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Sample records for rocket engine performance

  1. Radiation effect on rocket engine performance

    NASA Technical Reports Server (NTRS)

    Chiu, Huei-Huang

    1988-01-01

    The effects of radiation on the performance of modern rocket propulsion systems operating at high pressure and temperature were recognized as a key issue in the design and operation of various liquid rocket engines of the current and future generations. Critical problem areas of radiation coupled with combustion of bipropellants are assessed and accounted for in the formulation of a universal scaling law incorporated with a radiation-enhanced vaporization combustion model. Numerical algorithms are developed and the pertaining data of the Variable Thrust Engine (VTE) and Space Shuttle Main Engine (SSME) are used to conduct parametric sensitivity studies to predict the principal intercoupling effects of radiation. The analysis reveals that low enthalpy engines, such as the VTE, are vulnerable to a substantial performance set back by the radiative loss, whereas the performance of high enthalpy engines such as the SSME, are hardly affected over a broad range of engine operation. Additionally, combustion enhancement by the radiative heating of the propellant has a significant impact in those propellants with high absorptivity. Finally, the areas of research related with radiation phenomena in bipropellant engines are identified.

  2. Numerical Modeling of Pulse Detonation Rocket Engine Gasdynamics and Performance

    NASA Technical Reports Server (NTRS)

    2003-01-01

    This paper presents viewgraphs on the numerical modeling of pulse detonation rocket engines (PDRE), with an emphasis on the Gasdynamics and performance analysis of these engines. The topics include: 1) Performance Analysis of PDREs; 2) Simplified PDRE Cycle; 3) Comparison of PDRE and Steady-State Rocket Engines (SSRE) Performance; 4) Numerical Modeling of Quasi 1-D Rocket Flows; 5) Specific PDRE Geometries Studied; 6) Time-Accurate Thrust Calculations; 7) PDRE Performance (Geometries A B C and D); 8) PDRE Blowdown Gasdynamics (Geom. A B C and D); 9) PDRE Geometry Performance Comparison; 10) PDRE Blowdown Time (Geom. A B C and D); 11) Specific SSRE Geometry Studied; 12) Effect of F-R Chemistry on SSRE Performance; 13) PDRE/SSRE Performance Comparison; 14) PDRE Performance Study; 15) Grid Resolution Study; and 16) Effect of F-R Chemistry on SSRE Exit Species Mole Fractions.

  3. Performance of a RBCC Engine in Rocket-Operation

    NASA Astrophysics Data System (ADS)

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

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

  4. Numerical Modeling of Pulse Detonation Rocket Engine Gasdynamics and Performance

    NASA Technical Reports Server (NTRS)

    Morris, C. I.

    2003-01-01

    Pulse detonation engines (PDB) have generated considerable research interest in recent years as a chemical propulsion system potentially offering improved performance and reduced complexity compared to conventional gas turbines and rocket engines. The detonative mode of combustion employed by these devices offers a theoretical thermodynamic advantage over the constant-pressure deflagrative combustion mode used in conventional engines. However, the unsteady blowdown process intrinsic to all pulse detonation devices has made realistic estimates of the actual propulsive performance of PDES 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.

  5. Scaling of Performance in Liquid Propellant Rocket Engine Combustors

    NASA Technical Reports Server (NTRS)

    Hulka, James R.

    2007-01-01

    This paper discusses scaling of combustion and combustion performance in liquid propellant rocket engine combustion devices. In development of new combustors, comparisons are often made between predicted performance in a new combustor and measured performance in another combustor with different geometric and thermodynamic characteristics. Without careful interpretation of some key features, the comparison can be misinterpreted and erroneous information used in the design of the new device. This paper provides a review of this performance comparison, including a brief review of the initial liquid rocket scaling research conducted during the 1950s and 1960s, a review of the typical performance losses encountered and how they scale, a description of the typical scaling procedures used in development programs today, and finally a review of several historical development programs to see what insight they can bring to the questions at hand.

  6. Scaling of Performance in Liquid Propellant Rocket Engine Combustion Devices

    NASA Technical Reports Server (NTRS)

    Hulka, James R.

    2008-01-01

    This paper discusses scaling of combustion and combustion performance in liquid propellant rocket engine combustion devices. In development of new combustors, comparisons are often made between predicted performance in a new combustor and measured performance in another combustor with different geometric and thermodynamic characteristics. Without careful interpretation of some key features, the comparison can be misinterpreted and erroneous information used in the design of the new device. This paper provides a review of this performance comparison, including a brief review of the initial liquid rocket scaling research conducted during the 1950s and 1960s, a review of the typical performance losses encountered and how they scale, a description of the typical scaling procedures used in development programs today, and finally a review of several historical development programs to see what insight they can bring to the questions at hand.

  7. Temperature measurement. [liquid monopropellant rocket engine performance tests

    NASA Technical Reports Server (NTRS)

    1979-01-01

    The design, installation, checkout, calibration, and operation of a temperature measuring system to be used during tests of a liquid monopropellant rocket engine are discussed. Appendixes include: (1) temperature measurement system elemental uncertainties, and (2) tables and equations for use with thermocouples and resistance thermometers. Design guidelines are given for the critical components of each portion of the system to provide an optimum temperature measurement system which meets the performance criteria specified.

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

    NASA Technical Reports Server (NTRS)

    Westra, Douglas G.; West, Jeffrey S.

    2014-01-01

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

  9. Large liquid rocket engine transient performance simulation system

    NASA Technical Reports Server (NTRS)

    Mason, J. R.; Southwick, R. D.

    1991-01-01

    A simulation system, ROCETS, was designed and developed to allow cost-effective computer predictions of liquid rocket engine transient performance. The system allows a user to generate a simulation of any rocket engine configuration using component modules stored in a library through high-level input commands. The system library currently contains 24 component modules, 57 sub-modules and maps, and 33 system routines and utilities. FORTRAN models from other sources can be operated in the system upon inclusion of interface information on comment cards. Operation of the simulation is simplified for the user by run, execution, and output processors. The simulation system makes available steady-state trim balance, transient operation, and linear partial generation. The system utilizes a modern equation solver for efficient operation of the simulations. Transient integration methods include integral and differential forms for the trapezoidal, first order Gear, and second order Gear corrector equations. A detailed technology test bed engine (TTBE) model was generated to be used as the acceptance test of the simulation system. The general level of model detail was that reflected in the Space Shuttle Main Engine DTM. The model successfully obtained steady-state balance in main stage operation and simulated throttle transients, including engine starts and shutdown. A NASA FORTRAN control model was obtained, ROCETS interface installed in comment cards, and operated with the TTBE model in closed-loop transient mode.

  10. Solid rocket booster performance evaluation model. Volume 1: Engineering description

    NASA Technical Reports Server (NTRS)

    1974-01-01

    The space shuttle solid rocket booster performance evaluation model (SRB-II) is made up of analytical and functional simulation techniques linked together so that a single pass through the model will predict the performance of the propulsion elements of a space shuttle solid rocket booster. The available options allow the user to predict static test performance, predict nominal and off nominal flight performance, and reconstruct actual flight and static test performance. Options selected by the user are dependent on the data available. These can include data derived from theoretical analysis, small scale motor test data, large motor test data and motor configuration data. The user has several options for output format that include print, cards, tape and plots. Output includes all major performance parameters (Isp, thrust, flowrate, mass accounting and operating pressures) as a function of time as well as calculated single point performance data. The engineering description of SRB-II discusses the engineering and programming fundamentals used, the function of each module, and the limitations of each module.

  11. Pressure measurement. [liquid monopropellant rocket engine performance tests

    NASA Technical Reports Server (NTRS)

    1979-01-01

    Practices are outlined for the design, installation, checkout, calibration, and operation of a pressure measuring system to be used during tests of a liquid monopropellant rocket engine. Appendixes include: (1) pressure measurement system elemental uncertainties; (2) short- and long-term pressure measurement system uncertainty; (3) shunt calibration of pressure transducers; (4) special considerations for vacuum measurement; and (5) methods of determining the dynamic characteristics of pressure transducers. Design guidelines are provided for the critical components of each portion of the system to provide a pressure measurement system which meets the performance criteria specified.

  12. Russian Rocket Engine Test

    NASA Technical Reports Server (NTRS)

    1998-01-01

    NASA engineers successfully tested a Russian-built rocket engine on November 4, 1998 at the Marshall Space Flight Center (MSFC) Advanced Engine Test Facility, which had been used for testing the Saturn V F-1 engines and Space Shuttle Main engines. The MSFC was under a Space Act Agreement with Lockheed Martin Astronautics of Denver to provide a series of test firings of the Atlas III propulsion system configured with the Russian-designed RD-180 engine. The tests were designed to measure the performance of the Atlas III propulsion system, which included avionics and propellant tanks and lines, and how these components interacted with the RD-180 engine. The RD-180 is powered by kerosene and liquid oxygen, the same fuel mix used in Saturn rockets. The RD-180, the most powerful rocket engine tested at the MSFC since Saturn rocket tests in the 1960s, generated 860,000 pounds of thrust.

  13. Numerical Modeling of Pulse Detonation Rocket Engine Gasdynamics And Performance

    NASA Technical Reports Server (NTRS)

    Morris, Christopher I.

    2004-01-01

    Pulse detonation rocket engines (PDREs) offer potential performance improvements over conventional designs, but represent a challenging modeling task. A quasi-1-D, finite-rate chemistry computational fluid dynamics model for PDREs is described and implemented. Four different PDRE geometries are evaluated in this work: a baseline detonation tube, a detonation tube with a straight extension, and a detonation tube with two types of converging-diverging (C-D) nozzles. The effect of extension length and C-D nozzle area ratio on the single-shot gasdynamics and performance of a PDRE is studied over a wide range of blowdown pressure ratios (1-1000). The results indicate that a C-D nozzle is generally more effective than a straight extension in improving PDRE performance, particularly at higher pressure ratios. Additionally, the results show that the blowdown process of the C-D nozzle systems could be beneficially cut off well before the pressure at the end-wall reaches the ambient value. The performance results are also compared to a steady-state rocket system using similar modeling assumptions.

  14. Critical Performance of Turbopump Mechanical Elements for Rocket Engine

    NASA Astrophysics Data System (ADS)

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

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

  15. Scaling of Performance in Liquid Propellant Rocket Engine Combustors

    NASA Technical Reports Server (NTRS)

    Hulka, James

    2008-01-01

    The objectives are: a) Re-introduce to you the concept of scaling; b) Describe the scaling research conducted in the 1950s and early 1960s, and present some of their conclusions; c) Narrow the focus to scaling for performance of combustion devices for liquid propellant rocket engines; and d) Present some results of subscale to full-scale performance from historical programs. Scaling is "The ability to develop new combustion devices with predictable performance on the basis of test experience with old devices." Scaling can be used to develop combustion devices of any thrust size from any thrust size. Scaling is applied mostly to increase thrust. Objective is to use scaling as a development tool. - Move injector design from an "art" to a "science"

  16. Large liquid rocket engine transient performance simulation system

    NASA Technical Reports Server (NTRS)

    Mason, J. R.; Southwick, R. D.

    1989-01-01

    Phase 1 of the Rocket Engine Transient Simulation (ROCETS) program consists of seven technical tasks: architecture; system requirements; component and submodel requirements; submodel implementation; component implementation; submodel testing and verification; and subsystem testing and verification. These tasks were completed. Phase 2 of ROCETS consists of two technical tasks: Technology Test Bed Engine (TTBE) model data generation; and system testing verification. During this period specific coding of the system processors was begun and the engineering representations of Phase 1 were expanded to produce a simple model of the TTBE. As the code was completed, some minor modifications to the system architecture centering on the global variable common, GLOBVAR, were necessary to increase processor efficiency. The engineering modules completed during Phase 2 are listed: INJTOO - main injector; MCHBOO - main chamber; NOZLOO - nozzle thrust calculations; PBRNOO - preburner; PIPE02 - compressible flow without inertia; PUMPOO - polytropic pump; ROTROO - rotor torque balance/speed derivative; and TURBOO - turbine. Detailed documentation of these modules is in the Appendix. In addition to the engineering modules, several submodules were also completed. These submodules include combustion properties, component performance characteristics (maps), and specific utilities. Specific coding was begun on the system configuration processor. All functions necessary for multiple module operation were completed but the SOLVER implementation is still under development. This system, the Verification Checkout Facility (VCF) allows interactive comparison of module results to store data as well as provides an intermediate checkout of the processor code. After validation using the VCF, the engineering modules and submodules were used to build a simple TTBE.

  17. Performance Increase Verification for a Bipropellant Rocket Engine

    NASA Technical Reports Server (NTRS)

    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.

  18. Measurements for liquid rocket engine performance code verification

    NASA Technical Reports Server (NTRS)

    Praharaj, Sarat C.; Palko, Richard L.

    1986-01-01

    The goal of the rocket engine performance code verification tests is to obtain the I sub sp with an accuracy of 0.25% or less. This needs to be done during the sequence of four related tests (two reactive and two hot gas simulation) to best utilize the loss separation technique recommended in this study. In addition to I sub sp, the measurements of the input and output parameters for the codes are needed. This study has shown two things in regard to obtaining the I sub sp uncertainty within the 0.25% target. First, this target is generally not being realized at the present time, and second, the instrumentation and testing technology does exist to obtain this 0.25% uncertainty goal. However, to achieve this goal will require carefully planned, designed, and conducted testing. In addition, the test-stand (or system) dynamics must be evaluated in the pre-test and post-test phases of the design of the experiment and data analysis, respectively always keeping in mind that a .25% overall uncertainty in I sub sp is targeted. A table gives the maximum allowable uncertainty required for obtaining I sub sp with 0.25% uncertainty, the currently-quoted instrument specification, and present test uncertainty for the parameters. In general, it appears that measurement of the mass flow parameter within the required uncertainty may be the most difficult.

  19. Russian Rocket Engine Test

    NASA Technical Reports Server (NTRS)

    1998-01-01

    NASA engineers successfully tested a Russian-built rocket engine on November 4, 1998 at the Marshall Space Flight Center (MSFC) Advanced Engine Test Facility, which had been used for testing the Saturn V F-1 engines and Space Shuttle Main engines. The MSFC was under a Space Act Agreement with Lockheed Martin Astronautics of Denver to provide a series of test firings of the Atlas III propulsion system configured with the Russian-designed RD-180 engine. The tests were designed to measure the performance of the Atlas III propulsion system, which included avionics and propellant tanks and lines, and how these components interacted with the RD-180 engine. The RD-180 is powered by kerosene and liquid oxygen, the same fuel mix used in Saturn rockets. The RD-180, the most powerful rocket engine tested at the MSFC since Saturn rocket tests in the 1960s, generated 860,000 pounds of thrust. The test was the first test ever anywhere outside Russia of a Russian designed and built engine.

  20. Quasi 1-D Study of Pulse Detonation Rocket Engine Blowdown Gasdynamics and Performance

    NASA Technical Reports Server (NTRS)

    Morris, Christopher

    2002-01-01

    This viewgraph representation provides an overview of research which develops a quasi one dimensional chemistry computational fluid dynamics code to study the effect of nozzle design on the performance of pulse detonation rocket engines (PDREs). Topics considered include: PDREs vs. steady-state rocket engines (SSREs), PDRE cycles, numerical models of idealized PDRE performance, thrust determination of PDRE, specific geometries, and nozzle design and geometry.

  1. Liquid rocket engine turbines

    NASA Technical Reports Server (NTRS)

    1974-01-01

    Criteria for the design and development of turbines for rocket engines to meet specific performance, and installation requirements are summarized. The total design problem, and design elements are identified, and the current technology pertaining to these elements is described. Recommended practices for achieving a successful design are included.

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

    NASA Technical Reports Server (NTRS)

    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.

  3. Small, high-performance engine component technology status. [liquid rocket engine for spacecraft

    NASA Technical Reports Server (NTRS)

    Yost, M. C.; Csomor, A.

    1976-01-01

    Rocketdyne, under contract to NASA-Lewis Research Center, is engaged in developing the technology for major subsystems of a small, high-performance, liquid rocket engine with the capabilities required for high-energy, upper-stage applications. Included in these efforts are the high-pressure turbopumps, preburner, igniter components, and high-area-ratio thrust chamber. A brief description of the engine system and its subsystems is given. The status of the programs is reviewed and the results of the testing conducted are presented.

