting in launch-cost reduction of 74%.

  11. Effects of baffle on combustion acoustic characteristics of liquid rocket engine%隔板对燃烧室声学特性的影响

    Institute of Scientific and Technical Information of China (English)

    李丹琳; 田原; 孙纪国

    2012-01-01

    为了研究液体火箭发动机燃烧室出现的横向一阶切向燃烧不稳定,通过冷态声学试验和理论算例的计算,研究了不同参数的隔板装置对一阶切向声学频率及阻尼特性的影响,结果表明:增加轴向隔板长度和径向隔板数目均会降低一阶切向声学频率,同时增强声阻尼效果;喷嘴式隔板产生的声阻尼效果,比典型直板形状的隔板要好得多,隔板喷嘴最佳间隙在0.1-0.4mm,采用最佳隔板喷嘴间隙能够在较短的轴向隔板长度上得到较高的阻尼能力,从而改善冷却问题.%Cold acoustic tests have been performed to elucidate the effect of baffle on the damping characteristics of the first-tangential acoustic mode in a liquid rocket engine. Differ- ent kinds of baffle parameters were researched by acoustic tests. The results agree well with the theory typical example and show that when increasing the axial baffle length and the ra- dial baffle number, the acoustic frequency of the first-tangential acoustic mode decreases and the acoustic damping capacity increases. Injector-forming baffles have some advantages over the typical straight baffles in acoustic damping capability; an optimal acoustic damping ca- pacitance has been achieved in 0.1-0. 4mm; axial baffle length can be reduced by using the optimal baffle gap, providing a possible solution of thermal cooling problems.

  12. Mini-Rocket User Guide

    Science.gov (United States)

    2007-08-01

    Missile Research , Development, and Engineering Center and Ray Sells DESE Research , Inc. 315 Wynn Drive Huntsville, AL 35805 August 2007...with the minirock command, you are prompted for a filename: Mini-Rocket v1.01 by Ray Sells, DESE Research , Inc. Input file: - Output is printed...nancv.bucher@us.army.mil Commander, U.S. Army ARDEC Picatinny Arsenal, NJ 07806-5000 ATTN: AMSRD-AR-AIS -SA DESE Research , Inc. 3 15 Wynn Drive

  13. Engineering

    National Research Council Canada - National Science Library

    Includes papers in the following fields: Aerospace Engineering, Agricultural Engineering, Chemical Engineering, Civil Engineering, Electrical Engineering, Environmental Engineering, Industrial Engineering, Materials Engineering, Mechanical...

  14. Solar Thermal Rocket Propulsion

    Science.gov (United States)

    Sercel, J. C.

    1986-01-01

    Paper analyzes potential of solar thermal rockets as means of propulsion for planetary spacecraft. Solar thermal rocket uses concentrated Sunlight to heat working fluid expelled through nozzle to produce thrust.

  15. 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

  16. The 2003 Goddard Rocket Replica Project: A Reconstruction of the World's First Functional Liquid Rocket System

    Science.gov (United States)

    Farr, R. A.; Elam, S. K.; Hicks, G. D.; Sanders, T. M.; London, J. R.; Mayne, A. W.; Christensen, D. L.

    2003-01-01

    As a part of NASA s 2003 Centennial of Flight celebration, engineers and technicians at Marshall Space Flight Center (MSFC), Huntsville, Alabama, in cooperation with the Alabama-Mississippi AIAA Section, have reconstructed historically accurate, functional replicas of Dr. Robert H. Goddard s 1926 first liquid- fuel rocket. The purposes of this project were to clearly understand, recreate, and document the mechanisms and workings of the 1926 rocket for exhibit and educational use, creating a vital resource for researchers studying the evolution of liquid rocketry for years to come. The MSFC team s reverse engineering activity has created detailed engineering-quality drawings and specifications describing the original rocket and how it was built, tested, and operated. Static hot-fire tests, as well as flight demonstrations, have further defined and quantified the actual performance and engineering actual performance and engineering challenges of this major segment in early aerospace history.

  17. Preliminary experimental investigation on pulse detonation rocket engine with central cone configuration%中心锥体结构脉冲爆震火箭发动机初步实验

    Institute of Scientific and Technical Information of China (English)

    严宇; 范玮; 王可; 穆杨

    2011-01-01

    In order to improve the atomization of liquid fuel and mixing of reactants in side the pulse detonation rocket engine using liquid fuel, a pulse detonation rocket engine with a different configuration was invented. A central cone instead of Shchelkin spiral was used in this engine. Reactants could be injected into the engine both through the engine head and the central cone. With kerosene used as fuel, oxygen as oxidizer and nitrogen as purge gas, fully developed detonation waves were generated in this engine, which could operate steadily on multi cycle mode. The result also indicates that this engine could greatly shorten the DDT (deflagration to detonation transition) run-up distance, and the DDT ruwup distance is approximately five times of the inner diameter of detonation tube. Compared with the approach of installing Shchelkin spiral in the detonation tube as DDT enhancement de vice, the DDT run-up distance of this engine was shortened by 57.5%.%为了改善采用液态燃料的脉冲爆震火箭发动机内部燃料的雾化以及燃料混合物的掺混状况,采用了一种中心锥体结构.该结构发动机不采用Shchelkin螺旋增爆装置,而采用中心锥体结构、二级供应方式.采用航空煤油为燃料、压缩氧气为氧化剂、压缩氮气为隔离气体,在该结构脉冲爆震火箭发动机上获得了充分发展的爆震波并且能够在多循环条件下稳定工作.实验结果表明,该结构可以大大缩短DDT(deflagra-tion to detonation transition)距离,在实验条件下爆燃向爆震转变距离约为管径的5倍.较之同一管径采用Shchelkin螺旋增爆装置的脉冲爆震火箭发动机,该结构发动机的爆燃向爆震转变距离缩短了57.5%.

  18. Digital Machining System for Nozzle Cooling Channel of Large Liquid Rocket Engine%大型液体火箭发动机喷管数字化铣槽加工系统

    Institute of Scientific and Technical Information of China (English)

    王永青; 刘海波; 李护林; 贾振元

    2012-01-01

    Rocket nozzle is a key structural part of high-thrust liquid rocket engine. There are a hundreds of cooling channels around the nozzle, to ensure the reliable cooling and preheat the fuel. However, it is very difficult to machine the cooling channel due to large size, complex profile, low rigidity, etc. In this article, an integrated digital method for cooling channel machining composed of profile measuring, data processing and channel milling is proposed. Because of large difference between the actual contour and the design model, the channel bottom should be redesigned by using the measured geometric information. Therefore, the varying-thickness and varying-depth cooling channel of nozzle with high order contour or parametric shapes can be machined. Further, a special digital machining system is developed based on an open numerical control platform for the dual-channel vertical milling machine. Finally, an experiment utilizing a typical rocket nozzle is implemented to verify the feasibility of the system. It has been proved that the digital machining system can meet the machining requirements for liquid rocket engine nozzle.%针对大型液体火箭发动机喷管几何尺寸大、廓形复杂、结构刚度低致使其冷却通道加工质量难以保证的难题,提出一种集“测量-数据处理-铣槽”于一体的喷管冷却通道数字化加工新方法,并在开放式数控平台上开发出喷管专用数字化铣槽加工系统.该方法利用喷管几何外廓的实际测量信息再设计出槽底曲面,进而实现高次曲线或参数曲线廓形、变壁厚变槽深喷管冷却通道的数字化加工.通过某型号火箭发动喷管的实际加工,表明所研制的双通道立式铣槽加工专用装备与系统可满足我国新一代大推力液体火箭发动机喷管冷却通道高质量、高效、高可靠的制造要求.

  19. Solution-chemical route to generalized synthesis of metal germanate nanowires with room-temperature, light-driven hydrogenation activity of CO2 into renewable hydrocarbon fuels.

    Science.gov (United States)

    Liu, Qi; Zhou, Yong; Tu, Wenguang; Yan, Shicheng; Zou, Zhigang

    2014-01-06

    A facile solution-chemical route was developed for the generalized preparation of a family of highly uniform metal germanate nanowires on a large scale. This route is based on the use of hydrazine monohydrate/H2O as a mixed solvent under solvothermal conditions. Hydrazine has multiple effects on the generation of the nanowires: as an alkali solvent, a coordination agent, and crystal anisotropic growth director. Different-percentage cobalt-doped Cd2Ge2O6 nanowires were also successfully obtained through the addition of Co(OAc)2·4H2O to the initial reaction mixture for future investigation of the magnetic properties of these nanowires. The considerably negative conduction band level of the Cd2Ge2O6 nanowire offers a high driving force for photogenerated electron transfer to CO2 under UV-vis illumination, which facilitates CO2 photocatalytic reduction to a renewable hydrocarbon fuel in the presence of water vapor at room temperature.

  20. 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

  1. 火箭发动机随机推力调节控制驱动器的研制%Research on random thrust adjustable controller of rocket engine

    Institute of Scientific and Technical Information of China (English)

    马兵兵; 翟丽婷; 孙璐

    2012-01-01

    为满足某型号液体火箭发动机定混合比随机无极变推力工作要求,研制了基于DSP处理器的随机推力调节控制驱动器。该控制驱动器实时接收随机变推力指令,在定混合比条件下,协调控制发动机系统上的燃料及氧化剂路调节阀,从而控制燃料及氧化剂流量,完成发动机的随机变推力控制。其参加多次发动机系统冷调试验及地面全程热试车,工作稳定可靠,实现了变推力双组元推进剂流量同步控制,精确控制发动机混合比,快速响应随机变推力控制要求。%To meet the thrust control requirements of a liquid rocket engine,a random thrust adjustable controller based on DSP is developed.It receives random variable thrust instructions and varies the engine thrust accordingly by means of controlling the fuel and oxidant valves,while the mixture ratio is fixed.The controller showed good performances during cold-flow tests and full-duration hot firing tests.With the high stability and reliability,the controller achieved the synchronization control of variable thrust bipropellant flows in liquid rocket engine,in which the mixture ratio was precisely controlled,and swift response to random variable thrust control demand was realyzed.

  2. Hybrid Rocket Technology

    National Research Council Canada - National Science Library

    Sankaran Venugopal; K K Rajesh; V Ramanujachari

    2011-01-01

    With their unique operational characteristics, hybrid rockets can potentially provide safer, lower-cost avenues for spacecraft and missiles than the current solid propellant and liquid propellant systems...