  4. Simplified Analysis of Pulse Detonation Rocket Engine Blowdown Gasdynamics and Performance

    NASA Technical Reports Server (NTRS)

    Morris, C. I.; Rodgers, Stephen L. (Technical Monitor)

    2002-01-01

    Pulse detonation rocket engines (PDREs) offer potential performance improvements over conventional designs, but represent a challenging modellng task. A simplified model for an idealized, straight-tube, single-shot PDRE blowdown process and thrust determination is described and implemented. In order to form an assessment of the accuracy of the model, the flowfield time history is compared to experimental data from Stanford University. Parametric Studies of the effect of mixture stoichiometry, initial fill temperature, and blowdown pressure ratio on the performance of a PDRE are performed using the model. PDRE performance is also compared with a conventional steady-state rocket engine over a range of pressure ratios using similar gasdynamic assumptions.

  5. Quasi 1-D Study of Pulse Detonation Rocket Engine Blowdown Gasdynamics and Performance

    NASA Technical Reports Server (NTRS)

    Morris, Christopher I.

    2002-01-01

    Pulse detonation rocket engines (PDREs) offer potential performance improvements over conventional designs, but represent a challenging modeling task. A quasi 1-D, finite-rate chemistry CFD model for a PDRE is described and implemented. A parametric study of the effect of blowdown pressure ratio on the performance of several different PDRE nozzle configurations is reported.

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

    NASA Astrophysics Data System (ADS)

    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

  7. Air-Breathing Rocket Engines

    NASA Technical Reports Server (NTRS)

    1998-01-01

    This photograph depicts an air-breathing rocket engine prototype in the test bay at the General Applied Science Lab facility in Ronkonkoma, New York. Air-breathing engines, known as rocket based, combined-cycle engines, get their initial take-off power from specially designed rockets, called air-augmented rockets, that boost performance about 15 percent over conventional rockets. When the vehicle's velocity reaches twice the speed of sound, the rockets are turned off and the engine relies totally on oxygen in the atmosphere to burn hydrogen fuel, as opposed to a rocket that must carry its own oxygen, thus reducing weight and flight costs. Once the vehicle has accelerated to about 10 times the speed of sound, the engine converts to a conventional rocket-powered system to propel the craft into orbit or sustain it to suborbital flight speed. NASA's Advanced Space Transportation Program at Marshall Space Flight Center, along with several industry partners and collegiate forces, is developing this technology to make space transportation affordable for everyone from business travelers to tourists. The goal is to reduce launch costs from today's price tag of $10,000 per pound to only hundreds of dollars per pound. NASA's series of hypersonic flight demonstrators currently include three air-breathing vehicles: the X-43A, X-43B and X-43C.

  8. Rocket Engine Oscillation Diagnostics

    NASA Technical Reports Server (NTRS)

    Nesman, Tom; Turner, James E. (Technical Monitor)

    2002-01-01

    Rocket engine oscillating data can reveal many physical phenomena ranging from unsteady flow and acoustics to rotordynamics and structural dynamics. Because of this, engine diagnostics based on oscillation data should employ both signal analysis and physical modeling. This paper describes an approach to rocket engine oscillation diagnostics, types of problems encountered, and example problems solved. Determination of design guidelines and environments (or loads) from oscillating phenomena is required during initial stages of rocket engine design, while the additional tasks of health monitoring, incipient failure detection, and anomaly diagnostics occur during engine development and operation. Oscillations in rocket engines are typically related to flow driven acoustics, flow excited structures, or rotational forces. Additional sources of oscillatory energy are combustion and cavitation. Included in the example problems is a sampling of signal analysis tools employed in diagnostics. The rocket engine hardware includes combustion devices, valves, turbopumps, and ducts. Simple models of an oscillating fluid system or structure can be constructed to estimate pertinent dynamic parameters governing the unsteady behavior of engine systems or components. In the example problems it is shown that simple physical modeling when combined with signal analysis can be successfully employed to diagnose complex rocket engine oscillatory phenomena.

  9. Robust Rocket Engine Concept

    NASA Technical Reports Server (NTRS)

    Lorenzo, Carl F.

    1995-01-01

    The potential for a revolutionary step in the durability of reusable rocket engines is made possible by the combination of several emerging technologies. The recent creation and analytical demonstration of life extending (or damage mitigating) control technology enables rapid rocket engine transients with minimum fatigue and creep damage. This technology has been further enhanced by the formulation of very simple but conservative continuum damage models. These new ideas when combined with recent advances in multidisciplinary optimization provide the potential for a large (revolutionary) step in reusable rocket engine durability. This concept has been named the robust rocket engine concept (RREC) and is the basic contribution of this paper. The concept also includes consideration of design innovations to minimize critical point damage.

  10. Rocket engine numerical simulation

    NASA Technical Reports Server (NTRS)

    Davidian, Ken

    1993-01-01

    The topics are presented in view graph form and include the following: a definition of the rocket engine numerical simulator (RENS); objectives; justification; approach; potential applications; potential users; RENS work flowchart; RENS prototype; and conclusions.

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

    NASA Technical Reports Server (NTRS)

    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

  12. Liquid rocket engine nozzles

    NASA Technical Reports Server (NTRS)

    1976-01-01

    The nozzle is a major component of a rocket engine, having a significant influence on the overall engine performance and representing a large fraction of the engine structure. The design of the nozzle consists of solving simultaneously two different problems: the definition of the shape of the wall that forms the expansion surface, and the delineation of the nozzle structure and hydraulic system. This monography addresses both of these problems. The shape of the wall is considered from immediately upstream of the throat to the nozzle exit for both bell and annular (or plug) nozzles. Important aspects of the methods used to generate nozzle wall shapes are covered for maximum-performance shapes and for nozzle contours based on criteria other than performance. The discussion of structure and hydraulics covers problem areas of regeneratively cooled tube-wall nozzles and extensions; it treats also nozzle extensions cooled by turbine exhaust gas, ablation-cooled extensions, and radiation-cooled extensions. The techniques that best enable the designer to develop the nozzle structure with as little difficulty as possible and at the lowest cost consistent with minimum weight and specified performance are described.

  13. Liquid Rocket Engine Testing

    NASA Technical Reports Server (NTRS)

    Rahman, Shamim

    2005-01-01

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

  14. Rocket engine condition monitoring system

    SciTech Connect

    Hagar, S.K.; Alcock, J.F.

    1989-01-01

    It is expected that the Rocket Engine Condition Monitoring System (RECMS) program will define engine monitoring technologies and an integration approach which can be applied to engine development in support of advanced launch system objectives. The RECMS program approaches engine monitoring as a system which is fully integrated with the engine controller, vehicle monitoring system, and ground processing systems to ensure mission success in addition to engine reliability. The system components are monitored through health and performance sensors; they are analyzed with the diagnostic and prognostic algorithms and demonstrated by system testing with hardware from other advanced development programs.

  15. Quasi-One-Dimensional Modeling of Pulse Detonation Rocket Engines

    NASA Technical Reports Server (NTRS)

    Morris, Christopher I.

    2003-01-01

    This viewgraph presentation provides information on the engine cycle of a pulse detonation rocket engine (PDRE), models for optimizing the performance of a PDRE, and the performance of PDREs in comparison to Solid State Rocket Engines (SSREs).

  16. Simplified Analysis of Pulse Detonation Rocket Engine B1owdown Gasdynamics and Performance

    NASA Technical Reports Server (NTRS)

    Morris, Christopher I.

    2001-01-01

    Pulsed detonation rocket engines (PDREs) have generated considerable 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 a great deal of 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 difficulties in comparing the available experimental measurements with numerical models. The goal of this paper is to improve understanding of PDRE blowdown gasdynamics and performance issues through use of a simplified model that captures the essential features of the unsteady blowdown process, and yet remains computationally inexpensive. The PDRE system studied here is highly idealized, consisting of a constant-area detonation tube with one end closed and the other end open to the environment. The tube is prefilled with a gaseous propellant mixture with no initial velocity or outflow to the environment. The detonation is initiated instantaneously at the closed end of the device. Chapman-Jouguet (C-J) post-detonation gas conditions are calculated using the CET89 version of the NASA thermochemical code. The I-D, unsteady method of characteristics is used to calculate the flowfield following the detonation front. See the compressible flow texts by Thompson and Zucrow and Hoffman for details of this method. Parametric studies of the effect of mixture stoichiometry, fill temperature, and blowdown pressure ratio on performance are reported. A comparison of the performance of an idealized straight-tube PDRE with a conventional steady

  17. Liquid Rocket Engine Testing Overview

    NASA Technical Reports Server (NTRS)

    Rahman, Shamim

    2005-01-01

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

  18. Program For Optimization Of Nuclear Rocket Engines

    NASA Technical Reports Server (NTRS)

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

    1994-01-01

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

  19. Thrust Performance Improvement for a Water/Liquid Nitrogen Rocket Engine

    NASA Astrophysics Data System (ADS)

    Watanabe, Rikio; Mikami, Ryo

    We propose a water/liquid nitrogen rocket engine as a new non-combustion type rocket engine. Liquid nitrogen is mixed with heated water and specific volume of nitrogen is increased by evaporation. Thrust force is obtained by exhaust of nitrogen gas through a nozzle with water particles. Results of previous experiments indicated a specific impulse is 60 % of the theoretically estimated value. By evaluating the characteristic exhaust velocity and other thrust characteristics, we found that the lower-than-expected specific impulse is due to insufficient propellant mixing and heat transfer between heated water and liquid nitrogen in the mixing chamber. We also performed high-speed imaging experiments to visualize impinging and mixing of propellants. Results indicate that in the original injection setup, heat conveyed by heated water is not adequately transferred to the liquid nitrogen. An alternative injection pattern was tested, which resulted in a 10% increase in the characteristic exhaust velocity. In addition, we tested a new type of injector designed for more efficient mixing and heat transfer that exhibited 30 % increase in characteristic exhaust velocity. Furthermore, we modified the theoretical expression for the characteristic exhaust velocity based on multi-phased flow theory so that it agrees well with the experimental results.

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

    NASA Technical Reports Server (NTRS)

    Wang, Ten-See

    2004-01-01

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

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

    NASA Technical Reports Server (NTRS)

    Wang, Ten-See

    2004-01-01

    The objectives of this study are to conduct a unified computational analysis for computing the design parameters such as axial thrust, convective and radiative wall heat fluxes for liquid rocket engine nozzles, so as to develop a computational strategy for computing those design parameters through parametric investigations. The computational methodology is based on a multidimensional, finite-volume, turbulent, chemically reacting, radiating, unstructured-grid, and pressure-based formulation, with grid refinement capabilities. Systematic parametric studies on effects of wail boundary conditions, combustion chemistry, radiation coupling, computational cell shape, and grid refinement were performed and assessed. Comparisons of the computed axial thrust performance, flow features, and wail heat fluxes with those of available test data and design calculations are presented.

  2. ION ROCKET ENGINE

    DOEpatents

    Ehlers, K.W.; Voelker, F. III

    1961-12-19

    A thrust generating engine utilizing cesium vapor as the propellant fuel is designed. The cesium is vaporized by heat and is passed through a heated porous tungsten electrode whereby each cesium atom is fonized. Upon emergfng from the tungsten electrode, the ions are accelerated rearwardly from the rocket through an electric field between the tungsten electrode and an adjacent accelerating electrode grid structure. To avoid creating a large negative charge on the space craft as a result of the expulsion of the positive ions, a source of electrons is disposed adjacent the ion stream to neutralize the cesium atoms following acceleration thereof. (AEC)

  3. SRB-3D Solid Rocket Booster performance prediction program. Volume 1: Engineering description/users information manual

    NASA Technical Reports Server (NTRS)

    Winkler, J. C.

    1976-01-01

    The modified Solid Rocket Booster Performance Evaluation Model (SRB-3D) was developed as an extension to the internal ballistics module of the SRB-2 performance program. This manual contains the engineering description of SRB-3D which describes the approach used to develop the 3D concept and an explanation of the modifications which were necessary to implement these concepts.

  4. Feasibility Demonstration of a 445N High-Performance Rocket Engine

    NASA Technical Reports Server (NTRS)

    Appel, Marshall A.; Schoenman, Leonard; Franklin, Jerrold E.; Lansaw, P. Tina

    1989-01-01

    A program to demonstrate the feasibility of a high-performance 445 Newton (100 lb(sub f)) NTO/NMH rocket engine is presented. An existing high-performance injector was coupled with three different thrust chambers to acquire test data for the program. A stainless-steel sea-level chamber was used for test stand checkout, calibration and initial injector performance determinations. Two high-temperature iridium lined rhenium thrust chambers were used to determine altitude performance and durability. The first chamber was tested for a total duration of 3381 s and the second was tested for a total duration of 15,000 s with no measurable degradation. Extrapolation of epsilon = 44:1 experimental data to the recommended epsilon = 467:1 expansion nozzle, using the JANNAF methodology, showed that a performance of 3138 to 3167 N-s/kg (320 to 323 lb(sub f-s)/lb(sub m) could be achieved with the present design. Recommendations are presented which would allow a redesigned engine to achieve the 3195 N-s/kg (326 lb(sub f-s)/lb(sub m)) program goal.

  5. Coal-Fired Rocket Engine

    NASA Technical Reports Server (NTRS)

    Anderson, Floyd A.

    1987-01-01

    Brief report describes concept for coal-burning hybrid rocket engine. Proposed engine carries larger payload, burns more cleanly, and safer to manufacture and handle than conventional solid-propellant rockets. Thrust changeable in flight, and stops and starts on demand.

  6. Air-breathing Rocket Engine Test

    NASA Technical Reports Server (NTRS)

    1999-01-01

    This Quick Time movie depicts the Rocketdyne static test of an air-breathing rocket. Air-breathing engines, known as rocket based, combined-cycle engines, get their initial take-off power from specially designed rockets, called air-augmented rockets, that boost performance about 15 percent over conventional rockets. When the vehicle's velocity reaches twice the speed of sound, the rockets are turned off and the engine relies totally on oxygen in the atmosphere to burn hydrogen fuel, as opposed to a rocket that must carry its own oxygen, thus reducing weight and flight costs. Once the vehicle has accelerated to about 10 times the speed of sound, the engine converts to a conventional rocket-powered system to propel the craft into orbit or sustain it to suborbital flight speed. NASA's advanced Transportation Program at the Marshall Space Flight Center, along with several industry partners and collegiate forces, is developing this technology to make space transportation affordable for everyone from business travelers to tourists. The goal is to reduce launch costs from today's price tag of $10,000 per pound to only hundreds of dollars per pound. NASA's series of hypersonic flight demonstrators currently include three air-breathing vehicles: the X-43A, X-43B and X-43C.

  7. J-2S rocket engine

    NASA Astrophysics Data System (ADS)

    Vilja, J. O.; Briley, G. L.; Murphy, T. H.

    1993-06-01

    The principal design characteristics and features of the J-2S rocket engine, developed as a simpler and more robust version of the J-2 engine, are described. The J-2S is a 265,000-lb vacuum thrust engine that delivers 436 sec vacuum thrust with a nozzle expansion ratio of 40 and operates at a chamber pressure of 1,200 psi. The most unique feature of the J-2S is that it incorporates a main chamber tap-off cycle which eliminates the need for a gas generator. Another simplification for the J-2S is the adoption of a centrifugal fuel turbopump to replace the J-2's axial turbopump. A schematic of the J-2S, engine test results, and performance options are presented.

  8. Nitrous Oxide/Paraffin Hybrid Rocket Engines

    NASA Technical Reports Server (NTRS)

    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.

  9. Air-Breathing Rocket Engine Test

    NASA Technical Reports Server (NTRS)

    2000-01-01

    This photograph depicts an air-breathing rocket engine that completed an hour or 3,600 seconds of testing at the General Applied Sciences Laboratory in Ronkonkoma, New York. Referred to as ARGO by its design team, the engine is named after the mythological Greek ship that bore Jason and the Argonauts on their epic voyage of discovery. Air-breathing engines, known as rocket based, combined-cycle engines, get their initial take-off power from specially designed rockets, called air-augmented rockets, that boost performance about 15 percent over conventional rockets. When the vehicle's velocity reaches twice the speed of sound, the rockets are turned off and the engine relies totally on oxygen in the atmosphere to burn hydrogen fuel, as opposed to a rocket that must carry its own oxygen, thus reducing weight and flight costs. Once the vehicle has accelerated to about 10 times the speed of sound, the engine converts to a conventional rocket-powered system to propel the craft into orbit or sustain it to suborbital flight speed. NASA's Advanced SpaceTransportation Program at Marshall Space Flight Center, along with several industry partners and collegiate forces, is developing this technology to make space transportation affordable for everyone from business travelers to tourists. The goal is to reduce launch costs from today's price tag of $10,000 per pound to only hundreds of dollars per pound. NASA's series of hypersonic flight demonstrators currently include three air-breathing vehicles: the X-43A, X-43B and X-43C.