  3. Application of C/C Composites in Rocket Engine Nozzles in Japan%日本火箭发动机喷管用C/C复合材料

    Institute of Scientific and Technical Information of China (English)

    李崇俊; 崔红; 李瑞珍

    2013-01-01

    The application state of C/C composites in solid rocket motor nozzle in Japan.The components include a rosetta carbon fabric laminated 2D-C/C exit cone for satellite apogee boost motor,and 3D-C/C throat inserts for solid rocket booster and launch vehicle.The rayon based carbon fiber is adopted to make the 2D-C/C exit cone.The 3D-C/C throat insert,used as the 1st stage of M-V solid rocket launch vehicle,has a dimension of Фll00 mm in outer diameter and a density of 1.95 g/em3.The 3 D preform is orthogonally weaved in a cylinder structure,and then densified by a repeating heat isostatic pressure-graphitization cycles.Applications of C/C composites in both solid and liquid rocket motor nozzle extendable exit cones are future development trend in this area.%介绍了C/C复合材料在日本固体火箭发动机喷管的应用情况,主要包括卫星远地点助推发动机用螺旋形状碳布铺层的2D-C/C扩张段、固体助推器及固体运载用3D-C/C喉衬.2D-C/C扩张段采用黏胶丝基碳纤维成型,M-V固体运载一级发动机C/C喉衬采用碳纤维三向正交圆筒编织结构,热等静压-石墨化致密,外径Ф1100 mm,密度达1.95 g/cm3.C/C复合材料在固体及液体火箭发动机喷管延伸出口锥的应用是未来的发展方向.

  4. Not just rocket science

    Energy Technology Data Exchange (ETDEWEB)

    MacAdam, S.; Anderson, R. [Celan Energy Systems, Rancho Cordova, CA (United States)

    2007-10-15

    The paper explains a different take on oxyfuel combustion. Clean Energy Systems (CES) has integrated aerospace technology into conventional power systems, creating a zero-emission power generation technology that has some advantages over other similar approaches. When using coal as a feedstock, the CES process burns syngas rather than raw coal. The process uses recycled water and steam to moderate the temperature, instead of recycled CO{sub 2}. With no air ingress, the CES process produces very pure CO{sub 2}. This makes it possible to capture over 99% of the CO{sub 2} resulting from combustion. CES uses the combustion products to drive the turbines, rather than indirectly raising steam for steam turbines, as in the oxyfuel process used by companies such as Vattenfall. The core of the process is a high-pressure oxy-combustor adapted from rocket engine technology. This combustor burns gaseous or liquid fuels with gaseous oxygen in the presence of water. Fuels include natural gas, coal or coke-derived synthesis gas, landfill and biodigester gases, glycerine solutions and oil/water emulsion. 2 figs.

  5. LDRD final report on "fundamentals of synthetic conversion of CO2 to simple hydrocarbon fuels" (LDRD 113486).

    Energy Technology Data Exchange (ETDEWEB)

    Maravelias, Christos T. (University of Wisconsin, Madison, WI); Kemp, Richard Alan; Mavrikakis, Manos (University of Wisconsin, Madison, WI); Miller, James Edward; Stewart, Constantine A.

    2009-11-01

    Energy production is inextricably linked to national security and poses the danger of altering the environment in potentially catastrophic ways. There is no greater problem than sustainable energy production. Our purpose was to attack this problem by examining processes, technology, and science needed for recycling CO{sub 2} back into transportation fuels. This approach can be thought of as 'bio-inspired' as nature employs the same basic inputs, CO{sub 2}/energy/water, to produce biomass. We addressed two key deficiencies apparent in current efforts. First, a detailed process analysis comparing the potential for chemical and conventional engineering methods to provide a route for the conversion of CO{sub 2} and water to fuel has been completed. No apparent 'showstoppers' are apparent in the synthetic route. Opportunities to improve current processes have also been identified and examined. Second, we have also specifically addressed the fundamental science of the direct production of methanol from CO{sub 2} using H{sub 2} as a reductant.

  6. The Ion Rocket

    Science.gov (United States)

    1961-05-29

    discharge velocity w and the speci- fic impulse lap respectively cannot be increased. At this limit condition the thermal rocket oecouos "choked up...structural quality is 900 t, 3) In the case of an atomic-driven thermal rocket ’,;lth specific Ipipulse ISjy«8C0 sec and thrust to weight ratio « 1, the

  7. Model Rockets and Microchips.

    Science.gov (United States)

    Fitzsimmons, Charles P.

    1986-01-01

    Points out the instructional applications and program possibilities of a unit on model rocketry. Describes the ways that microcomputers can assist in model rocket design and in problem calculations. Provides a descriptive listing of model rocket software for the Apple II microcomputer. (ML)

  8. Infrared spectroradiometer for rocket exhaust analysis

    Science.gov (United States)

    Herget, W. F.

    1968-01-01

    Infrared spectroradiometer measures high-resolution spectral absorption, emission, temperature, and concentration of chemical species in radically symmetric zones of the exhaust plumes of large rocket engines undergoing static firing tests. Measurements are made along predetermined lines of sight through the plume.

  9. Gaseous Helium Reclamation at Rocket Test Systems Project

    Data.gov (United States)

    National Aeronautics and Space Administration — GHe reclamation is critical in reducing operating costs at rocket engine test facilities. Increases in cost and shortages of helium will dramatically impact testing...

  10. Process Design and Economics for the Conversion of Lignocellulosic Biomass to Hydrocarbon Fuels: Thermochemical Research Pathways with In Situ and Ex Situ Upgrading of Fast Pyrolysis Vapors

    Energy Technology Data Exchange (ETDEWEB)

    Dutta, Abhijit [National Renewable Energy Lab. (NREL), Golden, CO (United States); Sahir, A. H. [National Renewable Energy Lab. (NREL), Golden, CO (United States); Tan, Eric [National Renewable Energy Lab. (NREL), Golden, CO (United States); Humbird, David [DWH Process Consulting, Denver, CO (United States); Snowden-Swan, Lesley J. [Pacific Northwest National Lab. (PNNL), Richland, WA (United States); Meyer, Pimphan A. [Pacific Northwest National Lab. (PNNL), Richland, WA (United States); Ross, Jeff [Harris Group, Inc., Seattle, WA (United States); Sexton, Danielle [Harris Group, Inc., Seattle, WA (United States); Yap, Raymond [Harris Group, Inc., Seattle, WA (United States); Lukas, John [Harris Group, Inc., Seattle, WA (United States)

    2015-03-01

    This report was developed as part of the U.S. Department of Energy’s Bioenergy Technologies Office’s efforts to enable the development of technologies for the production of infrastructure-compatible, cost-competitive liquid hydrocarbon fuels from biomass. Specifically, this report details two conceptual designs based on projected product yields and quality improvements via catalyst development and process integration. It is expected that these research improvements will be made within the 2022 timeframe. The two conversion pathways detailed are (1) in situ and (2) ex situ upgrading of vapors produced from the fast pyrolysis of biomass. While the base case conceptual designs and underlying assumptions outline performance metrics for feasibility, it should be noted that these are only two of many other possibilities in this area of research. Other promising process design options emerging from the research will be considered for future techno-economic analysis. Both the in situ and ex situ conceptual designs, using the underlying assumptions, project MFSPs of approximately $3.5/gallon gasoline equivalent (GGE). The performance assumptions for the ex situ process were more aggressive with higher distillate (diesel-range) products. This was based on an assumption that more favorable reaction chemistry (such as coupling) can be made possible in a separate reactor where, unlike in an in situ upgrading reactor, one does not have to deal with catalyst mixing with biomass char and ash, which pose challenges to catalyst performance and maintenance. Natural gas was used for hydrogen production, but only when off gases from the process was not sufficient to meet the needs; natural gas consumption is insignificant in both the in situ and ex situ base cases. Heat produced from the burning of char, coke, and off-gases allows for the production of surplus electricity which is sold to the grid allowing a reduction of approximately 5¢/GGE in the MFSP.

  11. Effect of strain rate on sooting limits in counterflow diffusion flames of gaseous hydrocarbon fuels: Sooting temperature index and sooting sensitivity index

    KAUST Repository

    Wang, Yu

    2014-05-01

    The effect of the strain rate on the sooting limits in counterflow diffusion flames was investigated in various gaseous hydrocarbon fuels by varying the nitrogen dilution in the fuel and oxidizer streams. The sooting limit was defined as the critical fuel and oxygen mole fraction at which soot started to appear in the elastic light scattering signal. The sooting region for normal alkane fuels at a specified strain rate, in terms of the fuel and oxygen mole fraction, expanded as the number of carbon atoms increased. The alkene fuels (ethylene, propene) tested had a higher propensity for sooting as compared with alkane fuels with the same carbon numbers (ethane, propane). Branched iso-butane had a higher propensity for sooting than did n-butane. An increase in the strain rate reduced the tendency for sooting in all the fuels tested. The sensitivity of the sooting limit to the strain rate was more pronounced for less sooting fuels. When plotted in terms of calculated flame temperature, the critical oxygen mole fraction exhibited an Arrhenius form under sooting limit conditions, which can be utilized to significantly reduce the effort required to determine sooting limits at different strain rates. We found that the limiting temperatures of soot formation flames are viable sooting metrics for quantitatively rating the sooting tendency of various fuels, based on comparisons with threshold soot index and normalized smoke point data. We also introduce a sooting temperature index and a sooting sensitivity index, two quantitative measures to describe sooting propensity and its dependence on strain rate. © 2013 The Combustion Institute.

  12. HZSM-5/MCM-41 composite molecular sieves for the catalytic cracking of endothermic hydrocarbon fuels: nano-ZSM-5 zeolites as the source

    Science.gov (United States)

    Sang, Yu; Jiao, Qingze; Li, Hansheng; Wu, Qin; Zhao, Yun; Sun, Kening

    2014-12-01

    A series of HZSM-5/MCM-41 composite molecular sieves (HZM-Ns ( x)) were prepared by employing nano-ZSM-5 zeolites with the SiO2/Al2O3 ratios ( x) of 50, 100 and 150 as the source. These materials were characterized by X-ray diffraction, Fourier transform infrared spectroscopy, scanning electron microscopy, energy dispersive spectroscopy, transmission electron microscopy, N2 adsorption-desorption measurement, and NH3 temperature-programmed desorption. The catalytic cracking of endothermic hydrocarbon fuels over the HZM-Ns with n-decane as model was evaluated at atmospheric pressure and 500 °C. The effect of the parent zeolite, mesopore and SiO2/Al2O3 ratio on the structure, acidity, and catalytic performance of HZM-Ns was investigated. The HZM-Ns exhibited a skeletal matrix with nano-sized HZSM-5 particles (200-300 nm) with a controllable acidity well dispersed in and microporous-mesoporous hierarchical pores. The mesoporous structure improved the diffusion of the reactants and products in the pores, and the HZSM-5 nanoparticles uniformly dispersed in the MCM-41 matrix supplied a proper acidity, shorter channels, and a higher specific surface area for reaction. These resulted in a high catalytic activity, a high selectivity to light olefins and a long lifetime for n-decane catalytic cracking. The HZM-N (150) exhibited the excellent conversion, a high selectivity to light olefins and a long lifetime due to low diffusion resistance, high specific surface area, and appropriate acid distribution and strength, with the increasing SiO2/Al2O3 ratio.

  13. 固体火箭发动机自动回转系统的设计与实现%Design and Implementation of Auto-rotation System for Solid Rocket Engine

    Institute of Scientific and Technical Information of China (English)

    赵锴; 何敏; 于殿泓; 郑毅

    2011-01-01

    In the process of high energy X-ray radiography detection, the disadvantages of low control accuracy and difficult radiation-proof for operators exist in manually controlling the rotation of solid rocket engine. Thus, the auto-rotation system based on OMRON CQM1H PLC has been designed. Two operating modes: auto and manual are equipped in this system to implement remote and high accurate automatic rotating function for solid rocket engine. The practice shows that the system features high stability and reliability, ease maintenance, and satisfies the requirements of explosion-proof, safety and high reliability.%在对固体火箭发动机进行高能X射线照相检测的过程中,针对采用人工方式存在回转固体火箭发动机存在控制精度不高、人员辐射防护困难等问题,设计了一种基于OMRON CQM1H PLC的自动回转系统.系统具备手动和自动两种运行模式,实现了固体火箭发动机的远程、高精度和自动化回转控制功能.实际应用表明,系统稳定性好、可靠性高且易于维护,符合检测现场防爆安全和高可靠性的要求.

  14. Research progresses on turbulent mixing and combustion for air-turbo-rocket engine%空气涡轮火箭发动机掺混燃烧研究进展

    Institute of Scientific and Technical Information of China (English)

    李文龙; 李平; 郭海波

    2011-01-01

    It is absolutely crucial for the performance of air-turbo-rocket engine in which forms an efficient and steady mixing combustion of air and fuel-rich gas in the combustion chamber. The experimental investigations on schemes of turbulent mixing and combustion for air-turbo-rocket engine, which are proceeded by worldwide institutes, are reviewed. Typical applications and various research achievements of aeronautic lobed mixers for enhancing molecular mixing between bypass and core flows are assessed. An essential summary of characteristics and key problems about turbulent mixing and combustion is carried out. Furthermore, the focuses on turbulent mixing and combustion in subsequent researches which should be paid extra attention are analyzed.%在混流燃烧室内组织富燃燃气与空气的高效稳定掺混燃烧对于空气涡轮火箭发动机(ATR)性能至关重要。回顾了国外各研究机构关于ATR掺混燃烧方案的试验研究,对波瓣混流器在航空领域强化内、外涵气流掺混中的典型应用及研究成果进行总结评述,总结并提出ATR掺混燃烧的特点和关键问题,分析了后续掺混燃烧研究中需重点关注的问题。

  15. Simulation of 3-D Ultraviolet Radiation from Liquid Rocket Engine Plume%液体火箭发动机羽烟三维紫外辐射仿真研究

    Institute of Scientific and Technical Information of China (English)

    国爱燕; 唐义; 白廷柱; 黄刚

    2012-01-01

    针对液体火箭发动机羽烟紫外辐射的空间分布问题,建立三维数值计算模型.该模型采用标准κ-ε湍流模型和PDF模型仿真羽烟流场的状态参数,根据HITRAN数据库计算流场内吸收系数分布,并利用离散坐标法求解辐射传输方程,计算三维空间的紫外辐射分布.测试结果表明:三维紫外辐射模型的计算结果与实验数据一致,能够反应不同视角下液体火箭发动机羽烟紫外辐射强度的空间变化.%To analyze the spacial distribution of UV radiation from liquid rocket engine plume, a 3-D numerical model has been built. The standard κ -e model and probability density function (PDF) model were adopted to compute the flow-field properties of the plume. Combined with HITRAN database, the distribution of absorption coefficient was calculated, and the radiation transfer equation could be solved by the discrete ordinate method. Test results show that the 3-D computation model can provide numerical data that agree well with measured experimental data. It also could reflect the spacial distribution of UV radiation from the hydrogen-oxygen rocket engine plume at different angles of view.

  16. 轨姿控液体火箭发动机水击仿真模拟%Simulation of water hammer in liquid rocket engine of orbit and attitude control system

    Institute of Scientific and Technical Information of China (English)

    张峥岳; 康乃全

    2012-01-01

    Taking the liquid rocket engine of orbit and attitude control system as the study object, an emulator was established with AMESim according to the modular modeling idea. The simulation computation of water hammer pressure in the pipeline while the engine system was working was per- formed. The results show that the running of orbit control engine is a major factor creating high water hammer. The compared result of theoretical calculation and test data indicate that the simulation mod- els can give reasonable descriptions for generative process of water hammer. The measure to reduce the amount of water hammer is introduced.%以轨姿控液体火箭发动机为研究对象,根据模块化思想,利用AMESim建立了仿真平台,仿真计算了发动机系统工作中管路的水击压力。结果表明:轨控发动机的工作是引起大水击的主要因素。通过与理论计算和试验数据的对比表明,仿真模型较好地描述了管路水击的生成过程。介绍了减小系统水击量的措施。

  17. CODEX sounding rocket wire grid collimator design

    Science.gov (United States)

    Shipley, Ann; Zeiger, Ben; Rogers, Thomas

    2011-05-01

    CODEX is a sounding rocket payload designed to operate in the soft x-ray (0.1-1.0 kV) regime. The instrument has a 3.25 degree square field of view that uses a one meter long wire grid collimator to create a beam that converges to a line in the focal plane. Wire grid collimator performance is directly correlated to the geometric accuracy of actual grid features and their relative locations. Utilizing a strategic combination of manufacturing and assembly techniques, this design is engineered for precision within the confines of a typical rocket budget. Expected resilience of the collimator under flight conditions is predicted by mechanical analysis.

  18. Heterogeneous fuel for hybrid rocket

    Science.gov (United States)

    Stickler, David B. (Inventor)

    1996-01-01

    Heterogeneous fuel compositions suitable for use in hybrid rocket engines and solid-fuel ramjet engines, The compositions include mixtures of a continuous phase, which forms a solid matrix, and a dispersed phase permanently distributed therein. The dispersed phase or the matrix vaporizes (or melts) and disperses into the gas flow much more rapidly than the other, creating depressions, voids and bumps within and on the surface of the remaining bulk material that continuously roughen its surface, This effect substantially enhances heat transfer from the combusting gas flow to the fuel surface, producing a correspondingly high burning rate, The dispersed phase may include solid particles, entrained liquid droplets, or gas-phase voids having dimensions roughly similar to the displacement scale height of the gas-flow boundary layer generated during combustion.

  19. 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…

  20. 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…

  1. 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

  2. Nuclear rocket using indigenous Martian fuel NIMF

    Science.gov (United States)

    Zubrin, Robert

    1991-01-01

    In the 1960's, Nuclear Thermal Rocket (NTR) engines were developed and ground tested capable of yielding isp of up to 900 s at thrusts up to 250 klb. Numerous trade studies have shown that such traditional hydrogen fueled NTR engines can reduce the inertial mass low earth orbit (IMLEO) of lunar missions by 35 percent and Mars missions by 50 to 65 percent. The same personnel and facilities used to revive the hydrogen NTR can also be used to develop NTR engines capable of using indigenous Martian volatiles as propellant. By putting this capacity of the NTR to work in a Mars descent/acent vehicle, the Nuclear rocket using Indigenous Martian Fuel (NIMF) can greatly reduce the IMLEO of a manned Mars mission, while giving the mission unlimited planetwide mobility.

  3. 基于工作流的液体火箭发动机虚拟试验流程管理%Workflow-based process management in virtual test of liquid rocket engine

    Institute of Scientific and Technical Information of China (English)

    段娜; 朱子环; 于海磊; 张黎辉

    2012-01-01

    以液体火箭发动机虚拟试验为对象,研究了数字化试验流程管理的解决方案。通过分析发动机虚拟试验流程,建立了基于工作流的试验管理系统过程仿真模型,提出了一个支撑整个试验管理系统的层次化体系架构。该架构为整个虚拟试验过程提供了统一的试验信息集成平台和应用服务环境。最后,以试验准备阶段为例,给出了该系统的一个集成应用,介绍了基于图形管理系统的实现过程,得到了过程中资源消耗优化的结论。所提出的液体火箭发动机虚拟试验流程管理体系为真实试验的资源优化提供了理论依据,在航天领域的数字化试验方面进行了探索性研究。%The digitalized test process management in the virtual test of liquid rocket engines is studied. By analyzing the virtual test process of such engines, the workflow-based process simulation model of test management system is established, and a hierarchical architecture is proposed to support the test process management system. The architecture provides a unified test information integrated platform and an application service environment for the entire virtual test process. Taking preparatory stage of the test as an example, the integrated application of the the management system is shown, and the implementation process based on graphics management system is introduced. A conclusion of optimized consumption of resources is achieved. The test process management system presented in this paper for the virtual test of liquid rocket engines can provide a theoretic foundation for resource utilization in real test. It conduces to farther research in digitalized test process management in the field of space technology.

  4. 液氧煤油发动机稳态故障仿真分析%Steady-state fault simulation and effect analysis of LOX/kerosene rocket engine

    Institute of Scientific and Technical Information of China (English)

    党锋刚; 马红宇; 李春红; 宋春

    2012-01-01

    根据液氧煤油补燃循环发动机的特点,建立了稳态工作过程故障仿真数学模型,并针对比较典型的几种故障模式,进行了仿真计算与效应分析,最后进行了故障参数特征的初步提取。结果表明,选定的10个缓变热力参数,可对泄漏、堵塞及涡轮泵等典型故障模式进行有效识别和分离。%The simulation mathematical model of fault occurring in the steady-state working process is built and simulation software is designed according to the characteristics of LOX/kerosene staged combustion cycled rocket engine.Several typical failure modes are simulatively calculated.The effect of the fault is analyzed and characteristics of the fault modes are extracted.Ten slow variable thermal parameters were selected on the basis of the anaysis result to identify the faut modes of leakage and jamming in the fuel pipe of engine.