  10. Performance and heat transfer characteristics of a carbon monoxide/oxygen rocket engine

    NASA Technical Reports Server (NTRS)

    Linne, Diane L.

    1993-01-01

    The combustion and heat transfer characteristics of a carbon monoxide and oxygen rocket engine were evaluated. The test hardware consisted of a calorimeter combustion chamber with a heat sink nozzle and an eighteen element concentric tube injector. Experimental results are given at chamber pressures of 1070 and 3070 kPa, and over a mixture ratio range of 0.3 to 1.0. Experimental C efficiency was between 95 and 96.5 percent. Heat transfer results are discussed both as a function of mixture ratio and axial distance in the chamber. They are also compared to a Nusselt number correlation for fully developed turbulent flow.

  11. Combustion performance of bipropellant liquid rocket engine combustors with fuel-impingement cooling

    SciTech Connect

    Jiang, T.L.; Chiang, W.; Jang, S.

    1995-05-01

    In order to obtain an accurate combustion analyses which are important in the thruster design of modern advanced liquid rocket engine, flow analysis should be conducted from the injector phase down to the propulsive nozzle throat. Thus, in the present study, flow analysis for the axisymmetric thrust chamber of an OMV(exp 3) installed with a pintle-type ring-shaped injector and a conical convergent nozzle is conducted. Liquid monomethyl hydrazine (MMH) and nitrogen tetroxide (NTO) storable bipropellants are used as fuel and oxidizer sources. An optimum injected fuel and oxidizer droplet-size combination is proposed. Finally, the results obtained are presented. 4 refs.

  12. Future Modeling Needs in Pulse Detonation Rocket Engine Design

    NASA Technical Reports Server (NTRS)

    Meade, Brian; Talley, Doug; Mueller, Donn; Tew, Dave; Guidos, Mike; Seymour, Dave

    2001-01-01

    This paper presents a performance model rocket engine design that takes advantage of pulse detonation to generate thrust. The contents include: 1) Introduction to the Pulse Detonation Rocket Engine (PDRE); 2) PDRE modeling issues and options; 3) Discussion of the PDRE Performance Workshop held at Marshall Space Flight Center; and 4) Identify needs involving an open performance model for Pulse Detonation Rocket Engines. This paper is in viewgraph form.

  13. Composite Material Application to Liquid Rocket Engines

    NASA Technical Reports Server (NTRS)

    Judd, D. C.

    1982-01-01

    The substitution of reinforced plastic composite (RPC) materials for metal was studied. The major objectives were to: (1) determine the extent to which composite materials can be beneficially used in liquid rocket engines; (2) identify additional technology requirements; and (3) determine those areas which have the greatest potential for return. Weight savings, fabrication costs, performance, life, and maintainability factors were considered. Two baseline designs, representative of Earth to orbit and orbit to orbit engine systems, were selected. Weight savings are found to be possible for selected components with the substitution of materials for metal. Various technology needs are identified before RPC material can be used in rocket engine applications.

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

    NASA Technical Reports Server (NTRS)

    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.

  15. Hybrid rocket performance

    NASA Astrophysics Data System (ADS)

    Frederick, Robert A., Jr.

    1992-12-01

    A hybrid rocket is a system consisting of a solid fuel grain and a gaseous or liquid oxidizer. Figure 1 shows three popular hybrid propulsion cycles that are under current consideration. NASA MSFC has teamed with industry to test two hybrid propulsion systems that will allow scaling to motors of potential interest for Titan and Atlas systems, as well as encompassing the range of interest for SEI lunar ascent stages and National Launch System Cargo Transfer Vehicle (NLS CTV) and NLS deorbit systems. Hybrid systems also offer advantages as moderate-cost, environmentally acceptable propulsion system. The objective of this work was to recommend a performance prediction methodology for hybrid rocket motors. The scope included completion of: a literature review, a general methodology, and a simplified performance model.

  16. Hybrid rocket performance

    NASA Technical Reports Server (NTRS)

    Frederick, Robert A., Jr.

    1992-01-01

    A hybrid rocket is a system consisting of a solid fuel grain and a gaseous or liquid oxidizer. Figure 1 shows three popular hybrid propulsion cycles that are under current consideration. NASA MSFC has teamed with industry to test two hybrid propulsion systems that will allow scaling to motors of potential interest for Titan and Atlas systems, as well as encompassing the range of interest for SEI lunar ascent stages and National Launch System Cargo Transfer Vehicle (NLS CTV) and NLS deorbit systems. Hybrid systems also offer advantages as moderate-cost, environmentally acceptable propulsion system. The objective of this work was to recommend a performance prediction methodology for hybrid rocket motors. The scope included completion of: a literature review, a general methodology, and a simplified performance model.

  17. The Strutjet Rocket Based Combined Cycle Engine

    NASA Technical Reports Server (NTRS)

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

    1998-01-01

    The multi stage chemical rocket has been established over many years as the propulsion System for space transportation vehicles, while, at the same time, there is increasing concern about its continued affordability and rather involved reusability. Two broad approaches to addressing this overall launch cost problem consist in one, the further development of the rocket motor, and two, the use of airbreathing propulsion to the maximum extent possible as a complement to the limited use of a conventional rocket. In both cases, a single-stage-to-orbit (SSTO) vehicle is considered a desirable goal. However, neither the "all-rocket" nor the "all-airbreathing" approach seems realizable and workable in practice without appreciable advances in materials and manufacturing. An affordable system must be reusable with minimal refurbishing on-ground, and large mean time between overhauls, and thus with high margins in design. It has been suggested that one may use different engine cycles, some rocket and others airbreathing, in a combination over a flight trajectory, but this approach does not lead to a converged solution with thrust-to-mass, specific impulse, and other performance and operational characteristics that can be obtained in the different engines. The reason is this type of engine is simply a combination of different engines with no commonality of gas flowpath or components, and therefore tends to have the deficiencies of each of the combined engines. A further development in this approach is a truly combined cycle that incorporates a series of cycles for different modes of propulsion along a flight path with multiple use of a set of components and an essentially single gas flowpath through the engine. This integrated approach is based on realizing the benefits of both a rocket engine and airbreathing engine in various combinations by a systematic functional integration of components in an engine class usually referred to as a rocket-based combined cycle (RBCC) engine

  18. Low-thrust chemical rocket engine study

    NASA Technical Reports Server (NTRS)

    Mellish, J. A.

    1981-01-01

    Engine data and information are presented to perform system studies on cargo orbit-transfer vehicles which would deliver large space structures to geosynchronous equatorial orbit. Low-thrust engine performance, weight, and envelope parametric data were established, preliminary design information was generated, and technologies for liquid rocket engines were identified. Two major engine design drivers were considered in the study: cooling and engine cycle options. Both film-cooled and regeneratively cooled engines were evaluated. The propellant combinations studied were hydrogen/oxygen, methane/oxygen, and kerosene/oxygen.

  19. Centrifugal pumps for rocket engines

    NASA Technical Reports Server (NTRS)

    Campbell, W. E.; Farquhar, J.

    1974-01-01

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

  20. Outbrief - Long Life Rocket Engine Panel

    NASA Technical Reports Server (NTRS)

    Quinn, Jason Eugene

    2004-01-01

    This white paper is an overview of the JANNAF Long Life Rocket Engine (LLRE) Panel results from the last several years of activity. The LLRE Panel has met over the last several years in order to develop an approach for the development of long life rocket engines. Membership for this panel was drawn from a diverse set of the groups currently working on rocket engines (Le. government labs, both large and small companies and university members). The LLRE Panel was formed in order to determine the best way to enable the design of rocket engine systems that have life capability greater than 500 cycles while meeting or exceeding current performance levels (Specific Impulse and Thrust/Weight) with a 1/1,OOO,OOO likelihood of vehicle loss due to rocket system failure. After several meetings and much independent work the panel reached a consensus opinion that the primary issues preventing LLRE are a lack of: physics based life prediction, combined loads prediction, understanding of material microphysics, cost effective system level testing. and the inclusion of fabrication process effects into physics based models. With the expected level of funding devoted to LLRE development, the panel recommended that fundamental research efforts focused on these five areas be emphasized.

  1. Transpiration And Regenerative Cooling Of Rocket Engine

    NASA Technical Reports Server (NTRS)

    Obrien, Charles J.

    1989-01-01

    Transpiration cooling extends limits of performance. Addition of transpiration cooling to regeneratively-cooled rocket-engine combustion chamber proposed. Modification improves performance of engine by allowing use of higher chamber pressure. Throat section of combustion-chamber wall cooled by transpiration, while chamber and nozzle sections cooled by fluid flowing in closed channels. Concept applicable to advanced, high-performance terrestrial engines or some kinds of industrial combustion chambers. With proper design, cooling scheme makes possible to achieve higher chamber pressure and higher overall performance in smaller engine.

  2. Developments in REDES: The rocket engine design expert system

    NASA Technical Reports Server (NTRS)

    Davidian, Kenneth O.

    1990-01-01

    The Rocket Engine Design Expert System (REDES) is being developed at the 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 is built in the Knowledge Engineering Environment (KEE) expert system shell and runs on a Sun 4/110 computer.

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

    NASA Technical Reports Server (NTRS)

    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.

  4. Transpiration cooled throat for hydrocarbon rocket engines

    NASA Technical Reports Server (NTRS)

    May, Lee R.; Burkhardt, Wendel M.

    1991-01-01

    The objective for the Transpiration Cooled Throat for Hydrocarbon Rocket Engines Program was to characterize the use of hydrocarbon fuels as transpiration coolants for rocket nozzle throats. The hydrocarbon fuels investigated in this program were RP-1 and methane. To adequately characterize the above transpiration coolants, a program was planned which would (1) predict engine system performance and life enhancements due to transpiration cooling of the throat region using analytical models, anchored with available data; (2) a versatile transpiration cooled subscale rocket thrust chamber was designed and fabricated; (3) the subscale thrust chamber was tested over a limited range of conditions, e.g., coolant type, chamber pressure, transpiration cooled length, and coolant flow rate; and (4) detailed data analyses were conducted to determine the relationship between the key performance and life enhancement variables.

  5. Rocket Engine Altitude Simulation Technologies

    NASA Technical Reports Server (NTRS)

    Woods, Jody L.; Lansaw, John

    2010-01-01

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

  6. Experimental Performance Evaluation of a Supersonic Turbine for Rocket Engine Applications

    NASA Technical Reports Server (NTRS)

    Snellgrove, Lauren M.; Griffin, Lisa W.; Sieja, James P.; Huber, Frank W.

    2003-01-01

    In order to mitigate the risk of rocket propulsion development, efficient, accurate, detailed fluid dynamics analysis and testing of the turbomachinery is necessary. To support this requirement, a task was developed at NASA Marshall Space Flight Center (MSFC) to improve turbine aerodynamic performance through the application of advanced design and analysis tools. These tools were applied to optimize a supersonic turbine design suitable for a reusable launch vehicle (RLV). The hot gas path and blading were redesigned-to obtain an increased efficiency. The goal of the demonstration was to increase the total-to- static efficiency of the turbine by eight points over the baseline design. A sub-scale, cold flow test article modeling the final optimized turbine was designed, manufactured, and tested in air at MSFC s Turbine Airflow Facility. Extensive on- and off- design point performance data, steady-state data, and unsteady blade loading data were collected during testing.

  7. Rocket-in-a-Duct Performance Analysis

    NASA Technical Reports Server (NTRS)

    Schneider, Steven J.; Reed, Brian D.

    1999-01-01

    An axisymmetric, 110 N class, rocket configured with a free expansion between the rocket nozzle and a surrounding duct was tested in an altitude simulation facility. The propellants were gaseous hydrogen and gaseous oxygen and the hardware consisted of a heat sink type copper rocket firing through copper ducts of various diameters and lengths. A secondary flow of nitrogen was introduced at the blind end of the duct to mix with the primary rocket mass flow in the duct. This flow was in the range of 0 to 10% of the primary massflow and its effect on nozzle performance was measured. The random measurement errors on thrust and massflow were within +/-1%. One dimensional equilibrium calculations were used to establish the possible theoretical performance of these rocket-in-a-duct nozzles. Although the scale of these tests was small, they simulated the relevant flow expansion physics at a modest experimental cost. Test results indicated that lower performance was obtained at higher free expansion area ratios and longer ducts, while, higher performance was obtained with the addition of secondary flow. There was a discernable peak in specific impulse efficiency at 4% secondary flow. The small scale of these tests resulted in low performance efficiencies, but prior numerical modeling of larger rocket-in-a-duct engines predicted performance that was comparable to that of optimized rocket nozzles. This remains to be proven in large-scale, rocket-in-a-duct tests.

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

    NASA Technical Reports Server (NTRS)

    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.

  9. Measuring Model Rocket Engine Thrust Curves

    ERIC Educational Resources Information Center

    Penn, Kim; Slaton, William V.

    2010-01-01

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

  10. Integrated engineering modeling for air breathing rockets

    NASA Astrophysics Data System (ADS)

    Chitilappilly, Lazar T.; Subramanyam, J. D. A.

    An innovative aerodynamic-propulsion-flight integrated modeling is carried out for airbreathing rockets, the propulsion of which has primary dependence on flight conditions. The integrated modeling is highly beneficial for design and analysis of accelerating air breathing rockets characterized by continuously varying flight conditions. The details of the modeling is described; the force accounting, trajectory analysis, solving the flow in the sub-systems (air intake, primary rocket, secondary combustion chamber and secondary nozzle), matching the subsystem flow fields and determining the mode of operation. Operational features are listed of the computer software developed, air breathing integrated design and analysis engineering software. It gives all the propulsion and flight parameters from take-off of the rocket to end of flight and has been instrumental in the design of the research air breathing rocket ABR-200(I). The hundreds of flight performance analyses required for design is possible by the engineering approach adopted for solving the propulsor flow field. The software results are compared with ejector mode and connected pipe mode static tests. The overall validation of the software is achieved by flight tests; the performance predictions have matched exactly with that measured during thee first and second flights of the ABR-200(I).

  11. Reusable Rocket Engine Maintenance Study

    NASA Technical Reports Server (NTRS)

    Macgregor, C. A.

    1982-01-01

    Approximately 85,000 liquid rocket engine failure reports, obtained from 30 years of developing and delivering major pump feed engines, were reviewed and screened and reduced to 1771. These were categorized into 16 different failure modes. Failure propagation diagrams were established. The state of the art of engine condition monitoring for in-flight sensors and between flight inspection technology was determined. For the 16 failure modes, the potential measurands and diagnostic requirements were identified, assessed and ranked. Eight areas are identified requiring advanced technology development.

  12. High-Temperature Rocket Engine

    NASA Technical Reports Server (NTRS)

    Schneider, Steven J.; Rosenberg, Sanders D.; Chazen, Melvin L.

    1994-01-01

    Two rocket engines that operate at temperature of 2,500 K designed to provide thrust for station-keeping adjustments of geosynchronous satellites, for raising and lowering orbits, and for changing orbital planes. Also useful as final propulsion stages of launch vehicles delivering small satellites to low orbits around Earth. With further development, engines used on planetary exploration missions for orbital maneuvers. High-temperature technology of engines adaptable to gas-turbine combustors, ramjets, scramjets, and hot components of many energy-conversion systems.

  13. Aluminum/hydrocarbon gel propellants: An experimental and theoretical investigation of secondary atomization and predicted rocket engine performance

    NASA Astrophysics Data System (ADS)

    Mueller, Donn Christopher

    1997-12-01

    Experimental and theoretical investigations of aluminum/hydrocarbon gel propellant secondary atomization and its potential effects on rocket engine performance were conducted. In the experimental efforts, a dilute, polydisperse, gel droplet spray was injected into the postflame region of a burner and droplet size distributions was measured as a function of position above the burner using a laser-based sizing/velocimetry technique. The sizing/velocimetry technique was developed to measure droplets in the 10-125 mum size range and avoids size-biased detection through the use of a uniformly illuminated probe volume. The technique was used to determine particle size distributions and velocities at various axial locations above the burner for JP-10, and 50 and 60 wt% aluminum gels. Droplet shell formation models were applied to aluminum/hydrocarbon gels to examine particle size and mass loading effects on the minimum droplet diameter that will permit secondary atomization. This diameter was predicted to be 38.1 and 34.7 mum for the 50 and 60 wt% gels, which is somewhat greater than the experimentally measured 30 and 25 mum diameters. In the theoretical efforts, three models were developed and an existing rocket code was exercised to gain insights into secondary atomization. The first model was designed to predict gel droplet properties and shell stresses after rigid shell formation, while the second, a one-dimensional gel spray combustion model was created to quantify the secondary atomization process. Experimental and numerical comparisons verify that secondary atomization occurs in 10-125 mum diameter particles although an exact model could not be derived. The third model, a one-dimensional gel-fueled rocket combustion chamber, was developed to evaluate secondary atomization effects on various engine performance parameters. Results show that only modest secondary atomization may be required to reduce propellant burnout distance and radiation losses. A solid propellant

  14. Low-thrust chemical rocket engine study

    NASA Technical Reports Server (NTRS)

    Shoji, J. M.

    1981-01-01

    An analytical study evaluating thrust chamber cooling engine cycles and preliminary engine design for low thrust chemical rocket engines for orbit transfer vehicles is described. Oxygen/hydrogen, oxygen/methane, and oxygen/RP-1 engines with thrust levels from 444.8 N to 13345 N, and chamber pressures from 13.8 N/sq cm to 689.5 N/sq cm were evaluated. The physical and thermodynamic properties of the propellant theoretical performance data, and transport properties are documented. The thrust chamber cooling limits for regenerative/radiation and film/radiation cooling are defined and parametric heat transfer data presented. A conceptual evaluation of a number of engine cycles was performed and a 2224.1 N oxygen/hydrogen engine cycle configuration and a 2224.1 N oxygen/methane configuration chosen for preliminary engine design. Updated parametric engine data, engine design drawings, and an assessment of technology required are presented.

  15. Unique nuclear thermal rocket engine

    NASA Astrophysics Data System (ADS)

    Culver, Donald W.; Rochow Wilcox Space; Nuclear Systems, Richard

    1993-01-01

    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.

  16. Unique nuclear thermal rocket engine

    NASA Astrophysics Data System (ADS)

    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.

  17. Unique nuclear thermal rocket engine

    SciTech Connect

    Culver, D.W. ); Rochow, R. )

    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.

  18. Liquid rocket engine turbopump gears

    NASA Technical Reports Server (NTRS)

    1974-01-01

    Design and fabrication of gear drives for rocket engine turbopumps are described in the sequence encountered during the design process as follows: (1) selection of overall arrangement; (2) selection of gear type; (3) preliminary sizing; (4) lubrication system design; (5) detail tooth design; (6) selection of gear materials; and (7) gear fabrication and testing as it affects the design. The description is oriented towards the use of involute spur gears, although reference material for helical gears is also cited.

  19. Rocket Engine Numerical Simulator (RENS)

    NASA Technical Reports Server (NTRS)

    Davidian, Kenneth O.

    1997-01-01

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

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

    NASA Technical Reports Server (NTRS)

    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.

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

    NASA Astrophysics Data System (ADS)

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

    NASA Technical Reports Server (NTRS)

    Wang, Ten-See; Chen, Yen-Sen

    1990-01-01

    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.

  3. Liquid rocket engine test facility engineering challenges

    NASA Astrophysics Data System (ADS)

    Ellerbrock, Hartwig; Ziegenhagen, Stefan

    2006-12-01

    Liquid rocket engines for launch vehicles and space crafts as well as their subsystems need to be verified and qualified during hot-runs. A high test cadence combined with a flexible test team helps to reduce the cost for test verification during development/qualification as well as during acceptance testing for production. Test facility intelligence allows to test subsystems in the same manner as during complete engine system tests and will therefore reduce development time and cost. This paper gives an overview of the maturing of test engineering know how for rocket engine test stands as well as high altitude test stands for small propulsion thrusters at EADS-ST in Ottobrunn and Lampoldshausen and is split into two parts: Part 1 gives a historical overview of the EADS-ST test stands at Ottobrunn and Lampoldshausen since the beginning of Rocket propulsion activities in the 1960s. Part 2 gives an overview of the actual test capabilities and the test engineering know-how for test stand construction/adaptation and their use during running programs. Examples of actual realised facility concepts are given to demonstrate cost saving potential for test programs in both cases for development/qualification issues as well as for production purposes.

  4. Development Status of Reusable Rocket Engine

    NASA Astrophysics Data System (ADS)

    Yoshida, Makoto; Takada, Satoshi; Naruo, Yoshihiro; Niu, Kenichi

    A 30-kN rocket engine, a pilot engine, is being developed in Japan. Development of this pilot engine has been initiated in relation to a reusable sounding rocket, which is also being developed in Japan. This rocket takes off vertically, reaches an altitude of 100 km, lands vertically at the launch site, and is launched again within several days. Due to advantage of reusability, successful development of this rocket will mean that observation missions can be carried out more frequently and economically. In order to realize this rocket concept, the engines installed on the rocket should be characterized by reusability, long life, deep throttling and health monitoring, features which have not yet been established in Japanese rocket engines. To solve the engineering factors entitled by those features, a new design methodology, advanced engine simulations and engineering testing are being focused on in the pilot engine development stage. Especially in engineering testing, limit condition data is acquired to facilitate development of new diagnostic techniques, which can be applied by utilizing the mobility of small-size hardware. In this paper, the development status of the pilot engine is described, including fundamental design and engineering tests of the turbopump bearing and seal, turbine rig, injector and combustion chamber, and operation and maintenance concepts for one hundred flights by a reusable rocket are examined.

  5. Casting propellant in rocket engine

    NASA Technical Reports Server (NTRS)

    Roach, J. E.; Froehling, S. C. (Inventor)

    1976-01-01

    A method is described for casting a solid propellant in the casing of a rocket engine having a continuous wall with a single opening which is formed by leaves of a material which melt at a temperature of the propellant and with curved edges concentric to the curvature of the spherical casing. The leaves are inserted into the spherical casing through the opening forming a core having a greater width than the width of the single opening and with curved peripheral edges. The cast propellant forms a solid mass and then heated to melt the leaves and provide a central opening with radial projecting flutes.

  6. Prediction of the Thrust Performance and the Flowfield of Liquid Rocket Engines

    NASA Technical Reports Server (NTRS)

    Wang, T.-S.

    1990-01-01

    In an effort to improve the current solutions in the design and analysis of liquid propulsive engines, a computational fluid dynamics (CFD) model capable of calculating the 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, was investigated as a sample case. The CFD model, FDNS, is a pressure based, non-staggered grid, viscous/inviscid, ideal gas/real gas, reactive code. An adaptive upwinding differencing scheme is employed for the spatial discretization. The upwind scheme is based on fourth order central differencing with fourth order damping for smooth regions, and second order central differencing with second order damping for shock capturing. It is equipped with a CHMQGM equilibrium chemistry algorithm and a PARASOL finite rate chemistry algorithm using the point implicit method. The computed flow results and performance compared well with those of other standard codes and engine hot fire test data. In addition, the transient nozzle flowfield calculation was also performed to demonstrate the ability of FDNS in capturing the flow separation during the startup process.

  7. Rocket-Based Combined Cycle Engine Concept Development

    NASA Technical Reports Server (NTRS)

    Ratekin, G.; Goldman, Allen; Ortwerth, P.; Weisberg, S.; McArthur, J. Craig (Technical Monitor)

    2001-01-01

    The development of rocket-based combined cycle (RBCC) propulsion systems is part of a 12 year effort under both company funding and contract work. The concept is a fixed geometry integrated rocket, ramjet, scramjet, which is hydrogen fueled and uses hydrogen regenerative cooling. The baseline engine structural configuration uses an integral structure that eliminates panel seals, seal purge gas, and closeout side attachments. Engine A5 is the current configuration for NASA Marshall Space Flight Center (MSFC) for the ART program. Engine A5 models the complete flight engine flowpath of inlet, isolator, airbreathing combustor, and nozzle. High-performance rocket thrusters are integrated into the engine enabling both low speed air-augmented rocket (AAR) and high speed pure rocket operation. Engine A5 was tested in GASL's new Flight Acceleration Simulation Test (FAST) facility in all four operating modes, AAR, RAM, SCRAM, and Rocket. Additionally, transition from AAR to RAM and RAM to SCRAM was also demonstrated. Measured performance demonstrated vision vehicle performance levels for Mach 3 AAR operation and ramjet operation from Mach 3 to 4. SCRAM and rocket mode performance was above predictions. For the first time, testing also demonstrated transition between operating modes.

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

    NASA Technical Reports Server (NTRS)

    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.

  9. Commercial Rocket Engine Readied for Test

    NASA Video Gallery

    Engineers at NASA’s John C. Stennis Space Center recently installed an Aerojet AJ26 rocket engine for qualification testing as part of a partnership that highlights the space agency’s commitment to...

  10. The prediction of nozzle performance and heat transfer in hydrogen/oxygen rocket engines with transpiration cooling, film cooling, and high area ratios

    NASA Technical Reports Server (NTRS)

    Kacynski, Kenneth J.; Hoffman, Joe D.

    1993-01-01

    An advanced engineering computational model has been developed to aid in the analysis and design of hydrogen/oxygen chemical rocket engines. The complete multi-species, chemically reacting and diffusing Navier-Stokes equations are modelled, finite difference approach that is tailored to be conservative in an axisymmetric coordinate system for both the inviscid and viscous terms. Demonstration cases are presented for a 1030:1 area ratio nozzle, a 25 lbf film cooled nozzle, and transpiration cooled plug-and-spool rocket engine. The results indicate that the thrust coefficient predictions of the 1030:1 nozzle and the film cooled nozzle are within 0.2 to 0.5 percent, respectively, of experimental measurements when all of the chemical reaction and diffusion terms are considered. Further, the model's predictions agree very well with the heat transfer measurements made in all of the nozzle test cases. The Soret thermal diffusion term is demonstrated to have a significant effect on the predicted mass fraction of hydrogen along the wall of the nozzle in both the laminar flow 1030:1 nozzle and the turbulent plug-and-spool rocket engine analysis cases performed. Further, the Soret term was shown to represent a significant fraction of the diffusion fluxes occurring in the transpiration cooled rocket engine.

  11. The prediction of nozzle performance and heat transfer in hydrogen/oxygen rocket engines with transpiration cooling, film cooling, and high area ratios

    NASA Astrophysics Data System (ADS)

    Kacynski, Kenneth J.; Hoffman, Joe D.

    1993-11-01

    An advanced engineering computational model has been developed to aid in the analysis and design of hydrogen/oxygen chemical rocket engines. The complete multi-species, chemically reacting and diffusing Navier-Stokes equations are modelled, finite difference approach that is tailored to be conservative in an axisymmetric coordinate system for both the inviscid and viscous terms. Demonstration cases are presented for a 1030:1 area ratio nozzle, a 25 lbf film cooled nozzle, and transpiration cooled plug-and-spool rocket engine. The results indicate that the thrust coefficient predictions of the 1030:1 nozzle and the film cooled nozzle are within 0.2 to 0.5 percent, respectively, of experimental measurements when all of the chemical reaction and diffusion terms are considered. Further, the model's predictions agree very well with the heat transfer measurements made in all of the nozzle test cases. The Soret thermal diffusion term is demonstrated to have a significant effect on the predicted mass fraction of hydrogen along the wall of the nozzle in both the laminar flow 1030:1 nozzle and the turbulent plug-and-spool rocket engine analysis cases performed. Further, the Soret term was shown to represent a significant fraction of the diffusion fluxes occurring in the transpiration cooled rocket engine.

  12. Rocket Engine Thrust Chamber Assembly

    NASA Technical Reports Server (NTRS)

    Cornelius, Charles S. (Inventor); Counts, Richard H. (Inventor); Myers, W. Neill (Inventor); Lackey, Jeffrey D. (Inventor); Peters, Warren (Inventor); Shadoan, Michael D. (Inventor); Sparks, David L. (Inventor); Lawrence, Timothy W. (Inventor)

    2001-01-01

    A thrust chamber assembly for liquid fueled rocket engines and the method of making it wherein a two-piece mandrel wrapped with a silica tape saturated with a phenolic resin, the tape extending along the mandrel and covering the combustion chamber portion of the mandrel to the throat portion. The phenolic in the tape is cured and the end of the wrap is machined. The remainder of the mandrel is wrapped with a third silica tape. The resin in the third tape is cured and the assembly is machined. The entire assembly is then wrapped with a tow of graphite fibers wetted with an epoxy resin and, after the epoxy resin is cured, the graphite is machined to final dimensions.

  13. Orbital transfer rocket engine technology 7.5K-LB thrust rocket engine preliminary design

    NASA Technical Reports Server (NTRS)

    Harmon, T. J.; Roschak, E.

    1993-01-01

    A preliminary design of an advanced LOX/LH2 expander cycle rocket engine producing 7,500 lbf thrust for Orbital Transfer vehicle missions was completed. Engine system, component and turbomachinery analysis at both on design and off design conditions were completed. The preliminary design analysis results showed engine requirements and performance goals were met. Computer models are described and model outputs are presented. Engine system assembly layouts, component layouts and valve and control system analysis are presented. Major design technologies were identified and remaining issues and concerns were listed.

  14. Advanced oxygen-hydrocarbon rocket engine study

    NASA Technical Reports Server (NTRS)

    Obrien, C. J.; Ewen, R. L.

    1981-01-01

    This study identifies and evaluates promising LO2/HC rocket engine cycles, produces a consistent and reliable data base for vehicle optimization and design studies, demonstrates the significance of propulsion system improvements, and selects the critical technology areas necessary to realize an improved surface to orbit transportation system. Parametric LO2/HC engine data were generated over a range of thrust levels from 890 to 6672 kN (200K to 1.5M 1bF) and chamber pressures from 6890 to 34500 kN (1000 to 5000 psia). Engine coolants included RP-1, refined RP-1, LCH4, LC3H8, LO2, and LH2. LO2/RP-1 G.G. cycles were found to be not acceptable for advanced engines. The highest performing LO2/RP-1 staged combustion engine cycle utilizes LO2 as the coolant and incorporates an oxidizer rich preburner. The highest performing cycle for LO2/LCH4 and LO2/LC3H8 utilizes fuel cooling and incorporates both fuel and oxidizer rich preburners. LO2/HC engine cycles permitting the use of a third fluid LH2 coolant and an LH2 rich gas generator provide higher performance at significantly lower pump discharge pressures. The LO2/HC dual throat engine, because of its high altitude performance, delivers the highest payload for the vehicle configuration that was investigated.

  15. Devising rocket power for smaller engines

    SciTech Connect

    Burruss, R.

    1996-04-01

    Compact, high-power engines that burn fuel and oxygen could be made by winding copper tubing in a helix around boiler sections. With more than 1,000 horsepower per pound of engine weight, liquid-fueled rockets have the highest specific power of any engines designed for sustained operation. Yet those engines generally run for about only 1,000 seconds--nowhere near the sustained operation time for lower-power automotive and aircraft engines of more than 1,000 hours. In theory, at least, a fuel/oxygen rocket can be built that combines the best of both classes: high specific power (from perhaps two to 10 times that of a gas turbine) and a 1,000-hour service life. Such an engine would almost certainly be possible if the rocket`s exhaust gases could be simultaneously cooled and expanded by mixing water with the rocket`s exhaust and boiling it before it reaches the turbine. The technology itself is not new. variations of these rocket-turbine-type engines, for example, powered torpedoes during World War I. Some 30 years later, German V-2 rockets used fuel pumps, driven by the reaction of hydrogen peroxide with hydrocarbon fuels, to produce high-pressure steam that was directed against a turbine. Alternatively, fuel/oxygen combustion could produce steam to drive a piston engine. Either way, the challenge remains to construct a compact, long-service-life, high-specific-power boiler that burns fuel and oxygen. The new type of engine could be derived from recent research on electric vehicles (EVs).

  16. Electrodynamic actuators for rocket engine valves

    NASA Technical Reports Server (NTRS)

    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.

  17. Seismographic recording of large rocket engine operation

    NASA Technical Reports Server (NTRS)

    Dalins, I.; Mc Carty, V.

    1969-01-01

    Recording equipment for rocket engine vibration is adaptable to determining the structural strength of building materials. This seismographic system is portable and is capable of measuring displacements in the direction of three mutually perpendicular axes.

  18. Rocket Engine Innovations Advance Clean Energy

    NASA Technical Reports Server (NTRS)

    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.