  5. Scaled Rocket Testing in Hypersonic Flow

    Science.gov (United States)

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

    2015-01-01

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

  6. Design of Turning Equipment of Niobium Ring-ceramic Matrix Composites Rocket Engine Thrusters%铌环-C/SiC复合材料喷管的车削工装设计

    Institute of Scientific and Technical Information of China (English)

    黎明; 顾力强; 郭洪勤

    2014-01-01

    A model of rocket engine adopts the ceramic matrix composites nozzle, on which a niobium ring was deposited to connect to the engine by means of electron beam welding. For turning the welding stage of niobium ring, this paper introduces a kind of process equipment design train of thought. By using combining process equipment to find new positioning base, it effectively solved the problem of the fracture of ceramic matrix composites. This technology completely satisfied the requirements of tolerance of welding stage by turning results. And it improves the performance of the engine much further.%某火箭发动机采用 C/SiC复合材料的喷管,并在喷管上沉积一个带焊接头的铌环,通过该铌环与发动机金属头部进行电子束焊接。现需要车削完成喷管铌环的焊接台阶等特征。本文介绍了一种工装设计的新思路,即通过组合车削工装寻找定位基准。通过实际加工证明:该工艺技术完全满足焊接台阶的形位公差要求,并有效地解决了 C/SiC 复合材料脆性断裂等难题,最终满足了发动机的性能指标。

  7. Introduction to the Special Issue on Sounding Rockets and Instrumentation

    CERN Document Server

    Christe, Steven; 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 importantly, sounding rockets remain the only way to explore the tenuous regions of the Earth's atmosphere (the upper stratosphere, mesosphere, and lower ionosphere/thermosphere) above balloon altitudes ($\\sim$40 km) and below satellite orbits ($\\sim$160 km). They can lift remote sensing telescope payloads with masses up to 400 kg to altitudes of 350 km providing observing times of up to 6 minutes above the blocking influence of Earth's atmosphere. Though a number of sounding rocket research programs exist around the world, th...

  8. Thermo-acoustic instabilities of high-frequency combustion in rocket engines; Instabilites thermo-acoustiques de combustion haute-frequence dans les moteurs fusees

    Energy Technology Data Exchange (ETDEWEB)

    Cheuret, F.

    2005-10-15

    Rocket motors are confined environments where combustion occurs in extreme conditions. Combustion instabilities can occur at high frequencies; they are tied to the acoustic modes of the combustion chamber. A common research chamber, CRC, allows us to study the response of a turbulent two-phase flame to acoustic oscillations of low or high amplitudes. The chamber is characterised under cold conditions to obtain, in particular, the relative damping coefficient of acoustic oscillations. The structure and frequency of the modes are determined in the case where the chamber is coupled to a lateral cavity. We have used a powder gun to study the response to a forced acoustic excitation at high amplitude. The results guide us towards shorter flames. The injectors were then modified to study the combustion noise level as a function of injection conditions. The speed of the gas determines whether the flames are attached or lifted. The noise level of lifted flames is higher. That of attached flames is proportional to the Weber number. The shorter flames whose length is less than the radius of the CRC, necessary condition to obtain an effective coupling, are the most sensitive to acoustic perturbations. The use of a toothed wheel at different positions in the chamber allowed us to obtain informations on the origin of the thermo-acoustic coupling, main objective of this thesis. The flame is sensitive to pressure acoustic oscillations, with a quasi-zero response time. These observations suggest that under the conditions of the CRC, we observe essentially the response of chemical kinetics to pressure oscillations. (author)

  9. 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.

  10. Effects of high combustion chamber pressure on rocket noise environment

    Science.gov (United States)

    Pao, S. P.

    1972-01-01

    The acoustical environment for a high combustion chamber pressure engine was examined in detail, using both conventional and advanced theoretical analysis. The influence of elevated chamber pressure on the rocket noise environment was established, based on increase in exit velocity and flame temperature, and changes in basic engine dimensions. Compared to large rocket engines, the overall sound power level is found to be 1.5 dB higher, if the thrust is the same. The peak Strouhal number shifted about one octave lower to a value near 0.01. Data on apparent sound source location and directivity patterns are also presented.

  11. Flow separation in rocket nozzles under high altitude condition

    Science.gov (United States)

    Stark, R.; Génin, C.

    2017-01-01

    The knowledge of flow separation in rocket nozzles is crucial for rocket engine design and optimum performance. Typically, flow separation is studied under sea-level conditions. However, this disregards the change of the ambient density during ascent of a launcher. The ambient flow properties are an important factor concerning the design of altitude-adaptive rocket nozzles like the dual bell nozzle. For this reason an experimental study was carried out to study the influence of the ambient density on flow separation within conventional nozzles.

  12. 氢氧火箭发动机射流仿真与试验台热防护%Thermal Protection and Plume Simulation for Hydrogen/Oxygen Rocket Engine Test Stage

    Institute of Scientific and Technical Information of China (English)

    李茂; 王占林

    2014-01-01

    The plume characteristics of the hydrogen/oxygen rocket engine were studied with the method of CFD ( computational fluid dynamics ) .The influences of the geometry model , the combus-tion model and the turbulence model on the characteristics of the combustion flowfield were ana-lyzed .The results from the numerical simulation were compared with those from the experiments qualitatively and quantitatively .Based on the distributions of the temperature from the numerical simulation results in the different tests , the armor plate was designed for the thermal protection and applied for the test stage .It is confirmed that the approach is effective and reliable .%采用计算流体方法获得氢氧火箭发动机地面试验射流特征,开展了几何模型、燃烧模型和湍流模型对射流场的影响分析以及与试验结果的定性和定量对比。依据不同试验模式下的射流场温度的数值分布,提出试验台钢板防护方案并进行防护,试验证明方案可靠有效。

  13. Technological study and characterization of injectors with micro-orifices for bi-liquid rockets engines; Etude technologique et caracterisation d'injecteurs a micro-orifices pour propulseurs biliquides

    Energy Technology Data Exchange (ETDEWEB)

    Prevot, P. [Office National d' Etudes et de Recherches Aerospatiales (ONERA-DMAE/LP), 31 - Mauzac (France); Lecourt, R.; Foucaud, R.; D' Herbigny, F.X. [Office National d' Etudes et de Recherches Aerospatiales (ONERA-DMAE/MH), Centre du Fauga-Mauzac, 31 - Mauzac (France); Hervat, P. [Office National d' Etudes et de Recherches Aerospatiales (ONERA-DMTE/BET), Centre de Chatillon, 92 - Chatillon (France)

    2001-10-01

    Injectors for space rockets engines have been realised an tested at ONERA. Thanks to an original manufacturing technique, they allow one to get pulverization characteristics superior than those obtained with classic injectors. This improvement increases performances, allows a reduction of the size and mass of combustion chambers, while reducing manufacturing costs in the case of small series. These injectors are constituted by very thin plates fabricated by standard or chemical means. Individual plates are then assembled together by diffusion bond or brazing. This mode of fabrication allows one to conceive hydraulic circuits which would be impossible to realize with conventional metal machining. This technique, developed since the 70's in the United States by the AEROJET Company for the fabrication of platelets, complete combustion chambers and heat exchangers, has been adopted at ONERA for developing injectors at the request of CNES. An injector of this type, used at ONERA as a test tool on a very hot air generator, has given complete satisfaction. The industrial development of this technique will necessitate to achieve a technological step at the design level and to improve French capabilities in the diffusion bond process for materials others that titanium, such as stainless steel. (authors)

  14. Rockets in World War I

    Science.gov (United States)

    2004-01-01

    World War I enlisted rockets once again for military purposes. French pilots rigged rockets to the wing struts of their airplanes and aimed them at enemy observation balloons filled with highly inflammable hydrogen.

  15. An Evaluation Of Rocket Parameters

    Directory of Open Access Journals (Sweden)

    J. N. Beri

    1959-07-01

    Full Text Available The dependence of conventional parameters of internal ballistics of Solid Propellant Rockets using external burning cruciform charge, on the geometry of charge aad rocket motor is discussed and results applied in a special case.

  16. 径流式涡轮在膨胀循环发动机氢涡轮泵中的应用%Application of radial turbine in hydrogen turbopump of expander cycle rocket engine

    Institute of Scientific and Technical Information of China (English)

    杨凡; 叶小明

    2012-01-01

    In order to meet the high-performance and high reliability requirements in development of the expander cycle liquid rocket engine, the radial turbine was adopted in the scheme selection of hydrogen turbopumps. The feasibility was preliminarily verified through 1D thermal design and 3D structure design. With CFD analysis sottware, the full 3D viscosity numerical simulation was com- pleted under turbo' s design working conditions, proving that the performance could meet the qualifi- cation requirement. Two main technical problems in turbopump applications were solved through strength optimum design and axial force balance research. In combination with the results obtained in medium test and hot test, a conclusion that the radial turbine can be applied to the hydrogen turbopump of expander cycle engine was achieved.%为满足膨胀循环液体火箭发动机高性能和高可靠的研制要求,在氢涡轮泵方案的选择上采用了径流式氢涡轮方案。通过一维热力和三维结构设计,初步验证了径流式氢涡轮应用可行性。借助于CFD分析软件,完成了该涡轮设计工况全三维粘性数值模拟,证明性能满足指标要求。通过强度优化设计和轴向力平衡两方面研究,突破了涡轮泵应用的两大技术难点。结合该涡轮介质试验及发动机热试车考核情况,得出径流式涡轮能够应用于膨胀循环发动机氢涡轮泵的结论。

  17. Internal combustion engine

    Energy Technology Data Exchange (ETDEWEB)

    Helmich, M.J.; Hoagland, M.C.; Hubbard, R.L.; Schaub, F.S.

    1981-12-22

    A method of combusting natural gas fuel in a two cycle, turbocharged internal combustion engine substantially reduces the production of nitrogen-oxygen emissions. An improved turbocharger design provides increased air charging pressure, produces a controlled lean air/fuel mixture and lowers peak combustion temperatures. A jet cell ignition device ensures uniform, reliable ignition of the lean air/fuel mixture under all operating conditions and the lean air/fuel mixture in turn encourages complete fuel combustion and provides excellent combustion characteristics with methane, ethane and heavier paraffinic hydrocarbon fuels. These structural modifications and adjustment of other operating parameters combine to reduce nitric oxide (NO) and nitrogen dioxide (NO/sub 2/) emissions by as much as 75% while effecting only a negligible increase in fuel consumption.

  18. Engines for the Cosmos

    Science.gov (United States)

    Rodgers, Stephen L.; Reisz, Al; Wyckoff, James (Technical Monitor)

    2002-01-01

    Galactic forces spiral across the cosmos fueled by nuclear fission and fusion and atoms in plasmatic states with throes of constraints of gravitational forces and magnetic fields, In their wanderings these galaxies spew light, radiation, atomic and subatomic particles throughout the universe. Throughout the ages of man visions of journeying through the stars have been wondered. If humans and human devices from Earth are to go beyond the Moon and journey into deep space, it must be accomplished with like forces of the cosmos such as electrical fields, magnetic fields, ions, electrons and energies generated from the manipulation of subatomic and atomic particles. Forms of electromagnetic waves such as light, radio waves and lasers must control deep space engines. We won't get far on our Earth accustomed hydrocarbon fuels.