  19. Performance of a shaft seal system for the LE-7 rocket engine oxidizer turbopump

    NASA Astrophysics Data System (ADS)

    Oike, Mamoru; Nosaka, Masataka; Kikuchi, Masataka; Watanabe, Yoshiaki

    An experimental study on a rotating-shaft seal system for a high-pressure liquid oxygen (LOX) turbopump has been conducted to develop the LE-7 engine for the Japanese H-II launch vehicle. The LOX turbopump rotating-shaft seal system, which prevents LOX (4.9 MPa) and the high-pressure turbine-drive gas (16.6 MPa, 970 K) from mixing during high-speed operations (18,000 to 20,000 rpm), consists of the following seals: an LOX seal comprising a floating-ring and a wear-ring, a turbine gas seal comprising two floating-rings, and a helium purge seal comprising two segmented circumferential seal-rings. This report describes experimental and observational results concerning the rotating-shaft seal system obtained in the LOX turbopump operations and the seal tests. Based on comparisons between the measurements and the analytical results, sealing characteristics of the seal system are discussed. Inspections of the sealing surfaces after the engine firing tests demonstrated that the LOX turbopump rotating-shaft seal system has sufficient durability for use in the LE-7 engine for the H-II launch vehicle.

  20. Rocket Engine Thrust Chamber Assembly

    NASA Technical Reports Server (NTRS)

    Cornelius, Charles S. (Inventor); Counts, Richard H. (Inventor); Myers, W. Neill (Inventor); Lackey, Jeffrey D. (Inventor); Peters, Warren (Inventor); Shadoan, Michael (Inventor); Sparks, David L. (Inventor); Lawrence, Timothy W. (Inventor)

    2001-01-01

    A thrust chamber assembly for liquid fueled rocket engines and the method of making it wherein a two-piece mandrel having the configuration of an assembly having a combustion chamber portion connected to a nozzle portion through a throat portion is wrapped with a silica tape saturated with a phenolic resin, the tape extending along the mandrel and covering the combustion chamber portion of the mandrel to the throat portion. The width of the tape is positioned at an angle of 30 to 50 deg. to the axis of the mandrel such that one edge of the tape contacts the mandrel while the other edge is spaced from the mandrel. The phenolic in the tape is cured and the end of the wrap is machined to provide a frusto-conical surface extending at an angle of 15 to 30 deg. with respect to the axis of the mandrel for starting a second wrap on the mandrel to cover the throat portion. The remainder of the mandrel is wrapped with a third silica tape having its width positioned at a angle of 5 to 20 deg. from the axis of the mandrel. The resin in the third tape is cured and the assembly is machined to provide a smooth outer surface. The entire assembly is then wrapped with a tow of graphite fibers wetted with an epoxy resin and, after the epoxy resin is cured, the graphite is machined to final dimensions.

  1. Plasma Diagnostics Development for Advanced Rocket Engines

    NASA Astrophysics Data System (ADS)

    Glover, Timothy; Kittrell, Carter; Chan, Anthony; Chang-Diaz, Franklin

    2000-10-01

    The VASIMR (Variable Specific Impulse Magnetoplasma Rocket) engine is a next-generation rocket engine under development at the Johnson Space Center's Advanced Space Propulsion Laboratory. With an exhaust velocity up to 50 times that of chemical rocket engines such as the Space Shuttle Main Engine, the VASIMR concept promises fast, efficient interplanetary flight. Rice University has participated in VASIMR research since 1996 and at present is developing two new diagnostic probes: a retarding potential analyzer to measure the velocity of ions in the rocket's exhaust, and a moveable optical probe to examine the spectrum of the rocket's helicon plasma source. In support of the probe development, a test facility is under construction at Rice, consisting of a small electric rocket engine firing into a 2-m vacuum chamber. This engine, the MPD (magnetoplasmadynamic) thruster, dates from the 1960's and provides a well-characterized source plasma for testing of the probes under development. We present details of the ion energy analyzer and the facility under construction at Rice.

  2. GRCop-84 Developed for Rocket Engines

    NASA Technical Reports Server (NTRS)

    Ellis, David L.; Yun, Hee Man; Lerch, Bradley A.; Keller, Dennis A.; Holmes, Richard

    2001-01-01

    GRCop-84 (Cu-8 at.% Cr-4 at.% Nb) was developed at the NASA Glenn Research Center for use in regeneratively cooled rocket engines. This alloy possesses an excellent combination of conductivity, thermal expansion, strength, creep resistance, ductility, and low-cycle fatigue (LCF) life. In comparison to the currently used alloy, NARloy-Z (Cu-3 wt % Ag-0.5 wt % Zr), GRCop-84's significantly better properties give it the potential for significant gains in engine performance and reliability. For GRCop-84 to be safely used in a rocket engine, it is critical to develop a detailed data base of the important thermophysical and mechanical properties. Work has focused on five major properties: thermal conductivity, thermal expansion, tensile strength, creep resistance, and LCF life. The analysis went beyond simply measuring the data and reporting averages. A detailed statistical analysis was conducted that allowed regression of the data over the entire temperature range tested and establishment of design minimums. The design values were expressed as simple mathematical formulas that are highly amenable to use in computer codes such as in finite element analysis and related computer modeling work.

  3. Parallelization of Rocket Engine System Software (Press)

    NASA Technical Reports Server (NTRS)

    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.

  4. Pulse Detonation Rocket Engine Research at NASA Marshall

    NASA Technical Reports Server (NTRS)

    Morris, Christopher I.

    2003-01-01

    Pulse detonation rocket engines (PDREs) offer potential performance improvements over conventional designs, but represent a challenging modeling task. A quasi 1-D, finite-rate chemistry CFD model for a PDRE is described and implemented. A parametric study of the effect of blowdown pressure ratio on the performance of an optimized, fixed PDRE nozzle configuration is reported. The results are compared to a steady-state rocket system using similar modeling assumptions.

  5. Quasi-One-Dimensional Modeling of Pulse Detonation Rocket Engines

    NASA Technical Reports Server (NTRS)

    Morris, Christopher I.

    2003-01-01

    Pulse detonation rocket engines (PDREs) offer potential performance improvements over conventional designs, but represent a challenging modeling task. A quasi 1-D, finite-rate chemistry CFD model for a PDRE & described and implemented. A parametric study of the effect of blowdown pressure ratio on the performance of an optimized, fixed PDRE nozzle configuration is reported. The results are compared to a steady-state rocket system using similar modeling assumptions.

  6. Handbook of recommended practices for the determination of liquid monopropellant rocket engine performance

    NASA Technical Reports Server (NTRS)

    Bjorklund, R. A.; Rogero, R. S.; Baerwald, R. K.

    1979-01-01

    The design, installation, and operation of systems to be used for directly measuring quantities of fundamental importance to the determination of monopropellant thruster performance is described. Areas covered include: (1) force and impulse measurement; (2) propellant mass usage and flow measurement; (3) pressure measurement; (4) temperature measurement; (5) exhaust gas composition measurement; and (6) data reduction and performance determination.

  7. Pattern classification approach to rocket engine diagnostics

    SciTech Connect

    Tulpule, S.

    1989-01-01

    This paper presents a systems level approach to integrate state-of-the-art rocket engine technology with advanced computational techniques to develop an integrated diagnostic system (IDS) for future rocket propulsion systems. The key feature of this IDS is the use of advanced diagnostic algorithms for failure detection as opposed to the current practice of redline-based failure detection methods. The paper presents a top-down analysis of rocket engine diagnostic requirements, rocket engine operation, applicable diagnostic algorithms, and algorithm design techniques, which serve as a basis for the IDS. The concepts of hierarchical, model-based information processing are described, together with the use uf signal processing, pattern recognition, and artificial intelligence techniques which are an integral part of this diagnostic system. 27 refs.

  8. A Rocket Engine Design Expert System

    NASA Technical Reports Server (NTRS)

    Davidian, Kenneth J.

    1989-01-01

    The overall structure and capabilities of an expert system designed to evaluate rocket engine performance are described. The expert system incorporates a JANNAF standard reference computer code to determine rocket engine performance and a state of the art finite element computer code to calculate the interactions between propellant injection, energy release in the combustion chamber, and regenerative cooling heat transfer. Rule-of-thumb heuristics were incorporated for the H2-O2 coaxial injector design, including a minimum gap size constraint on the total number of injector elements. One dimensional equilibrium chemistry was used in the energy release analysis of the combustion chamber. A 3-D conduction and/or 1-D advection analysis is used to predict heat transfer and coolant channel wall temperature distributions, in addition to coolant temperature and pressure drop. Inputting values to describe the geometry and state properties of the entire system is done directly from the computer keyboard. Graphical display of all output results from the computer code analyses is facilitated by menu selection of up to five dependent variables per plot.

  9. A rocket engine design expert system

    NASA Technical Reports Server (NTRS)

    Davidian, Kenneth J.

    1989-01-01

    The overall structure and capabilities of an expert system designed to evaluate rocket engine performance are described. The expert system incorporates a JANNAF standard reference computer code to determine rocket engine performance and a state-of-the-art finite element computer code to calculate the interactions between propellant injection, energy release in the combustion chamber, and regenerative cooling heat transfer. Rule-of-thumb heuristics were incorporated for the hydrogen-oxygen coaxial injector design, including a minimum gap size constraint on the total number of injector elements. One-dimensional equilibrium chemistry was employed in the energy release analysis of the combustion chamber and three-dimensional finite-difference analysis of the regenerative cooling channels was used to calculate the pressure drop along the channels and the coolant temperature as it exits the coolant circuit. Inputting values to describe the geometry and state properties of the entire system is done directly from the computer keyboard. Graphical display of all output results from the computer code analyses is facilitated by menu selection of up to five dependent variables per plot.

  10. Analysis of a Rocket Based Combined Cycle Engine during Rocket Only Operation

    NASA Technical Reports Server (NTRS)

    Smith, T. D.; Steffen, C. J., Jr.; Yungster, S.; Keller, D. J.

    1998-01-01

    The all rocket mode of operation is a critical factor in the overall performance of a rocket based combined cycle (RBCC) vehicle. However, outside of performing experiments or a full three dimensional analysis, there are no first order parametric models to estimate performance. As a result, an axisymmetric RBCC engine was used to analytically 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 statistical regression analysis was used to build parametric models. The main parameters investigated in this study were: rocket chamber pressure, rocket exit area ratio, percent of injected secondary flow, mixer-ejector inlet area, mixer-ejector area ratio, and mixer-ejector length-to-inject diameter ratio. A perfect gas computational fluid dynamics analysis was performed to obtain values of vacuum specific impulse. Statistical regression analysis was performed based on both full flow and gas generator engine cycles. Results were also found to be dependent upon the entire cycle assumptions. The statistical regression analysis determined that there were five significant linear effects, six interactions, and one second-order effect. Two parametric models were created to provide performance assessments of an RBCC engine in the all rocket mode of operation.

  11. Rocket engine hot-spot detector

    NASA Astrophysics Data System (ADS)

    Collamore, F. N.

    1985-04-01

    On high performance devices such as rocket engines it is desirable to know if local hot spots or areas of reduced cooling margin exist. The objective of this program is to design, fabricate and test an electronic hot spot detector capable of sensing local hot spot on the exterior circumference of a regeneratively cooled combustion chamber in order to avoid hardware damage. The electronic hot spot sensor consists of an array of 120 thermocouple elements which are bonded in a flexible belt of polyimide film. The design temperature range is from +30 F to +400 F continuously with an intermittent temperature of 500 F maximum. The thermocouple belt consists of 120 equally spaced copper-Constantan thermocouple junctions which is wrapped around the OMS liquid rocket engine combustion chamber, to monitor temperatures of individual cooling channels. Each thermocouple is located over a cooling channel near the injector end of the combustion chamber. The thermocouple array sensor is held in place by a spring loaded clamp band. Analyses show that in the event of a blocked cooling channel the surface temperature of the chamber over the blocked channel will rise from a normal operating temperature of approx. 300 F to approx. 600 F. The hot spot detector will respond quickly to this change with a response time constant less than 0.05 seconds. The hot spot sensor assembly is fabricated with a laminated construction of layers of Kapton film and an outer protective layer of fiberglass reinforced silicone rubber.

  12. A Versatile Rocket Engine Hot Gas Facility

    NASA Technical Reports Server (NTRS)

    Green, James M.

    1993-01-01

    The capabilities of a versatile rocket engine facility, located in the Rocket Laboratory at the NASA Lewis Research Center, are presented. The gaseous hydrogen/oxygen facility can be used for thermal shock and hot gas testing of materials and structures as well as rocket propulsion testing. Testing over a wide range of operating conditions in both fuel and oxygen rich regimes can be conducted, with cooled or uncooled test specimens. The size and location of the test cell provide the ability to conduct large amounts of testing in short time periods with rapid turnaround between programs.

  13. Orbit transfer rocket engine technology program

    NASA Technical Reports Server (NTRS)

    Gustafson, N. B.; Harmon, T. J.

    1993-01-01

    An advanced near term (1990's) space-based Orbit Transfer Vehicle Engine (OTVE) system was designed, and the technologies applicable to its construction, maintenance, and operations were developed under Tasks A through F of the Orbit Transfer Rocket Engine Technology Program. Task A was a reporting task. In Task B, promising OTV turbomachinery technologies were explored: two stage partial admission turbines, high velocity ratio diffusing crossovers, soft wear ring seals, advanced bearing concepts, and a rotordynamic analysis. In Task C, a ribbed combustor design was developed. Possible rib and channel geometries were chosen analytically. Rib candidates were hot air tested and laser velocimeter boundary layer analyses were conducted. A channel geometry was also chosen on the basis of laser velocimeter data. To verify the predicted heat enhancement effects, a ribbed calorimeter spool was hot fire tested. Under Task D, the optimum expander cycle engine thrust, performance and envelope were established for a set of OTV missions. Optimal nozzle contours and quick disconnects for modularity were developed. Failure Modes and Effects Analyses, maintenance and reliability studies and component study results were incorporated into the engine system. Parametric trades on engine thrust, mixture ratio, and area ratio were also generated. A control system and the health monitoring and maintenance operations necessary for a space-based engine were outlined in Task E. In addition, combustor wall thickness measuring devices and a fiberoptic shaft monitor were developed. These monitoring devices were incorporated into preflight engine readiness checkout procedures. In Task F, the Integrated Component Evaluator (I.C.E.) was used to demonstrate performance and operational characteristics of an advanced expander cycle engine system and its component technologies. Sub-system checkouts and a system blowdown were performed. Short transitions were then made into main combustor ignition and

  14. AJ26 rocket engine testing news briefing

    NASA Technical Reports Server (NTRS)

    2010-01-01

    Operators at NASA's John C. Stennis Space Center are completing modifications to the E-1 Test Stand to begin testing Aerojet AJ26 rocket engines in early summer of 2010. Modifications include construction of a 27-foot-deep flame deflector trench. The AJ26 rocket engines will be used to power Orbital Sciences Corp.'s Taurus II space vehicles to provide commercial cargo transportation missions to the International Space Station for NASA. Stennis has partnered with Orbital to test all engines for the transport missions.

  15. Iridium/Rhenium Parts For Rocket Engines

    NASA Technical Reports Server (NTRS)

    Schneider, Steven J.; Harding, John T.; Wooten, John R.

    1991-01-01

    Oxidation/corrosion of metals at high temperatures primary life-limiting mechanism of parts in rocket engines. Combination of metals greatly increases operating temperature and longevity of these parts. Consists of two transition-element metals - iridium and rhenium - that melt at extremely high temperatures. Maximum operating temperature increased to 2,200 degrees C from 1,400 degrees C. Increases operating lifetimes of small rocket engines by more than factor of 10. Possible to make hotter-operating, longer-lasting components for turbines and other heat engines.

  16. Acoustic/infrasonic rocket engine signatures

    NASA Astrophysics Data System (ADS)

    Tenney, Stephen M.; Noble, John M.; Whitaker, Rodney W.; ReVelle, Douglas O.

    2003-09-01

    Infrasonics offers the potential of long-range acoustic detection of explosions, missiles and even sounds created by manufacturing plants. The atmosphere attenuates acoustic energy above 20 Hz quite rapidly, but signals below 10 Hz can propagate to long ranges. Space shuttle launches have been detected infrasonically from over 1000 km away and the Concorde airliner from over 400 km. This technology is based on microphones designed to respond to frequencies from .1 to 300 Hz that can be operated outdoors for extended periods of time with out degrading their performance. The US Army Research Laboratory and Los Alamos National Laboratory have collected acoustic and infrasonic signatures of static engine testing of two missiles. Signatures were collected of a SCUD missile engine at Huntsville, AL and a Minuteman engine at Edwards AFB. The engines were fixed vertically in a test stand during the burn. We will show the typical time waveform signals of these static tests and spectrograms for each type. High resolution, 24-bit data were collected at 512 Hz and 16-bit acoustic data at 10 kHz. Edwards data were recorded at 250 Hz and 50 Hz using a Geotech Instruments 24 bit digitizer. Ranges from the test stand varied from 1 km to 5 km. Low level and upper level meteorological data was collected to provide full details of atmospheric propagation during the engine test. Infrasonic measurements were made with the Chaparral Physics Model 2 microphone with porous garden hose attached for wind noise suppression. A B&K microphone was used for high frequency acoustic measurements. Results show primarily a broadband signal with distinct initiation and completion points. There appear to be features present in the signals that would allow identification of missile type. At 5 km the acoustic/infrasonic signal was clearly present. Detection ranges for the types of missile signatures measured will be predicted based on atmospheric modeling. As part of an experiment conducted by ARL

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

    NASA Technical Reports Server (NTRS)

    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.

  18. Digital controller for high pressure rocket engine.

    NASA Technical Reports Server (NTRS)

    Thompson, Z.; Cummings, W. J.; Hall, D. M.

    1972-01-01

    Description of a general approach for the design of an adaptive digital control system for liquid bipropellant rocket engines. The technique employs linearized transfer functions derived from perturbations of an engine simulation. The linear models serve as a basis on which to develop candidate closed-loop control laws quickly and economically.

  19. Catalyst Development for Hydrogen Peroxide Rocket Engines

    NASA Technical Reports Server (NTRS)

    Morlan, P. W.; Wu, P.-K.; Ruttle, D. W.; Fuller, R. P.; Nejad, A. S.; Anderson, W. E.

    1999-01-01

    The development of various catalysts of hydrogen peroxide was conducted for the applications of liquid rocket engines. The catalyst development includes silver screen technology, solid catalyst technology, and homogeneous catalyst technology. The silver screen technology development was performed with 85% (by weight) hydrogen peroxide. The results of this investigation were used as the basis for the catalyst design of a pressure-fed liquid-fueled upper stage engine. Both silver-plated nickel 200 screens and pure silver screens were used as the active metal catalyst during the investigation, The data indicate that a high decomposition efficiency (greater than 90%) of 85% hydrogen peroxide can be achieved at a bed loading of 0.5 lbm/sq in/sec with both pure silver and silver plated screens. Samarium oxide coating, however, was found to retard the decomposition process and the catalyst bed was flooded at lower bed loading. A throughput of 200 lbm of hydrogen peroxide (1000 second run time) was tested to evaluate the catalyst aging issue and performance degradation was observed starting at approximately 400 seconds. Catalyst beds of 3.5 inch in diameter was fabricated using the same configuration for a 1,000-lbf rocket engine. High decomposition efficiency was obtained with a low pressure drop across the bed. Solid catalyst using precious metal was also developed for the decomposition of hydrogen peroxide from 85% to 98% by weight. Preliminary results show that the catalyst has a strong reactivity even after 15 minutes of peroxide decomposition. The development effort also includes the homogeneous catalyst technology. Various non-toxic catalysts were evaluated with 98% peroxide and hydrocarbon fuels. The results of open cup drop tests indicate an ignition delay around 11 ms.

  20. Radiative forcing caused by rocket engine emissions

    NASA Astrophysics Data System (ADS)

    Ross, Martin N.; Sheaffer, Patti M.

    2014-04-01

    Space transportation plays an important and growing role in Earth's economic system. Rockets uniquely emit gases and particles directly into the middle and upper atmosphere where exhaust from hundreds of launches accumulates, changing atmospheric radiation patterns. The instantaneous radiative forcing (RF) caused by major rocket engine emissions CO2, H2O, black carbon (BC), and Al2O3 (alumina) is estimated. Rocket CO2 and H2O emissions do not produce significant RF. BC and alumina emissions, under some scenarios, have the potential to produce significant RF. Absorption of solar flux by BC is likely the main RF source from rocket launches. In a new finding, alumina particles, previously thought to cool the Earth by scattering solar flux back to space, absorb outgoing terrestrial longwave radiation, resulting in net positive RF. With the caveat that BC and alumina microphysics are poorly constrained, we find that the present-day RF from rocket launches equals 16 ± 8 mW m-2. The relative contributions from BC, alumina, and H2O are 70%, 28%, and 2%. respectively. The pace of rocket launches is predicted to grow and space transport RF could become comparable to global aviation RF in coming decades. Improved understanding of rocket emission RF requires more sophisticated modeling and improved data describing particle microphysics.

  1. Design considerations in clustering nuclear rocket engines

    NASA Technical Reports Server (NTRS)

    Sager, Paul H.

    1992-01-01

    An initial investigation of the design considerations in clustering nuclear rocket engines for space transfer vehicles has been made. The clustering of both propulsion modules (which include start tanks) and nuclear rocket engines installed directly to a vehicle core tank appears to be feasible. Special provisions to shield opposite run tanks and the opposite side of a core tank - in the case of the boost pump concept - are required; the installation of a circumferential reactor side shield sector appears to provide an effective solution to this problem. While the time response to an engine-out event does not appear to be critical, the gimbal displacement required appears to be important. Since an installation of three engines offers a substantial reduction in gimbal requirements for engine-out and it may be possible to further enhance mission reliability with the greater number of engines, it is recommended that a cluster of four engines be considered.

  2. Amateur Gas-Propelled Rocket Engine Development and Advanced Rocket Design

    NASA Astrophysics Data System (ADS)

    Souverein, L. J.; Twigt, D. J.; Engelen, S.

    The paper describes the design and manufacturing of a gaseous propellant rocket engine. It is an undertaking of the authors, performed on project basis with fellow aerospace engineering students under auspices of DARE (Delft Aerospace Rocket Engineering). This paper describes the requirements, the engine development, and the design considerations and calculations as they were performed. Furthermore, the plans for engine tests and the parameters that will have to be measured during those tests are covered. The design process converged to a 1800 N thrust gaseous oxygen/methane (GOX/CH4) engine made of electrolytic copper. GOX/CH4 was selected based on its relatively high specific impulse, its availability and because of its potential as a green propellant. A test engine was produced with a specific impulse of 287 s and a propellant mass flow of 637 g/s. From a point of view of strength, the focus was mainly on robustness rather than light weight. The main aim now is to perform tests with the current engine, based on which the performance can be verified and vital information for future design efforts can be acquired. The ultimate goal is to have an operational rocket and to attempt an amateur altitude record.

  3. Fiber-Reinforced Superalloys For Rocket Engines

    NASA Technical Reports Server (NTRS)

    Lewis, Jack R.; Yuen, Jim L.; Petrasek, Donald W.; Stephens, Joseph R.

    1990-01-01

    Report discusses experimental studies of fiber-reinforced superalloy (FRS) composite materials for use in turbine blades in rocket engines. Intended to withstand extreme conditions of high temperature, thermal shock, atmospheres containing hydrogen, high cycle fatigue loading, and thermal fatigue, which tax capabilities of even most-advanced current blade material - directionally-solidified, hafnium-modified MAR M-246 {MAR M-246 (Hf) (DS)}. FRS composites attractive combination of properties for use in turbopump blades of advanced rocket engines at temperatures from 870 to 1,100 degrees C.

  4. Robust Strategy for Rocket Engine Health Monitoring

    NASA Technical Reports Server (NTRS)

    Santi, L. Michael

    2001-01-01

    Monitoring the health of rocket engine systems is essentially a two-phase process. The acquisition phase involves sensing physical conditions at selected locations, converting physical inputs to electrical signals, conditioning the signals as appropriate to establish scale or filter interference, and recording results in a form that is easy to interpret. The inference phase involves analysis of results from the acquisition phase, comparison of analysis results to established health measures, and assessment of health indications. A variety of analytical tools may be employed in the inference phase of health monitoring. These tools can be separated into three broad categories: statistical, rule based, and model based. Statistical methods can provide excellent comparative measures of engine operating health. They require well-characterized data from an ensemble of "typical" engines, or "golden" data from a specific test assumed to define the operating norm in order to establish reliable comparative measures. Statistical methods are generally suitable for real-time health monitoring because they do not deal with the physical complexities of engine operation. The utility of statistical methods in rocket engine health monitoring is hindered by practical limits on the quantity and quality of available data. This is due to the difficulty and high cost of data acquisition, the limited number of available test engines, and the problem of simulating flight conditions in ground test facilities. In addition, statistical methods incur a penalty for disregarding flow complexity and are therefore limited in their ability to define performance shift causality. Rule based methods infer the health state of the engine system based on comparison of individual measurements or combinations of measurements with defined health norms or rules. This does not mean that rule based methods are necessarily simple. Although binary yes-no health assessment can sometimes be established by

  5. Reliability Estimation Methods for Liquid Rocket Engines

    NASA Astrophysics Data System (ADS)

    Hirata, Kunio; Masuya, Goro; Kamijo, Kenjiro

    Reliability estimation using the dispersive, binominal distribution method has been traditionally used to certify the reliability of liquid rocket engines, but its estimation sometimes disagreed with the failure rates of flight engines. In order to take better results, the reliability growth model and the failure distribution method are applied to estimate the reliability of LE-7A engines, which have propelled the first stage of H-2A launch vehicles.

  6. Experimental study of rocket engine model with gaseous polyethylene fuel

    NASA Astrophysics Data System (ADS)

    Yemets, V. V.

    Experimental results for liquid rocket engine models with gaseous polyethylene fuel that is hard before its consumption are considered. The possibility of hard design element combustion in a liquid rocket engine is demonstrated.

  7. 12. Historic plot plan and drawings index for rocket engine ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    12. Historic plot plan and drawings index for rocket engine test facility, June 28, 1956. NASA GRC drawing number CE-101810. On file at NASA Glenn Research Center. - Rocket Engine Testing Facility, NASA Glenn Research Center, Cleveland, Cuyahoga County, OH

  8. 7. Historic aerial photo of rocket engine test facility complex, ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    7. Historic aerial photo of rocket engine test facility complex, June 1962. On file at NASA Plumbrook Research Center, Sandusky, Ohio. NASA GRC photo number C-60674. - Rocket Engine Testing Facility, NASA Glenn Research Center, Cleveland, Cuyahoga County, OH

  9. Development of nuclear rocket engine technology

    SciTech Connect

    Gunn, S.V.

    1989-01-01

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

  10. AXISYMMETRIC, THROTTLEABLE NON-GIMBALLED ROCKET ENGINE

    NASA Technical Reports Server (NTRS)

    Sackheim, Robert L. (Inventor); Hutt, John J. (Inventor); Anderson, William E. (Inventor); Dressler, Gordon A. (Inventor)

    2005-01-01

    A rocket engine assembly is provided for a vertically launched rocket vehicle. A rocket engine housing of the assembly includes two or more combustion chambers each including an outlet end defining a sonic throat area. A propellant supply for the combustion chambers includes a throttling injector, associated with each of the combustion chambers and located opposite to sonic throat area, which injects the propellant into the associated combustion chamber. A modulator, which may form part of the injector, and which is controlled by a controller, modulates the flow rate of the propellant to the combustion chambers so that the chambers provide a vectorable net thrust. An expansion nozzle or body located downstream of the throat area provides expansion of the combustion gases produced by the combustion chambers so as to increase the net thrust.

  11. Hybrid rocket engine, theoretical model and experiment

    NASA Astrophysics Data System (ADS)

    Chelaru, Teodor-Viorel; Mingireanu, Florin

    2011-06-01

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

  12. CFD Simulation of Liquid Rocket Engine Injectors

    NASA Technical Reports Server (NTRS)

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

    2001-01-01

    these investigators to be very valuable for code validation because combustion kinetics, turbulence models and atomization models based on low pressure experiments of hydrogen air combustion do not adequately verify analytical or CFD submodels which are necessary to simulate rocket engine combustion. We wish to emphasize that the simulations which we prepared for this meeting are meant to test the accuracy of the approximations used in our general purpose spray combustion models, rather than represent a definitive analysis of each of the experiments which were conducted. Our goal is to accurately predict local temperatures and mixture ratios in rocket engines; hence predicting individual experiments is used only for code validation. To replace the conventional JANNAF standard axisymmetric finite-rate (TDK) computer code 2 for performance prediction with CFD cases, such codes must posses two features. Firstly, they must be as easy to use and of comparable run times for conventional performance predictions. Secondly, they must provide more detailed predictions of the flowfields near the injector face. Specifically, they must accurately predict the convective mixing of injected liquid propellants in terms of the injector element configurations.

  13. Pulse Detonation Rocket Engine Research at NASA Marshall

    NASA Technical Reports Server (NTRS)

    Morris, Christopher I.

    2003-01-01

    This viewgraph representation provides an overview of research being conducted on Pulse Detonation Rocket Engines (PDRE) by the Propulsion Research Center (PRC) at the Marshall Space Flight Center. PDREs have a theoretical thermodynamic advantage over Steady-State Rocket Engines (SSREs) although unsteady blowdown processes complicate effective use of this advantage in practice; PRE is engaged in a fundamental study of PDRE gas dynamics to improve understanding of performance issues. Topics covered include: simplified PDRE cycle, comparison of PDRE and SSRE performance, numerical modeling of quasi 1-D rocket flows, time-accurate thrust calculations, finite-rate chemistry effects in nozzles, effect of F-R chemistry on specific impulse, effect of F-R chemistry on exit species mole fractions and PDRE performance optimization studies.

  14. Predicting performance of axial pump inducer of LOX booster turbo-pump of staged combustion cycle based rocket engine using CFD

    NASA Astrophysics Data System (ADS)

    Mishra, Arpit; Ghosh, Parthasarathi

    2015-12-01

    For low cost, high thrust, space missions with high specific impulse and high reliability, inert weight needs to be minimized and thereby increasing the delivered payload. Turbopump feed system for a liquid propellant rocket engine (LPRE) has the highest power to weight ratio. Turbopumps are primarily equipped with an axial flow inducer to achieve the high angular velocity and low suction pressure in combination with increased system reliability. Performance of the turbopump strongly depends on the performance of the inducer. Thus, for designing a LPRE turbopump, demands optimization of the inducer geometry based on the performance of different off-design operating regimes. In this paper, steady-state CFD analysis of the inducer of a liquid oxygen (LOX) axial pump used as a booster pump for an oxygen rich staged combustion cycle rocket engine has been presented using ANSYS® CFX. Attempts have been made to obtain the performance characteristic curves for the LOX pump inducer. The formalism has been used to predict the performance of the inducer for the throttling range varying from 80% to 113% of nominal thrust and for the different rotational velocities from 4500 to 7500 rpm. The results have been analysed to determine the region of cavitation inception for different inlet pressure.

  15. Combustion Characteristics of a Swirling LOX Type Hybrid Rocket Engine

    NASA Astrophysics Data System (ADS)

    Kitagawa, Koki; Yuasa, Saburo

    We have proposed a swirling oxidizer type hybrid rocket engine. In this paper, liquid oxygen (LOX) was used as oxidizer. Combustion tests of a hybrid rocket engine with a swirling LOX flow were conducted by changing the swirl strength. Ignition was rapid and reliable, and combustion of PMMA with swirling LOX was stable. Fuel regression rates, C* efficiency and specific impulse of the hybrid rocket engine with swirling LOX flow were smaller than those with swirling gaseous oxygen (GOX). This low performance may be restraint of atomization and vaporization of LOX by formation of a liquid layer on the PMMA fuel and a decline of angular momentum of the swirling LOX during vaporization. Combustion oscillation occurred when the ratios of differential pressure between injector pressure and chamber pressure to chamber pressure were small. This combustion oscillation was confirmed to be a “Chugging” mode due to combustion time lag of LOX.

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

    NASA Technical Reports Server (NTRS)

    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.

  17. Remote rocket engine leak detection techniques

    NASA Astrophysics Data System (ADS)

    Maram, J. M.

    1993-06-01

    Optical imaging techniques have been successfully used for detection of leaks in rocket engine components using introduced gases such as sulfur hexafluoride and nitrous oxide. The approach used and past applications of the technology are described. The potential for direct detection of launch system propellant leaks is discussed.

  18. Advanced oxygen-hydrocarbon rocket engine study

    NASA Technical Reports Server (NTRS)

    Obrien, C. J.

    1980-01-01

    Preliminary identification and evaluation of promising liquid oxygen/ hydrocarbon (LO2/HC) rocket engine cycles is reported. A consistent and reliable data base for vehicle optimization and design studies, to demonstrate the significance of propulsion system improvements, and to select the critical technology areas necessary to realize such advances is presented.

  19. Nondestructive testing of brazed rocket engine components

    NASA Technical Reports Server (NTRS)

    Adams, C. J.; Hagemaier, D. J.; Meyer, J. A.

    1968-01-01

    Report details study made of nondestructive radiographic, ultrasonic, thermographic, and leak test methods used to inspect and evaluate the quality of the various brazed joints in liquid-propellant rocket engine components and assemblies. Descriptions of some of the unique equipment and methods developed are included.