  19. 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…

  20. 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…

  1. Introduction to Rocket Propulsion

    Science.gov (United States)

    1991-12-01

    Von Braun; 1966. 4. Introduction to Ordnance Technology; IHSP 76-129; 1976. 5. Physics; D. Halliday and R. Resnick ; 1963. 6. Physics Tells Why: 0...to Luke Sky- walker in Star Wars when he said "Don’t get cocky." We never plan for EVERYTHING, though we like to think we do. As we’ve said, rocket

  2. Low toxicity rocket propellants

    NARCIS (Netherlands)

    Wink, J.

    2014-01-01

    Hydrazine (N2H4) and its hypergolic mate nitrogen tetroxide (N2O4) are used on virtually all spacecraft and on a large number of launch vehicles. In recent years however, there has been an effort in identifying and developing alternatives to replace hydrazine as a rocket propellant.

  3. 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…

  4. 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…

  5. Rocketing to the Skies

    Institute of Scientific and Technical Information of China (English)

    1999-01-01

    ONE sunny morning,we startedfor Yanqi Lake,Huairou District,Beijing,to try“rocket bungy”,so farthe only facility for this sport inChina.On the way there,wequestioned our courage and heartendurance. Entering the gate we saw,towering over a banner saying,

  6. Low toxicity rocket propellants

    NARCIS (Netherlands)

    Wink, J.

    2014-01-01

    Hydrazine (N2H4) and its hypergolic mate nitrogen tetroxide (N2O4) are used on virtually all spacecraft and on a large number of launch vehicles. In recent years however, there has been an effort in identifying and developing alternatives to replace hydrazine as a rocket propellant.

  7. 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.

  8. Emission behavior of OH radical in internal EGR using a 2-cycle engine; 2 cycle engine wo mochiita naibu EGR no OH radical no hakko kyodo

    Energy Technology Data Exchange (ETDEWEB)

    Hashimoto, S.; Amino, Y.; Yoshida, K.; Shoji, H.; Saima, A. [Nihon University, Tokyo (Japan)

    1997-10-01

    The purpose of this study was to examine and consider the influence, which the remained gas exercised on combustion. 2-cycle engine was Used as the test engine. Internal EGR was run. The means was that the test engine was fitted the back pressure control plate on the exhaust port. The conditions, which were run with internal EGR and without internal EGR, were compared. The OH radical, which plays important role in combustion of hydrocarbon fuels, was measured with emission spectroscopy. In internal EGR, the unburned end gas on exhaust port side was susceptible to the remained gas. 6 refs., 8 figs., 1 tab.

  9. 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.

  10. Introduction to the Special Issue on Sounding Rockets and Instrumentation

    Science.gov (United States)

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

    2016-03-01

    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 importantly, sounding rockets remain the only way to explore the tenuous regions of the Earth’s atmosphere (the upper stratosphere, mesosphere, and lower ionosphere/thermosphere) above balloon altitudes (˜40km) and below satellite orbits (˜160km). They can lift remote sensing telescope payloads with masses up to 400kg to altitudes of 350km providing observing times of up to 6min above the blocking influence of Earth’s atmosphere. Though a number of sounding rocket research programs exist around the world, this article focuses on the NASA Sounding Rocket Program, and particularly on the astrophysical and solar sounding rocket payloads.

  11. NASA's Advanced solid rocket motor

    Science.gov (United States)

    Mitchell, Royce E.

    The Advanced Solid Rocket Motor (ASRM) will not only bring increased safety, reliability and performance for the Space Shuttle Booster, it will enhance overall Shuttle safety by effectively eliminating 174 failure points in the Space Shuttle Main Engine throttling system and by reducing the exposure time to aborts due to main engine loss or shutdown. In some missions, the vulnerability time to Return-to-Launch Site aborts is halved. The ASRM uses case joints which will close or remain static under the effects of motor ignition and pressurization. The case itself is constructed of the weldable steel alloy HP 9-4-0.30, having very high strength and with superior fracture toughness and stress corrosion resistance. The internal insulation is strip-wound and is free of asbestos. The nozzle employs light weight ablative parts and is some 5,000 pounds lighter than the Shuttle motor used to date. The payload performance of the ASRM-powered Shuttle is 12,000 pounds higher than that provided by the present motor. This is of particular benefit for payloads delivered to higher inclinations and/or altitudes. The ASRM facility uses state-of-the-art manufacturing techniques, including continuous propellant mixing and direct casting.

  12. NASA's Advanced solid rocket motor

    Science.gov (United States)

    Mitchell, Royce E.

    1993-01-01

    The Advanced Solid Rocket Motor (ASRM) will not only bring increased safety, reliability and performance for the Space Shuttle Booster, it will enhance overall Shuttle safety by effectively eliminating 174 failure points in the Space Shuttle Main Engine throttling system and by reducing the exposure time to aborts due to main engine loss or shutdown. In some missions, the vulnerability time to Return-to-Launch Site aborts is halved. The ASRM uses case joints which will close or remain static under the effects of motor ignition and pressurization. The case itself is constructed of the weldable steel alloy HP 9-4-0.30, having very high strength and with superior fracture toughness and stress corrosion resistance. The internal insulation is strip-wound and is free of asbestos. The nozzle employs light weight ablative parts and is some 5,000 pounds lighter than the Shuttle motor used to date. The payload performance of the ASRM-powered Shuttle is 12,000 pounds higher than that provided by the present motor. This is of particular benefit for payloads delivered to higher inclinations and/or altitudes. The ASRM facility uses state-of-the-art manufacturing techniques, including continuous propellant mixing and direct casting.

  13. Microwave-assisted Pyrolysis of Swida wilsoniana Fruit Oil Soap for Preparing Renewable Hydrocarbon Fuel via Selective Decarboxylation%微波裂解光皮树油皂化物脱羧制备烃类燃料研究

    Institute of Scientific and Technical Information of China (English)

    刘玉环; 王允圃; 王应宽; 万益琴; 张锦胜; 阮榕生

    2012-01-01

    以光皮树果实油脂通过皂化反应获得的皂类为研究对象,利用微波裂解选择性加热优势开展皂类脱羧制备烃类燃料研究.通过气质联用等方法对裂解产物的分析表明,单纯钠皂微波裂解所得到液态产物一般都在皂类干质量的70%以上,裂解液态产物的密度为0.850~0.875 g/cm3,运动粘度为2.09 ~2.85 mm2/s,与柴油的性质基本相似.裂解液态产物中最高峰是十五碳烯,证实脱羧是皂类微波裂解的主要反应形式.%In order to prepare hydrocarbon fuel, sodium soap made from Swida wilsoniana fruit oil was chosen as a model compound with significant molecular polarity for preparing hydrocarbon fuel by microwave-assisted pyrolysis. The result showed that the hydrocarbon content were usually above 70% . In addition, its density and dynamic viscosity were 0. 850 ~ 0. 875 g/cm and 2. 09 ~ 2. 85 mm /s, respectively, similar with petroleum diesel. The highest peak in the GC - MS profile of liquid product was fifteen carbon alkenes, decarboxylation was proved to be main reaction in microwave-assisted pyrolysis.

  14. Turbopump Design and Analysis Approach for Nuclear Thermal Rockets

    Science.gov (United States)

    Chen, Shu-cheng 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 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 characteristics 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.

  15. Turbopump Design and Analysis Approach for Nuclear Thermal Rockets

    Science.gov (United States)

    Chen, Shu-Cheng 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 & 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.

  16. 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.

  17. 液体火箭发动机液膜冷却研究综述%Review of Research on Liquid Film Cooling for Liquid-propellant Rocket Engine

    Institute of Scientific and Technical Information of China (English)

    周红玲; 杨成虎; 刘犇

    2012-01-01

    液膜冷却对降低燃烧室和喷注器头部温度有显著作用,而且通道结构比较简单,因此在载人航天液体推进系统用姿轨控发动机中得到了广泛应用。液膜冷却的传热过程主要包括对流传热和沸腾传热两种形式,传质过程主要包括液膜的蒸发和中心主气流对液膜的携带。对液膜冷却过程的研究工作进行了综述,讨论了液膜冷却的异常升温现象和发生机理。%Liquid film cooling is widely used in attitude and ahitude liquid rocket engine for manned spaceflight propulsion sys- tem. With a comparatively simple channel structure, it is an effective method to protect the combustor and injector from high temperature. The two main heat transfer patterns in liquid film cooling are convection transfer and boiling heat transfer, while the process of mass transfer mainly includes liquid film evaporating and the carrying of the film by main flow. Research on the heat and mass transfer during liquid film cooling is reviewed, and the phenomenon and mecha- nism of abnormal temperature rise of liquid film cooling are discussed.

  18. Design Rules and Issues with Respect to Rocket Based Combined Cycles

    Science.gov (United States)

    2010-09-01

    Section Analysis As we seek for the accelerator, the inlet design is quite art of compromise. To make benefits due to air- breathing propulsion, the...Design Rules and Issues with Respect to Rocket Based Combined Cycles 3 - 4 RTO-EN-AVT-185 2.1.2 Combustor Section Analysis Embedded rocket chamber...cause thrust augmentation due to the ejector effects, which in turn, can reduce the requirement for the rocket engine output. In the speed regime with

  19. Taming Liquid Hydrogen: The Centaur Upper Stage Rocket

    Science.gov (United States)

    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.

  20. 基于动态云BP网络的液体火箭发动机故障诊断方法%Fault diagnosis method for liquid-propellant rocket engines based on the dynamic cloud-BP neural network

    Institute of Scientific and Technical Information of China (English)

    刘垠杰; 黄强; 程玉强; 吴建军

    2012-01-01

    将云模型与BP(backpropagation)神经网络以串联方式有机结合,首先利用云变换方法进行网络的结构辨识和云模型的特征提取,同时通过在输入层引入单位延时环节描述发动机工作过程动态特性,研究提出了基于动态云BP网络的液体火箭发动机故障诊断方法.结合实际试车数据的验证结果表明,该方法能够准确识别发动机已有的3种故障模式,通过在试车数据中添加0期望、0.2标准差的随机噪声的方法来模拟环境噪声和测试过程中产生的随机噪声,根据持续性原则,方法仍能够正确进行故障检测与分类.方法单步运行时长为1.124x10-4,完全能够满足实时性要求.%A fault diagnosis method for liquid-propellant rocket engines was proposed based on the dynamic cloud-BP(back propagation) neural network in the way of the integration of cloud model and BP neural network.The Cloud transform method was used to identify the network configuration and to extract the cloud features.And a unit time-delay was also introduced into the input layer to describe the dynamic characteristics of the engine.Results with test data show that the method can isolate the existed 3 fault modes precisely.A 0 expectation,0.2 standard deviation noise was used to simulate the entironmental noise and stochastic noise,and the method can still detect and classify the fault accurately acount to lasting-rule.The method can run in real-time with the single processing time being 1.124×10-4 s.