  20. Turbopump systems for liquid rocket engines

    NASA Technical Reports Server (NTRS)

    1974-01-01

    The turbopump system, from preliminary design through rocket engine testing is examined. Selection of proper system type for each application and integration of the components into a working system are dealt with. Details are also given on the design of various components including inducers, pumps, turbines, gears, and bearings.

  1. Heat transfer in a nuclear rocket engine

    SciTech Connect

    Konyukhov, G.V.; Petrov, A.I.

    1995-02-01

    Special features of heat transfer in the reactor of a nuclear rocket engine (NRE) are dealt with. It is shown that the design of the cooling system of the NRE reactor is governed by its stability to small deviations of the parameters from the corresponding calculated values and the possibility of compensating for effects due to nonuniformities and distrubances of various types and scales.

  2. Use of Soft Computing Technologies For Rocket Engine Control

    NASA Technical Reports Server (NTRS)

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

    2003-01-01

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

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

    NASA Technical Reports Server (NTRS)

    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.

  4. Mean Line Pump Flow Model in Rocket Engine System Simulation

    NASA Technical Reports Server (NTRS)

    Veres, Joseph P.; Lavelle, Thomas M.

    2000-01-01

    A mean line pump flow modeling method has been developed to provide a fast capability for modeling turbopumps of rocket engines. Based on this method, a mean line pump flow code PUMPA has been written that can predict the performance of pumps at off-design operating conditions, given the loss of the diffusion system at the design point. The pump code can model axial flow inducers, mixed-flow and centrifugal pumps. The code can model multistage pumps in series. The code features rapid input setup and computer run time, and is an effective analysis and conceptual design tool. The map generation capability of the code provides the map information needed for interfacing with a rocket engine system modeling code. The off-design and multistage modeling capabilities of the code permit parametric design space exploration of candidate pump configurations and provide pump performance data for engine system evaluation. The PUMPA code has been integrated with the Numerical Propulsion System Simulation (NPSS) code and an expander rocket engine system has been simulated. The mean line pump flow code runs as an integral part of the NPSS rocket engine system simulation and provides key pump performance information directly to the system model at all operating conditions.

  5. AJ26 Rocket Engine Test

    NASA Video Gallery

    Engineers at NASA’s John C. Stennis Space Center conducts the second in a series of verification tests on an Aerojet AJ26 engine that will power the first stage of the Orbital Sciences Corporatio...

  6. Analytical concepts for health management systems of liquid rocket engines

    NASA Technical Reports Server (NTRS)

    Williams, Richard; Tulpule, Sharayu; Hawman, Michael

    1990-01-01

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

  7. Liquid rocket engine centrifugal flow turbopumps. [design criteria

    NASA Technical Reports Server (NTRS)

    1973-01-01

    Design criteria and recommended practices are discussed for the following configurations selected from the design sequence of a liquid rocket engine centrifugal flow turbopump: (1) pump performance including speed, efficiency, and flow range; (2) impeller; (3) housing; and (4) thrust balance system. Hydrodynamic, structural, and mechanical problems are addressed for the achievement of required pump performance within the constraints imposed by the engine/turbopump system. Materials and fabrication specifications are also discussed.

  8. Multivariable optimization of liquid rocket engines using particle swarm algorithms

    NASA Astrophysics Data System (ADS)

    Jones, Daniel Ray

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

  9. Pretest uncertainty analysis for chemical rocket engine tests

    NASA Technical Reports Server (NTRS)

    Davidian, Kenneth J.

    1987-01-01

    A parametric pretest uncertainty analysis has been performed for a chemical rocket engine test at a unique 1000:1 area ratio altitude test facility. Results from the parametric study provide the error limits required in order to maintain a maximum uncertainty of 1 percent on specific impulse. Equations used in the uncertainty analysis are presented.

  10. AJ26 rocket engine testing news briefing

    NASA Technical Reports Server (NTRS)

    2010-01-01

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

  11. Design and Testing of a Liquid Nitrous Oxide and Ethanol Fueled Rocket Engine

    SciTech Connect

    Youngblood, Stewart

    2015-08-01

    A small-scale, bi-propellant, liquid fueled rocket engine and supporting test infrastructure were designed and constructed at the Energetic Materials Research and Testing Center (EMRTC). This facility was used to evaluate liquid nitrous oxide and ethanol as potential rocket propellants. Thrust and pressure measurements along with high-speed digital imaging of the rocket exhaust plume were made. This experimental data was used for validation of a computational model developed of the rocket engine tested. The developed computational model was utilized to analyze rocket engine performance across a range of operating pressures, fuel-oxidizer mixture ratios, and outlet nozzle configurations. A comparative study of the modeling of a liquid rocket engine was performed using NASA CEA and Cantera, an opensource equilibrium code capable of being interfaced with MATLAB. One goal of this modeling was to demonstrate the ability of Cantera to accurately model the basic chemical equilibrium, thermodynamics, and transport properties for varied fuel and oxidizer operating conditions. Once validated for basic equilibrium, an expanded MATLAB code, referencing Cantera, was advanced beyond CEAs capabilities to predict rocket engine performance as a function of supplied propellant flow rate and rocket engine nozzle dimensions. Cantera was found to comparable favorably to CEA for making equilibrium calculations, supporting its use as an alternative to CEA. The developed rocket engine performs as predicted, demonstrating the developedMATLAB rocket engine model was successful in predicting real world rocket engine performance. Finally, nitrous oxide and ethanol were shown to perform well as rocket propellants, with specific impulses experimentally recorded in the range of 250 to 260 seconds.

  12. Combustion Processes in Hybrid Rocket Engines

    NASA Technical Reports Server (NTRS)

    Venkateswaran,S.; Merkle, C. L.

    1996-01-01

    In recent years, there has been a resurgence of interest in the development of hybrid rocket engines for advanced launch vehicle applications. Hybrid propulsion systems use a solid fuel such as hydroxyl-terminated polybutadiene (HTPB) along with a gaseous/liquid oxidizer. The performance of hybrid combustors depends on the convective and radiative heat fluxes to the fuel surface, the rate of pyrolysis in the solid phase, and the turbulent combustion processes in the gaseous phases. These processes in combination specify the regression rates of the fuel surface and thereby the utilization efficiency of the fuel. In this paper, we employ computational fluid dynamics (CFD) techniques in order to gain a quantitative understanding of the physical trends in hybrid rocket combustors. The computational modeling is tailored to ongoing experiments at Penn State that employ a two dimensional slab burner configuration. The coordinated computational/experimental effort enables model validation while providing an understanding of the experimental observations. Computations to date have included the full length geometry with and with the aft nozzle section as well as shorter length domains for extensive parametric characterization. HTPB is sed as the fuel with 1,3 butadiene being taken as the gaseous product of the pyrolysis. Pure gaseous oxygen is taken as the oxidizer. The fuel regression rate is specified using an Arrhenius rate reaction, which the fuel surface temperature is given by an energy balance involving gas-phase convection and radiation as well as thermal conduction in the solid-phase. For the gas-phase combustion, a two step global reaction is used. The standard kappa - epsilon model is used for turbulence closure. Radiation is presently treated using a simple diffusion approximation which is valid for large optical path lengths, representative of radiation from soot particles. Computational results are obtained to determine the trends in the fuel burning or

  13. The beginnings. [Of Nuclear Engine for Rocket Vehicles Application

    SciTech Connect

    Bohl, R.J.; Kirk, W.L.; Holman, R.R.; Westinghouse Electric Corp., Pittsburgh, PA )

    1989-06-01

    The development of the nuclear rocket engine called NERVA (Nuclear Engine for Rocket Vehicle Application) is described. The choice of fuel element, required rocket parameters, NERVA project objectives, division of responsibilities among different organizations, and NERVA design configuration are reviewed. Progress that has been made in the development of NERVA is addressed.

  14. Air breathing engine/rocket trajectory optimization

    NASA Technical Reports Server (NTRS)

    Smith, V. K., III

    1979-01-01

    This research has focused on improving the mathematical models of the air-breathing propulsion systems, which can be mated with the rocket engine model and incorporated in trajectory optimization codes. Improved engine simulations provided accurate representation of the complex cycles proposed for advanced launch vehicles, thereby increasing the confidence in propellant use and payload calculations. The versatile QNEP (Quick Navy Engine Program) was modified to allow treatment of advanced turboaccelerator cycles using hydrogen or hydrocarbon fuels and operating in the vehicle flow field.

  15. Combustion dynamics in liquid rocket engines

    NASA Technical Reports Server (NTRS)

    Mclain, W. H.

    1971-01-01

    A chemical analysis of the emission and absorption spectra in the combustion chamber of a nitrogen tetroxide/aerozine-50 rocket engine was conducted. Measurements were made under conditions of preignition, ignition, and post combustion operating periods. The cause of severe ignition overpressures sporadically observed during the vacuum startup of the Apollo reaction control system engine was investigated. The extent to which residual propellants or condensed intermediate reaction products remain after the engine has been operated in a pulse mode duty cycle was determined.

  16. Linear quadratic servo control of a reusable rocket engine

    NASA Technical Reports Server (NTRS)

    Musgrave, Jeffrey L.

    1991-01-01

    A design method for a servo compensator is developed in the frequency domain using singular values. The method is applied to a reusable rocket engine. An intelligent control system for reusable rocket engines was proposed which includes a diagnostic system, a control system, and an intelligent coordinator which determines engine control strategies based on the identified failure modes. The method provides a means of generating various linear multivariable controllers capable of meeting performance and robustness specifications and accommodating failure modes identified by the diagnostic system. Command following with set point control is necessary for engine operation. A Kalman filter reconstructs the state while loop transfer recovery recovers the required degree of robustness while maintaining satisfactory rejection of sensor noise from the command error. The approach is applied to the design of a controller for a rocket engine satisfying performance constraints in the frequency domain. Simulation results demonstrate the performance of the linear design on a nonlinear engine model over all power levels during mainstage operation.

  17. Liquid-hydrogen rocket engine development at Aerojet, 1944 - 1950

    NASA Technical Reports Server (NTRS)

    Osborn, G. H.; Gordon, R.; Coplen, H. L.; James, G. S.

    1977-01-01

    This program demonstrated the feasibility of virtually all the components in present-day, high-energy, liquid-rocket engines. Transpiration and film-cooled thrust chambers were successfully operated. The first liquid-hydrogen tests of the coaxial injector was conducted and the first pump to successfully produce high pressures in pumping liquid hydrogen was tested. A 1,000-lb-thrust gaseous propellant and a 3,000-lb-thrust liquid-propellant thrust chamber were operated satisfactorily. Also, the first tests were conducted to evaluate the effects of jet overexpansion and separation on performance of rocket thrust chambers with hydrogen-oxygen propellants.

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

    NASA Technical Reports Server (NTRS)

    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.

  19. Ground test facility for SEI nuclear rocket engines

    SciTech Connect

    Harmon, C.D.; Ottinger, C.A.; Sanchez, L.C.; Shipers, L.R.

    1992-08-01

    Nuclear Thermal Propulsion (NTP) has been identified as a critical technology in support of the NASA Space Exploration Initiative (SEI). In order to safely develop a reliable, reusable, long-lived flight engine, facilities are required that will support ground tests to qualify the nuclear rocket engine design. Initial nuclear fuel element testing will need to be performed in a facility that supports a realistic thermal and neutronic environment in which the fuel elements will operate at a fraction of the power of a flight weight reactor/engine. Ground testing of nuclear rocket engines is not new. New restrictions mandated by the National Environmental Protection Act of 1970, however, now require major changes to be made in the manner in which reactor engines are now tested. These new restrictions now preclude the types of nuclear rocket engine tests that were performed in the past from being done today. A major attribute of a safely operating ground test facility is its ability to prevent fission products from being released in appreciable amounts to the environment. Details of the intricacies and complications involved with the design of a fuel element ground test facility are presented in this report with a strong emphasis on safety and economy.

  20. 1. ROCKET ENGINE TEST STAND, LOCATED IN THE NORTHEAST ¼ ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    1. ROCKET ENGINE TEST STAND, LOCATED IN THE NORTHEAST ¼ OF THE X-15 ENGINE TEST COMPLEX. Looking northeast. - Edwards Air Force Base, X-15 Engine Test Complex, Rocket Engine & Complete X-15 Vehicle Test Stands, Rogers Dry Lake, east of runway between North Base & South Base, Boron, Kern County, CA

  1. Quantifying Instability Sources in Liquid Rocket Engines

    NASA Technical Reports Server (NTRS)

    Farmer, Richard C.; Cheng, Gary C.

    2000-01-01

    Computational fluid dynamics methodology to predict the effects of combusting flows on acoustic pressure oscillations in liquid rocket engines (LREs) is under development. 'Me intent of the investigation is to develop the causal physics of combustion driven acoustic resonances in LREs. The crux of the analysis is the accurate simulation of pressure/density/sound speed in a combustor which when used by the FDNS-RFV CFD code will produce realistic flow phenomena. An analysis of a gas generator considered for the Fastrac engine will be used as a test validation case.

  2. Key technology for reusable rocket engine turbopump

    NASA Astrophysics Data System (ADS)

    Okayasu, A.; Ohta, T.; Kamijyo, A.; Yamada, H.

    2002-03-01

    Recently, there has been an increased need for evolved space transportation and the research of reusable rocket which enable low cost and high reliability and is generating a lot of interest all over the world. In the USA, the development of reusable launch vehicle "Venture Star" which will be used instead of space shuttle is planned and its half scale model "X-33" was developed for the first flight in 1999. In Japan, there has been agreement on the main points to develop the rocket type RLV based on the technology of H-IIA, HOPE-X before developing space plane type RLV. The planned reusable rocket engine was LOX/LH2 as propellant, has 100- 200 ton thrust and has a throttling capability. In addition, long life and high reliability are required for the engine system including LOX/LH2 turbopump. The paper introduces some key technologies for the reusable turbopump which IHI is promoting for research and development with NAL and Tohoku University.

  3. Turbo Pump Fed Micro-Rocket Engine

    NASA Astrophysics Data System (ADS)

    Miotti, P.; Tajmar, M.; Seco, F.; Guraya, C.; Perennes, F.; Soldati, A.; Lang, M.

    2004-10-01

    Micro-satellites (from 10kg up to 100kg) have mass, volume, and electrical power constraints due to their low dimensions. These limitations lead to the lack in currently available active orbit control systems in micro-satellites. Therefore, a micro-propulsion system with a high thrust to mass ratio is required to increase the potential functionality of small satellites. Mechatronic is presently working on a liquid bipropellant micro-rocket engine under contract with ESA (Contract No.16914/NL/Sfe - Micro-turbo-machinery Based Bipropellant System Using MNT). The advances in Mechatronic's project are to realise a micro-rocket engine with propellants pressurised by micro-pumps. The energy for driving the pumps would be extracted from a micro-turbine. Cooling channels around the nozzle would be also used in order to maintain the wall material below its maximum operating temperature. A mass budget comparison with more traditional pressure-fed micro-rockets shows a real benefit from this system in terms of mass reduction. In the paper, an overview of the project status in Mechatronic is presented.

  4. Software for Collaborative Engineering of Launch Rockets

    NASA Technical Reports Server (NTRS)

    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.

  5. Efficiency of the rocket engines with a supersonic afterburner

    NASA Astrophysics Data System (ADS)

    Sergienko, A. A.

    1992-08-01

    The paper is concerned with the problem of regenerative cooling of the liquid-propellant rocket engine combustion chamber at high pressures of the working fluid. It is shown that high combustion product pressures can be achieved in the liquid-propellant rocket engine with a supersonic afterburner than in a liquid-propellant rocket engine with a conventional subsonic combustion chamber for the same allowable heat flux density. However, the liquid-propellant rocket engine with a supersonic afterburner becomes more economical than the conventional engine only at generator gas temperatures of 1700 K and higher.

  6. Low-thrust chemical rocket engine study

    NASA Technical Reports Server (NTRS)

    Mellish, J. A.

    1980-01-01

    Parametric data and preliminary designs on liquid rocket engines for low thrust cargo orbit-transfer-vehicles are described and those items where technology is required to enhance the designs are identified. The results of film cooling studies to establish the upper chamber pressure limit are given. The study showed that regen cooling with RP-1 was not feasible over the entire thrust and chamber pressure ranges. The thermal data showed that the RP-1 bulk temperature exceeded the study coking temperature limit of 1010 R. Based upon the results presented, O2/H2 and O2/CH4 regen engine systems and O2/H2 film cooled engines were selected for further study in the system analysis. Six engine design concepts are examined.

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

    NASA Technical Reports Server (NTRS)

    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.