  1. Nuclear Thermal Rocket Propulsion Systems

    Science.gov (United States)

    2007-11-02

    NUCLEAR THERMAL ROCKET PROPULSION SYSTEMS, IAA WHITE PAPER PARIS, FRANCE, MARCH 2005 Lt Col Timothy J. Lawrence U.S. Air Force Academy...YYYY) 18-03-2005 2. REPORT TYPE White Paper 3. DATES COVERED (From - To) 18 Mar 2005 4. TITLE AND SUBTITLE NUCLEAR THERMAL ROCKET PROPULSION...reduce radiation exposure, is to have a high energy system like a nuclear thermal rocket that can get the payload to the destination in the fastest

  2. Influence of gunpowder start system on starting performance of liquid rocket engines%火药起动系统对发动机起动性能的影响分析

    Institute of Scientific and Technical Information of China (English)

    孙海雨; 刘志让

    2012-01-01

    Aiming at the pumping pressure open cycle liquid rocket engine which is started by solid start cartridge (SSC), the performance of the start system is studied in this paper. The simulated model of SSC in the start system and the numerical model of the powder gas pipeline were established to simulate the process of gunpowder start. The influence of SSC and gas pipeline parameters on start performance of engine is analyzed to ensure the main influence parameters and regular patterns. It is found that the powder quantity of SSC and the diameter of the first throat in the powder gas pipeline are the most effective factors to the engine's start characteristic. The diameter of the powder gas pipeline's second throat and the diameter of the powder gas pipeline's outlet are the least ones in the case of that the powder gas pipeline's flow field keeps rated condition. The simulated result of the start system was proven in engine hot tests.%针对采用火药起动器起动的泵压开式循环液体火箭发动机,对其起动系统进行了分析和研究。建立了液体火箭发动机火药起动器计算模型和起动系统燃气管路流场计算模型。将所建立的起动系统模型应用于发动机系统仿真,对发动机火药起动过程进行仿真,分析了起动系统中火药起动器参数和燃气管路参数对发动机起动性能的影响,确定了主要影响参数和影响规律。火药起动器火药药柱内径、火药药柱长度以及燃气管路火药起动器喷管喉部直径为强影响因素;燃气管路涡轮喷嘴喉部直径和管路出口直径在确保发动机火药起动主要工况段燃气管路流场流态为额定工况流态的前提下,为弱影响因素。试验数据验证表明,发动机起动系统的仿真结果正确、可信。

  3. 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;...

  4. Engineering Encounters: Blasting off with Engineering

    Science.gov (United States)

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

    2014-01-01

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

  5. Methane Dual Expander Aerospike Nozzle Rocket Engine

    Science.gov (United States)

    2012-03-22

    v Acknowledgments I would like to thank my thesis advisor, Lt. Col Hartsfield, who has put great effort in discussing, teaching , and...42 NPSS Thermochemistry ................................................................................................44 Methane FPT Generation...Next is a more focused description of how NPSS uses thermochemistry and an explanation on the development of the different fluid property tables

  6. Powder metallurgy bearings for advanced rocket engines

    Science.gov (United States)

    Fleck, J. N.; Killman, B. J.; Munson, H.E.

    1985-01-01

    Traditional ingot metallurgy was pushed to the limit for many demanding applications including antifriction bearings. New systems require corrosion resistance, better fatigue resistance, and higher toughness. With conventional processing, increasing the alloying level to achieve corrosion resistance results in a decrease in other properties such as toughness. Advanced powder metallurgy affords a viable solution to this problem. During powder manufacture, the individual particle solidifies very rapidly; as a consequence, the primary carbides are very small and uniformly distributed. When properly consolidated, this uniform structure is preserved while generating a fully dense product. Element tests including rolling contact fatigue, hot hardness, wear, fracture toughness, and corrosion resistance are underway on eleven candidate P/M bearing alloys and results are compared with those for wrought 440C steel, the current SSME bearing material. Several materials which offer the promise of a significant improvement in performance were identified.

  7. Standardization of Rocket Engine Pulse Time Parameters

    Science.gov (United States)

    Larin, Max E.; Lumpkin, Forrest E.; Rauer, Scott J.

    2001-01-01

    Plumes of bipropellant thrusters are a source of contamination. Small bipropellant thrusters are often used for spacecraft attitude control and orbit correction. Such thrusters typically operate in a pulse mode, at various pulse lengths. Quantifying their contamination effects onto spacecraft external surfaces is especially important for long-term complex-geometry vehicles, e.g. International Space Station. Plume contamination tests indicated the presence of liquid phase contaminant in the form of droplets. Their origin is attributed to incomplete combustion. Most of liquid-phase contaminant is generated during the startup and shutdown (unsteady) periods of thruster pulse. These periods are relatively short (typically 10-50 ms), and the amount of contaminant is determined by the thruster design (propellant valve response, combustion chamber size, thruster mass flow rate, film cooling percentage, dribble volume, etc.) and combustion process organization. Steady-state period of pulse is characterized by much lower contamination rates, but may be lengthy enough to significantly conh'ibute to the overall contamination effect. Because there was no standard methodology for thruster pulse time division, plume contamination tests were conducted at various pulse durations, and their results do not allow quantifying contaminant amounts from each portion of the pulse. At present, the ISS plume contamination model uses an assumption that all thrusters operate in a pulse mode with the pulse length being 100 ms. This assumption may lead to a large difference between the actual amounts of contaminant produced by the thruster and the model predictions. This paper suggests a way to standardize thruster startup and shutdown period definitions, and shows the usefulness of this approach to better quantify thruster plume contamination. Use of the suggested thruster pulse time-division technique will ensure methodological consistency of future thruster plume contamination test programs, and allow accounting for thruster pulse length when modeling plume contamination and erosion effects.

  8. Construction and Design of Rocket Engines,

    Science.gov (United States)

    1981-02-12

    l.ntring together with the turbine intc the so-called tirbopump aggrqgat,’ (TNA). For tho wnrk c- turbina , .riz d for l- ’ . .vi, "-t is necessary to have... turbina by qasqous vcrking m-liJum/proop! L[ant-, e) branch system cf the exhaust (crushed) gas. TNA is intqnded for increasing the pressure of the...Starting systems of turbine are examined ir §14.1. In the power-supply system of turbina by gaseous working medium/propellant are 3ncluded ZhGG, the

  9. Nuclear Thermal Rocket Element Environmental Simulator (NTREES)

    Science.gov (United States)

    Emrich, William J.

    2008-01-01

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

  10. Computational simulation of liquid rocket injector anomalies

    Science.gov (United States)

    Przekwas, A. J.; Singhal, A. K.; Tam, L. T.; Davidian, K.

    1986-01-01

    A computer model has been developed to analyze the three-dimensional two-phase reactive flows in liquid fueled rocket combustors. The model is designed to study the influence of liquid propellant injection nonuniformities on the flow pattern, combustion and heat transfer within the combustor. The Eulerian-Lagrangian approach for simulating polidisperse spray flow, evaporation and combustion has been used. Full coupling between the phases is accounted for. A nonorthogonal, body fitted coordinate system along with a conservative control volume formulation is employed. The physical models built into the model include a kappa-epsilon turbulence model, a two-step chemical reaction, and the six-flux radiation model. Semiempirical models are used to describe all interphase coupling terms as well as chemical reaction rates. The purpose of this study was to demonstrate an analytical capability to predict the effects of reactant injection nonuniformities (injection anomalies) on combustion and heat transfer within the rocket combustion chamber. The results show promising application of the model to comprehensive modeling of liquid propellant rocket engines.

  11. Nuclear Thermal Rocket Element Environmental Simulator (NTREES)

    Science.gov (United States)

    Emrich, William J., Jr.

    2008-01-01

    To support the eventual development of a nuclear thermal rocket engine, a state-of-the-art 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 will 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.

  12. NASA Space Rocket Logistics Challenges

    Science.gov (United States)

    Neeley, James R.; Jones, James V.; Watson, Michael D.; Bramon, Christopher J.; Inman, Sharon K.; Tuttle, Loraine

    2014-01-01

    The Space Launch System (SLS) is the new NASA heavy lift launch vehicle and is scheduled for its first mission in 2017. The goal of the first mission, which will be uncrewed, is to demonstrate the integrated system performance of the SLS rocket and spacecraft before a crewed flight in 2021. SLS has many of the same logistics challenges as any other large scale program. Common logistics concerns for SLS include integration of discreet programs geographically separated, multiple prime contractors with distinct and different goals, schedule pressures and funding constraints. However, SLS also faces unique challenges. The new program is a confluence of new hardware and heritage, with heritage hardware constituting seventy-five percent of the program. This unique approach to design makes logistics concerns such as commonality especially problematic. Additionally, a very low manifest rate of one flight every four years makes logistics comparatively expensive. That, along with the SLS architecture being developed using a block upgrade evolutionary approach, exacerbates long-range planning for supportability considerations. These common and unique logistics challenges must be clearly identified and tackled to allow SLS to have a successful program. This paper will address the common and unique challenges facing the SLS programs, along with the analysis and decisions the NASA Logistics engineers are making to mitigate the threats posed by each.

  13. Effects of turbulent and spray models on combustion process simultion of LOX/GH2 rocket engine%湍流、喷雾模型对氢氧火箭发动机燃烧仿真的影响

    Institute of Scientific and Technical Information of China (English)

    程钰锋; 聂万胜; 丰松江

    2011-01-01

    Based on the improved PISO algorithm, the numerical simulation for the combustion instability of a LOX/GH2 rocket engine was conducted by changing the turbulent and spay models of κ-ε equations. The compared results of the theoretical analysis and numerical simulation show that in the two-dimensional situation, both droplet collision model and TAB droplet breakup model are not suitable for the numerical simulation of LOX/GH2 combustion instability; the pressure oscillation in the combustion chamber can be simulated by combining the TVB droplet breakup model with the turbulent models of κ-ε equations, but the oscillation frequency can not be simulated; if the turbulent models of Realizable κ-ε equations are adopted without consideration of the droplet spray models, both the pressure oscillation in the combustion chamber and the distribution of the oscillation frequency can be simulated.%基于完善的压力隐式算子分裂(PISO)算法,通过改变κ-ε两方程湍流模型和喷雾模型,对氢氧火箭发动机不稳定燃烧进行数值仿真。比较理论分析和数值仿真的结果得出,在二维情况下,液滴碰撞模型和TAB液滴破碎模型不适于模拟氢氧火箭发动机不稳定燃烧;TVB液滴破碎模型与κ-ε两方程湍流模型的组合情况能够捕捉到燃烧室中的压力振荡,但不能体现出振荡频率;而采用Realizableκ-ε湍流模型不考虑液滴雾化模型时不但能够捕捉燃烧室内压力振荡情况,还能够很好地得出振荡频率的分布情况。

  14. Numerical investigations of thermal stratification in cooling channel of liquid rocket engine thrust chamber%液体火箭发动机推力室冷却通道温度分层数值研究

    Institute of Scientific and Technical Information of China (English)

    康玉东; 孙冰; 高翔宇

    2009-01-01

    为了研究冷却剂温度分层的形成机理及其对流动和换热的影响,应用雷诺应力模型(RSM)对液体火箭发动机推力室再生冷却通道的流动与传热进行了三维数值模拟,冷却剂为气氢,考虑其物性随温度和压力的变化.所得结果表明:冷却剂在非流动方向会出现温度分层现象,随着冷却剂的不断受热,温度分层现象越明显,由于喉部二次流加强了冷却剂间的混合,在喉部区域温度分层被减弱,温度分层对冷却剂温升及压降影响较小,严重影响气壁温度的估算.%To study the formation mechanism of thermal stratification in cooling channel and its effects on the flow and heat transfer,three dimensional turbulent fluid flow and heat transfer in a regenerative-cooling channel of liquid rocket engine were numerically investigated with Reynolds stress model(RSM) model,and the coolant was hydrogen,whose thermo-physical properties varied with both temperature and pressure.The results show that thermal stratification occurs at non-flow direction,the extent of thermal stratification becomes increasingly significant as the extent of heating increases,and the thermal stratification of coolant is weakened for the existence of secondary flow in throat region; the thermal stratification has little effect on the bulk temperature increase and hydrodynamic losses of hydrogen,but has significant effect on the calculation of the wall heat fluxes and temperature.

  15. What fuel for a rocket?

    CERN Document Server

    Miranda, E N

    2012-01-01

    Elementary concepts from general physics and thermodynamics have been used to analyze rocket propulsion. Making some reasonable assumptions, an expression for the exit velocity of the gases is found. From that expression one can conclude what are the desired properties for a rocket fuel.

  16. Rocket launchers as passive controllers

    Science.gov (United States)

    Cochran, J. E., Jr.; Gunnels, R. T.; McCutchen, R. K., Jr.

    1981-12-01

    A concept is advanced for using the motion of launchers of a free-flight launcher/rocket system which is caused by random imperfections of the rockets launched from it to reduce the total error caused by the imperfections. This concept is called 'passive launcher control' because no feedback is generated by an active energy source after an error is sensed; only the feedback inherent in the launcher/rocket interaction is used. Relatively simple launcher models with two degrees of freedom, pitch and yaw, were used in conjunction with a more detailed, variable-mass model in a digital simulation code to obtain rocket trajectories with and without thrust misalignment and dynamic imbalance. Angular deviations of rocket velocities and linear deviations of the positions of rocket centers of mass at burnout were computed for cases in which the launcher was allowed to move ('flexible' launcher) and was constrained so that it did not rotate ('rigid' launcher) and ratios of flexible to rigid deviations were determined. Curves of these error ratios versus launcher frequency are presented. These show that a launcher which has a transverse moment of inertia about its pivot point of the same magnitude as that of the centroidal transverse moments of inertia of the rockets launched from it can be tuned to passively reduce the errors caused by rocket imperfections.

  17. The Effect of Resistance on Rocket Injector Acoustics

    Science.gov (United States)

    Morgan, C. J.

    2015-01-01

    Combustion instability, where unsteady heat release couples with acoustic modes, has long been an area of concern in liquid rocket engines. Accurate modeling of the acoustic normal modes of the combustion chamber is important to understanding and preventing combustion instability. The injector resistance can have a significant influence on the chamber normal mode shape, and hence on the system stability.

  18. STS-27 Atlantis, OV-104, solid rocket booster (SRB) inspection

    Science.gov (United States)

    1988-01-01

    Engineers, kneeling inside a hollow solid rocket booster (SRB), closely inspect the SRB segments and seams in the Kennedy Space Center (KSC) rotation and processing facility. The SRB will be used on STS-27 Atlantis, Orbiter Vehicle (OV) 104. The booster segments were transported via rail car from Morton Thiokol's Utah manufacturing plant. View provided by KSC with alternate number KSC-88PC-492.

  19. Review of Systems for Photocatalytic Conversion of CO2 to Hydrocarbon Fuels%光催化CO2转化为碳氢燃料体系的综述

    Institute of Scientific and Technical Information of China (English)

    蓝奔月; 史海峰

    2014-01-01

    Increasing global warming and energy shortage caused by traditional fossil energy combustion to carbon dioxide (CO2) has become a significant global issue in view of humansʹcontinuing development. The photocatalytic reduction of CO2 produced from hydrocarbon fuels using solar light and semiconductor photocatalytic materials could not only decrease the concentration of carbon dioxide in the atmosphere and thus reduce the greenhouse warming effect, but also provide hydrocarbon fuels to partial y al eviate the energy shortage crisis. Hence, the photocatalysis technique has attracted considerable attention in industry and academic areas. In this paper, the fundamental principles of heterogeneous photocatalysis and the recent progress in the photocatalytic reduction of CO2 to hydrocarbon fuels are introduced and reviewed. Based on previous reports in the field of photocatalysis research, the main types of semiconductors capable of photocatalytic reduction of carbon dioxide can be summarized as fol ows: pure TiO2 photocatalysts, ABO3 perovskite-structured photocatalysts, spinel-structured photocatalysts, doped oxide photocatalysts, composite semiconductor photocatalysts, V-, W-, Ge-, Ga-based photocatalysts, and graphene-based photocatalysts. In addition, the characteristics of various photocatalytic materials and some factors affecting photocatalytic activities are reviewed and analyzed. Final y, the prospects and chal enges for developing new photocatalysts for CO2 reduction are presented.%传统化石能源燃烧产生CO2引起的地球变暖和能源短缺已经成为一个严重的全球性问题。利用太阳光和光催化材料将CO2还原为碳氢燃料,不仅可以减少空气中CO2浓度,降低温室效应的影响,还可以提供碳氢燃料,缓解能源短缺问题,因此日益受到各国科学家的高度关注。本文综述了光催化还原CO2为碳氢燃料的研究进展,介绍了光催化还原CO2的反应机理,并对

  20. Process-Hardened, Multi-Analyte Sensor for Characterizing Rocket Plum Constituents Under Test Environment Project

    Data.gov (United States)

    National Aeronautics and Space Administration — This STTR project aims to develop a process-hardened, simple and low cost multi-analyte sensor for detecting components of rocket engine plumes. The sensor will be...

  1. Laser-Induced Emissions Sensor for Soot Mass in Rocket Plumes Project

    Data.gov (United States)

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

  2. Airbreathing combined cycle engine systems

    Science.gov (United States)

    Rohde, John

    1992-01-01

    The Air Force and NASA share a common interest in developing advanced propulsion systems for commercial and military aerospace vehicles which require efficient acceleration and cruise operation in the Mach 4 to 6 flight regime. The principle engine of interest is the turboramjet; however, other combined cycles such as the turboscramjet, air turborocket, supercharged ejector ramjet, ejector ramjet, and air liquefaction based propulsion are also of interest. Over the past months careful planning and program implementation have resulted in a number of development efforts that will lead to a broad technology base for those combined cycle propulsion systems. Individual development programs are underway in thermal management, controls materials, endothermic hydrocarbon fuels, air intake systems, nozzle exhaust systems, gas turbines and ramjet ramburners.

  3. Concurrent engineering

    Science.gov (United States)

    Chamis, C. C.; Leger, L.; Hunter, D.; Jones, C.; Sprague, R.; Berke, L.; Newell, J.; Singhal, S.

    1991-01-01

    The following subject areas are covered: issues (liquid rocket propulsion - current development approach, current certification process, and costs of engineering changes); state of the art (DICE information management system, key government participants, project development strategy, quality management, and numerical propulsion system simulation); needs identified; and proposed program.

  4. Iridium-Coated Rhenium Radiation-Cooled Rockets

    Science.gov (United States)

    Reed, Brian D.; Biaglow, James A.; Schneider, Steven J.

    1997-01-01

    Radiation-cooled rockets are used for a range of low-thrust propulsion functions, including apogee insertion, attitude control, and repositioning of satellites, reaction control of launch vehicles, and primary propulsion for planetary space- craft. The key to high performance and long lifetimes for radiation-cooled rockets is the chamber temperature capability. The material system that is currently used for radiation-cooled rockets, a niobium alloy (C103) with a fused silica coating, has a maximum operating temperature of 1370 C. Temperature limitations of C103 rockets force the use of fuel film cooling, which degrades rocket performance and, in some cases, imposes a plume contamination issue from unburned fuel. A material system composed of a rhenium (Re) substrate and an iridium (Ir) coating has demonstrated operation at high temperatures (2200 C) and for long lifetimes (hours). The added thermal margin afforded by iridium-coated rhenium (Ir/Re) allows reduction or elimination of fuel film cooling. This, in turn, leads to higher performance and cleaner spacecraft environments. There are ongoing government- and industry-sponsored efforts to develop flight Ir/ Re engines, with the primary focus on 440-N, apogee insertion engines. Complementing these Ir/Re engine development efforts is a program to address specific concerns and fundamental characterization of the Ir/Re material system, including (1) development of Ir/Re rocket fabrication methods, (2) establishment of critical Re mechanical properly data, (3) development of reliable joining methods, and (4) characterization of Ir/Re life-limiting mechanisms.

  5. Application of Dry Coupling Ultrasonic Detecting Technology for the Rocket Engine Nozzle%干耦合超声检测技术在某火箭发动机喷管在役检测中的应用

    Institute of Scientific and Technical Information of China (English)

    穆洪彬; 吴朝军; 吴晨; 李剑

    2013-01-01

    应用干耦合超声检测方法对某火箭发动机喷管进行了在役检测试验,验证了干耦合超声检测技术对该火箭发动机喷管的在役检测能力.结果表明,干耦合超声检测方法能够有效地检测出该型号发动机喷管中φ 20 mm以上的脱粘缺陷,满足在役检测要求.自动与手动,两种检测方法的检测能力相当,但在操作方式和工作效率等方面各有优劣.%In order to study the feasibility of the in-service dry coupling ultrasonic test technology for rocket jet,the principle of the dry coupling ultrasonic testing was analyzed,and both the manual and automatic testing methods were applied to test the rocket jet specimen and the rocket jet parts in service.The results show that the debond defect,whose normal size is larger than 20mm,can be found using the two kinds of the dry coupling ultrasonic testing methods.It satisfies the quality request of the rocket jet being tested in service.It is the same to the manual testing and automatic testing in the ability of finding defect.And there is difference between the two testing methods in the operation and efficiency.The study is useful for testing rocket jet in service.

  6. Study of solid rocket motor for space shuttle booster, Volume 3: Program acquisition planning

    Science.gov (United States)

    1972-01-01

    The program planning acquisition functions for the development of the solid propellant rocket engine for the space shuttle booster is presented. The subjects discussed are: (1) program management, (2) contracts administration, (3) systems engineering, (4) configuration management, and (5) maintenance engineering. The plans for manufacturing, testing, and operations support are included.

  7. The Application of Counter-Rotating Turbine in Rocket Turbopump

    Directory of Open Access Journals (Sweden)

    Tang Fei

    2008-01-01

    Full Text Available Counter rotating turbine offers advantages on weight, volume, efficiency, and maneuverability relative to the conventional turbine because of its special architecture. Nowadays, it has been a worldwide research emphasis and has been used widely in the aeronautic field, while its application in the astronautic field is seldom investigated. Researches of counter rotating turbine for rocket turbopump are reviewed in this paper. A primary analysis of a vaneless counter rotating-turbine configuration with rotors of different diameters and rotational speeds is presented. This unconventional configuration meets the requirements of turbopump and may benefit the performance and reliability of rocket engines.

  8. Analytical study of nozzle performance for nuclear thermal rockets

    Science.gov (United States)

    Davidian, Kenneth O.; Kacynski, Kenneth J.

    1991-01-01

    A parametric study has been conducted by the NASA-Lewis Rocket Engine Design Expert System for the convergent-divergent nozzle of the Nuclear Thermal Rocket system, which uses a nuclear reactor to heat hydrogen to high temperature and then expands it through the nozzle. It is established by the study that finite-rate chemical reactions lower performance levels from theoretical levels. Major parametric roles are played by chamber temperature and chamber pressure. A maximum performance of 930 sec is projected at 2700 K, and of 1030 at 3100 K.

  9. Advanced materials for radiation-cooled rockets

    Science.gov (United States)

    Reed, Brian; Biaglow, James; Schneider, Steven

    1993-01-01

    The most common material system currently used for low thrust, radiation-cooled rockets is a niobium alloy (C-103) with a fused silica coating (R-512A or R-512E) for oxidation protection. However, significant amounts of fuel film cooling are usually required to keep the material below its maximum operating temperature of 1370 C, degrading engine performance. Also the R-512 coating is subject to cracking and eventual spalling after repeated thermal cycling. A new class of high-temperature, oxidation-resistant materials are being developed for radiation-cooled rockets, with the thermal margin to reduce or eliminate fuel film cooling, while still exceeding the life of silicide-coated niobium. Rhenium coated with iridium is the most developed of these high-temperature materials. Efforts are on-going to develop 22 N, 62 N, and 440 N engines composed of these materials for apogee insertion, attitude control, and other functions. There is also a complimentary NASA and industry effort to determine the life limiting mechanisms and characterize the thermomechanical properties of these materials. Other material systems are also being studied which may offer more thermal margin and/or oxidation resistance, such as hafnium carbide/tantalum carbide matrix composites and ceramic oxide-coated iridium/rhenium chambers.

  10. Lunar mission design using Nuclear Thermal Rockets

    Science.gov (United States)

    Stancati, Michael L.; Collins, John T.; Borowski, Stanley 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.

  11. British used Congreve Rockets to Attack Napoleon

    Science.gov (United States)

    2004-01-01

    Sir William Congreve developed a rocket with a range of about 9,000 feet. The incendiary rocket used black powder, an iron case, and a 16-foot guide stick. In 1806, British used Congreve rockets to attack Napoleon's headquarters in France. In 1807, Congreve directed a rocket attack against Copenhagen.

  12. Rockets and People. Volume 1

    Science.gov (United States)

    Chertok, Boris E; Siddiqi, Asif A. (Editor)

    2005-01-01

    Much has been written in the West on the history of the Soviet space program but few Westerners have read direct first-hand accounts of the men and women who were behind the many Russian accomplishments in exploring space.The memoirs of Academician Boris Chertok, translated from the original Russian, fills that gap.Chertok began his career as an electrician in 1930 at an aviation factory near Moscow.Twenty-seven years later, he became deputy to the founding figure of the Soviet space program, the mysterious Chief Designer Sergey Korolev. Chertok s sixty-year-long career and the many successes and failures of the Soviet space program constitute the core of his memoirs, Rockets and People. These writings are spread over four volumes. This is volume I. Academician Chertok not only describes and remembers, but also elicits and extracts profound insights from an epic story about a society s quest to explore the cosmos. In Volume 1, Chertok describes his early years as an engineer and ends with the mission to Germany after the end of World War II when the Soviets captured Nazi missile technology and expertise. Volume 2 takes up the story with the development of the world s first intercontinental ballistic missile ICBM) and ends with the launch of Sputnik and the early Moon probes. In Volume 3, Chertok recollects the great successes of the Soviet space program in the 1960s including the launch of the world s first space voyager Yuriy Gagarin as well as many events connected with the Cold War. Finally, in Volume 4, Chertok meditates at length on the massive Soviet lunar project designed to beat the Americans to the Moon in the 1960s, ending with his remembrances of the Energiya-Buran project.

  13. Alternate Propellant Thermal Rocket Project

    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...

  14. Magnesium and Carbon Dioxide - A Rocket Propellant for Mars Missions

    Science.gov (United States)

    Shafirovich, E. IA.; Shiriaev, A. A.; Goldshleger, U. I.

    1993-01-01

    A rocket engine for Mars missions is proposed that could utilize CO2 accumulated from the Martian atmosphere as an oxidizer. For use as possible fuel, various metals, their hydrides, and mixtures with hydrogen compounds are considered. Thermodynamic calculations show that beryllium fuels ensure the most impulse but poor inflammability of Be and high toxicity of its compounds put obstacles to their applications. Analysis of the engine performance for other metals together with the parameters of ignition and combustion show that magnesium seems to be the most promising fuel. Ballistic estimates imply that a hopper with the chemical rocket engine on Mg + CO2 propellant could be readily developed. This vehicle would be able to carry out 2-3 ballistic flights on Mars before the final ascent to orbit.

  15. '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

  16. Improved hybrid rocket fuel

    Science.gov (United States)

    Dean, David L.

    1995-01-01

    McDonnell Douglas Aerospace, as part of its Independent R&D, has initiated development of a clean burning, high performance hybrid fuel for consideration as an alternative to the solid rocket thrust augmentation currently utilized by American space launch systems including Atlas, Delta, Pegasus, Space Shuttle, and Titan. It could also be used in single stage to orbit or as the only propulsion system in a new launch vehicle. Compared to solid propellants based on aluminum and ammonium perchlorate, this fuel is more environmentally benign in that it totally eliminates hydrogen chloride and aluminum oxide by products, producing only water, hydrogen, nitrogen, carbon oxides, and trace amounts of nitrogen oxides. Compared to other hybrid fuel formulations under development, this fuel is cheaper, denser, and faster burning. The specific impulse of this fuel is comparable to other hybrid fuels and is between that of solids and liquids. The fuel also requires less oxygen than similar hybrid fuels to produce maximum specific impulse, thus reducing oxygen delivery system requirements.

  17. Rehabilitation of the Rocket Vehicle Integration Test Stand at Edwards Air Force Base

    Science.gov (United States)

    Jones, Daniel S.; Ray, Ronald J.; Phillips, Paul

    2005-01-01

    Since initial use in 1958 for the X-15 rocket-powered research airplane, the Rocket Engine Test Facility has proven essential for testing and servicing rocket-powered vehicles at Edwards Air Force Base. For almost two decades, several successful flight-test programs utilized the capability of this facility. The Department of Defense has recently demonstrated a renewed interest in propulsion technology development with the establishment of the National Aerospace Initiative. More recently, the National Aeronautics and Space Administration is undergoing a transformation to realign the organization, focusing on the Vision for Space Exploration. These initiatives provide a clear indication that a very capable ground-test stand at Edwards Air Force Base will be beneficial to support the testing of future access-to-space vehicles. To meet the demand of full integration testing of rocket-powered vehicles, the NASA Dryden Flight Research Center, the Air Force Flight Test Center, and the Air Force Research Laboratory have combined their resources in an effort to restore and upgrade the original X-15 Rocket Engine Test Facility to become the new Rocket Vehicle Integration Test Stand. This report describes the history of the X-15 Rocket Engine Test Facility, discusses the current status of the facility, and summarizes recent efforts to rehabilitate the facility to support potential access-to-space flight-test programs. A summary of the capabilities of the facility is presented and other important issues are discussed.

  18. Nuclear rockets: High-performance propulsion for Mars

    Science.gov (United States)

    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.

  19. Thrust Vector Control for Nuclear Thermal Rockets

    Science.gov (United States)

    Ensworth, Clinton B. F.

    2013-01-01

    Future space missions may use Nuclear Thermal Rocket (NTR) stages for human and cargo missions to Mars and other destinations. The vehicles are likely to require engine thrust vector control (TVC) to maintain desired flight trajectories. This paper explores requirements and concepts for TVC systems for representative NTR missions. Requirements for TVC systems were derived using 6 degree-of-freedom models of NTR vehicles. Various flight scenarios were evaluated to determine vehicle attitude control needs and to determine the applicability of TVC. Outputs from the models yielded key characteristics including engine gimbal angles, gimbal rates and gimbal actuator power. Additional factors such as engine thrust variability and engine thrust alignment errors were examined for impacts to gimbal requirements. Various technologies are surveyed for TVC systems for the NTR applications. A key factor in technology selection is the unique radiation environment present in NTR stages. Other considerations including mission duration and thermal environments influence the selection of optimal TVC technologies. Candidate technologies are compared to see which technologies, or combinations of technologies best fit the requirements for selected NTR missions. Representative TVC systems are proposed and key properties such as mass and power requirements are defined. The outputs from this effort can be used to refine NTR system sizing models, providing higher fidelity definition for TVC systems for future studies.

  20. The Chameleon Solid Rocket Propulsion Model

    Science.gov (United States)

    Robertson, Glen A.

    2010-01-01

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