  8. An RL10A-3-3A rocket engine model using the rocket engine transient simulator (ROCETS) software

    NASA Technical Reports Server (NTRS)

    Binder, Michael

    1993-01-01

    Steady-state and transient computer models of the RL10A-3-3A rocket engine have been created using the Rocket Engine Transient Simulation (ROCETS) code. These models were created for several purposes. The RL10 engine is a critical component of past, present, and future space missions; the model will give NASA an in-house capability to simulate the performance of the engine under various operating conditions and mission profiles. The RL10 simulation activity is also an opportunity to further validate the ROCETS program. The ROCETS code is an important tool for modeling rocket engine systems at NASA Lewis. ROCETS provides a modular and general framework for simulating the steady-state and transient behavior of any desired propulsion system. Although the ROCETS code is being used in a number of different analysis and design projects within NASA, it has not been extensively validated for any system using actual test data. The RL10A-3-3A has a ten year history of test and flight applications; it should provide sufficient data to validate the ROCETS program capability. The ROCETS models of the RL10 system were created using design information provided by Pratt & Whitney, the engine manufacturer. These models are in the process of being validated using test-stand and flight data. This paper includes a brief description of the models and comparison of preliminary simulation output against flight and test-stand data.

  9. Design issues for lunar in situ aluminum/oxygen propellant rocket engines

    NASA Technical Reports Server (NTRS)

    Meyer, Michael L.

    1992-01-01

    Design issues for lunar ascent and lunar descent rocket engines fueled by aluminum/oxygen propellant produced in situ at the lunar surface were evaluated. Key issues are discussed which impact the design of these rockets: aluminum combustion, throat erosion, and thrust chamber cooling. Four engine concepts are presented, and the impact of combustion performance, throat erosion and thrust chamber cooling on overall engine design are discussed. The advantages and disadvantages of each engine concept are presented.

  10. Performance potential of gas-core and fusion rockets - A mission applications survey.

    NASA Technical Reports Server (NTRS)

    Fishbach, L. H.; Willis, E. A., Jr.

    1971-01-01

    This paper reports an evaluation of the performance potential of five nuclear rocket engines for four mission classes. These engines are: the regeneratively cooled gas-core nuclear rocket; the light bulb gas-core nuclear rocket; the space-radiator cooled gas-core nuclear rocket; the fusion rocket; and an advanced solid-core nuclear rocket which is included for comparison. The missions considered are: earth-to-orbit launch; near-earth space missions; close interplanetary missions; and distant interplanetary missions. For each of these missions, the capabilities of each rocket engine type are compared in terms of payload ratio for the earth launch mission or by the initial vehicle mass in earth orbit for space missions (a measure of initial cost). Other factors which might determine the engine choice are discussed. It is shown that a 60 day manned round trip to Mars is conceivable.-

  11. 2. ROCKET ENGINE TEST STAND, SHOWING TANK (BUILDING 1929) AND ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    2. ROCKET ENGINE TEST STAND, SHOWING TANK (BUILDING 1929) AND GARAGE (BUILDING 1930) AT LEFT REAR. Looking to west. - Edwards Air Force Base, X-15 Engine Test Complex, Rocket Engine & Complete X-15 Vehicle Test Stands, Rogers Dry Lake, east of runway between North Base & South Base, Boron, Kern County, CA

  12. Very high thrust-to-weight rocket engines

    NASA Astrophysics Data System (ADS)

    Glass, James F.; Goracke, B. David; Levack, Daniel J. H.

    1998-01-01

    High delta-V earth-to-orbit missions have put a premium on high performance booster rocket engines. While significant improvements to specific impulse are unlikely, high thrust-to-weight design provides a promising avenue for improving mission and vehicle capabilities and margins. Several approaches can contribute to achieving such engine designs, including proper design optimization, simplification, geometry, propellant selection, and the application of advanced materials. Incorporation of the first four approaches can yield factors of about two improvements in current liquid engine designs. The utilization of emerging material capabilities could yield another factor of two improvement with the possibility of even larger gains with far-term materials and designs.

  13. Reusable rocket engine turbopump health management system

    NASA Astrophysics Data System (ADS)

    Surko, Pamela

    1994-10-01

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

  14. Reusable Rocket Engine Turbopump Health Management System

    NASA Technical Reports Server (NTRS)

    Surko, Pamela

    1994-01-01

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

  15. Prume Heating Analysis of Clustered Rocket Engines

    NASA Astrophysics Data System (ADS)

    Maemura, Takashi; Igarashi, Iwao

    The H-IIB launch vehicle is an upgraded version of the current H-IIA launch capacity, which has two liquid rocket engines (LE-7A) in the first-stage, instead of one for the H-IIA. It has four SRB-As attached to the body, while the standard version of H-IIA had two SRB-As. One of the major design issue of H-IIB launch vehicle is increased prume heating due to clustering two LE-7A engines and four SRB-As, especially the interaction of engine prumes at high altitude. This paper describes the prume analysys method of H-IIB launch vehicle which is based on the flight proven method of the current H-IIA launch.

  16. Rocket engine control and monitoring expert system

    NASA Technical Reports Server (NTRS)

    Ali, Moonis; Crawford, Roger

    1988-01-01

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

  17. Rocket engine diagnostics using qualitative modeling techniques

    NASA Technical Reports Server (NTRS)

    Binder, Michael; Maul, William; Meyer, Claudia; Sovie, Amy

    1992-01-01

    Researchers at NASA Lewis Research Center are presently developing qualitative modeling techniques for automated rocket engine diagnostics. A qualitative model of a turbopump interpropellant seal system was created. The qualitative model describes the effects of seal failures on the system steady state behavior. This model is able to diagnose the failure of particular seals in the system based on anomalous temperature and pressure values. The anomalous values input to the qualitative model are generated using numerical simulations. Diagnostic test cases include both single and multiple seal failures.

  18. Rocket engine diagnostics using qualitative modeling techniques

    NASA Technical Reports Server (NTRS)

    Binder, Michael; Maul, William; Meyer, Claudia; Sovie, Amy

    1992-01-01

    Researchers at NASA Lewis Research Center are presently developing qualitative modeling techniques for automated rocket engine diagnostics. A qualitative model of a turbopump interpropellant seal system has been created. The qualitative model describes the effects of seal failures on the system steady-state behavior. This model is able to diagnose the failure of particular seals in the system based on anomalous temperature and pressure values. The anomalous values input to the qualitative model are generated using numerical simulations. Diagnostic test cases include both single and multiple seal failures.

  19. Life extending control for rocket engines

    NASA Technical Reports Server (NTRS)

    Lorenzo, C. F.; Saus, J. R.; Ray, A.; Carpino, M.; Wu, M.-K.

    1992-01-01

    The concept of life extending control is defined. A brief discussion of current fatigue life prediction methods is given and the need for an alternative life prediction model based on a continuous functional relationship is established. Two approaches to life extending control are considered: (1) the implicit approach which uses cyclic fatigue life prediction as a basis for control design; and (2) the continuous life prediction approach which requires a continuous damage law. Progress on an initial formulation of a continuous (in time) fatigue model is presented. Finally, nonlinear programming is used to develop initial results for life extension for a simplified rocket engine (model).

  20. Deimos Methane-Oxygen Rocket Engine Test Results

    NASA Astrophysics Data System (ADS)

    Engelen, S.; Souverein, L. J.; Twigt, D. J.

    This paper presents the results of the first DEIMOS Liquid Methane/Oxygen rocket engine test campaign. DEIMOS is an acronym for `Delft Experimental Methane Oxygen propulsion System'. It is a project performed by students under the auspices of DARE (Delft Aerospace Rocket Engineering). The engine provides a theoretical design thrust of 1800 N and specific impulse of 287 s at a chamber pressure of 40 bar with a total mass flow of 637 g/s. It has links to sustainable development, as the propellants used are one of the most promising so-called `green propellants'-combinations, currently under scrutiny by the industry, and the engine is designed to be reusable. This paper reports results from the provisional tests, which had the aim of verifying the engine's ability to fire, and confirming some of the design assumptions to give confidence for further engine designs. Measurements before and after the tests are used to determine first estimates on feed pressures, propellant mass flows and achieved thrust. These results were rather disappointing from a performance point of view, with an average thrust of a mere 3.8% of the design thrust, but nonetheless were very helpful. The reliability of ignition and stability of combustion are discussed as well. An initial assessment as to the reusability, the flexibility and the adaptability of the engine was made. The data provides insight into (methane/oxygen) engine designs, leading to new ideas for a subsequent design. The ultimate goal of this project is to have an operational rocket and to attempt to set an amateur altitude record.

  1. Multiobjective Optimization of Rocket Engine Pumps Using Evolutionary Algorithm

    NASA Technical Reports Server (NTRS)

    Oyama, Akira; Liou, Meng-Sing

    2001-01-01

    A design optimization method for turbopumps of cryogenic rocket engines has been developed. Multiobjective Evolutionary Algorithm (MOEA) is used for multiobjective pump design optimizations. Performances of design candidates are evaluated by using the meanline pump flow modeling method based on the Euler turbine equation coupled with empirical correlations for rotor efficiency. To demonstrate the feasibility of the present approach, a single stage centrifugal pump design and multistage pump design optimizations are presented. In both cases, the present method obtains very reasonable Pareto-optimal solutions that include some designs outperforming the original design in total head while reducing input power by one percent. Detailed observation of the design results also reveals some important design criteria for turbopumps in cryogenic rocket engines. These results demonstrate the feasibility of the EA-based design optimization method in this field.

  2. Engine protection system for recoverable rocket booster

    NASA Technical Reports Server (NTRS)

    Shelby, Jr., Jerry A. (Inventor)

    1994-01-01

    A rocket engine protection system for a recoverable rocket booster which is arranged to land in a salt water body in substantially a nose down attitude. The system includes an inflatable bag which is stowed on a portion of a flat annular rim of the aft skirt of the booster. The bag is hinged at opposing sides and is provided with springs that urge the bag open. The bag is latched in a stowed position during launch and prior to landing for recovery is unlatched to permit the bag to be urged open and into sealing engagement with the rim. A source of pressurized gas further inflates the bag and urges it into sealing engagement with the rim of the skirt where it is locked into position. The gas provides a positive pressure upon the interior of the bag to preclude entry of salt water into the skirt and into contact with the engine. A flotation arrangement may assist in precluding the skirt of the booster from becoming submerged.

  3. Liquid fuel injection elements for rocket engines

    NASA Technical Reports Server (NTRS)

    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.

  4. Advanced oxygen-hydrocarbon rocket engine study

    NASA Technical Reports Server (NTRS)

    Obrien, C. J.; Salkeld, R.

    1980-01-01

    The advantages and disadvantages, system performance and operating limits, engine parametric data, and technology requirements for candidate high pressure LO2/Hydrocarbon engine systems are summarized. These summaries of parametric analysis and design provide a consistent engine system data base. Power balance data were generated for the eleven engine cycles. Engine cycle rating parameters were established and the desired condition and the effect of the parameter on the engine and/or vehicle are described.

  5. Modular Rocket Engine Control Software (MRECS)

    NASA Technical Reports Server (NTRS)

    Tarrant, C.; Crook, J.

    1998-01-01

    The Modular Rocket Engine Control Software (MRECS) Program is a technology demonstration effort designed to advance the state-of-the-art in launch vehicle propulsion systems. Its emphasis is on developing and demonstrating a modular software architecture for advanced engine control systems that will result in lower software maintenance (operations) costs. It effectively accommodates software requirement changes that occur due to hardware technology upgrades and engine development testing. Ground rules directed by MSFC were to optimize modularity and implement the software in the Ada programming language. MRECS system software and the software development environment utilize Commercial-Off-the-Shelf (COTS) products. This paper presents the objectives, benefits, and status of the program. The software architecture, design, and development environment are described. MRECS tasks are defined and timing relationships given. Major accomplishments are listed. MRECS offers benefits to a wide variety of advanced technology programs in the areas of modular software architecture, reuse software, and reduced software reverification time related to software changes. MRECS was recently modified to support a Space Shuttle Main Engine (SSME) hot-fire test. Cold Flow and Flight Readiness Testing were completed before the test was cancelled. Currently, the program is focused on supporting NASA MSFC in accomplishing development testing of the Fastrac Engine, part of NASA's Low Cost Technologies (LCT) Program. MRECS will be used for all engine development testing.

  6. Analytical study of nozzle performance for nuclear thermal rockets

    NASA Technical Reports Server (NTRS)

    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.

  7. 30. Historic view of twentythousandpound rocket test stand with engine ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    30. Historic view of twenty-thousand-pound rocket test stand with engine installation in test cell of Building 202, looking down from elevated location, September 1957. On file at NASA Plumbrook Research Center, Sandusky, Ohio. NASA GRC photo number C-45872. - Rocket Engine Testing Facility, GRC Building No. 202, NASA Glenn Research Center, Cleveland, Cuyahoga County, OH

  8. 13. Historic drawing of rocket engine test facility layout, including ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    13. Historic drawing of rocket engine test facility layout, including Buildings 202, 205, 206, and 206A, February 3, 1984. NASA GRC drawing number CF-101539. On file at NASA Glenn Research Center. - Rocket Engine Testing Facility, NASA Glenn Research Center, Cleveland, Cuyahoga County, OH

  9. 5. Historic photo of scale model of rocket engine test ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    5. Historic photo of scale model of rocket engine test facility, June 18, 1957. On file at NASA Plumbrook Research Center, Sandusky, Ohio. NASA GRC photo number C-45264. - Rocket Engine Testing Facility, NASA Glenn Research Center, Cleveland, Cuyahoga County, OH

  10. 10. Historic photo of rendering of rocket engine test facility ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    10. Historic photo of rendering of rocket engine test facility complex, April 28, 1964. On file at NASA Plumbrook Research Center, Sandusky, Ohio. NASA GRC photo number C-69472. - Rocket Engine Testing Facility, NASA Glenn Research Center, Cleveland, Cuyahoga County, OH

  11. 11. Historic photo of cutaway rendering of rocket engine test ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    11. Historic photo of cutaway rendering of rocket engine test facility complex, June 11, 1965. On file at NASA Plumbrook Research Center, Sandusky, Ohio. NASA GRC photo number C-74433. - Rocket Engine Testing Facility, NASA Glenn Research Center, Cleveland, Cuyahoga County, OH

  12. 6. Historic photo of rocket engine test facility Building 202 ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    6. Historic photo of rocket engine test facility Building 202 complex in operation at night, September 12, 1957. On file at NASA Plumbrook Research Center, Sandusky, Ohio. NASA GRC photo number C-45924. - Rocket Engine Testing Facility, NASA Glenn Research Center, Cleveland, Cuyahoga County, OH

  13. 9. Historic aerial photo of rocket engine test facility complex, ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    9. Historic aerial photo of rocket engine test facility complex, June 11, 1965. On file at NASA Plumbrook Research Center, Sandusky, Ohio. NASA GRC photo number C-65-1270. - Rocket Engine Testing Facility, NASA Glenn Research Center, Cleveland, Cuyahoga County, OH

  14. 8. Historic aerial photo of rocket engine test facility complex, ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    8. Historic aerial photo of rocket engine test facility complex, June 11, 1965. On file at NASA Plumbrook Research Center, Sandusky, Ohio. NASA GRC photo number C-65-1271. - Rocket Engine Testing Facility, NASA Glenn Research Center, Cleveland, Cuyahoga County, OH

  15. 29. Historic view of twentythousandpound rocket test stand with engine ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    29. Historic view of twenty-thousand-pound rocket test stand with engine installation in test cell of Building 202, September 1957. On file at NASA Plumbrook Research Center, Sandusky, Ohio. NASA GRC photo number C-45870. - Rocket Engine Testing Facility, GRC Building No. 202, NASA Glenn Research Center, Cleveland, Cuyahoga County, OH

  16. Quasi-One-Dimensional Modeling of Pulse Detonation Rocket Engines

    NASA Technical Reports Server (NTRS)

    Morris, Christopher I.

    2002-01-01

    Pulsed detonation rocket engines (PDREs) have generated considerable 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 a great deal of 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 difficulties in comparing the available experimental measurements with numerical models. In a previous paper by the author, parametric studies of the performance of a single, straight-tube PDRE were reported. A 1-D, unsteady method of characteristics code, employing a constant-gamma assumption behind the detonation front, was developed for that study. Models of this type are computationally inexpensive, and are particularly useful for parametric performance comparisons. For example, a plot showing the specific impulse of various PDRE and steady-state rocket engine (SSRE) configurations as a function of blowdown pressure ratio. The performance curves clearly indicate that a straight-tube PDRE is superior in specific impulse to a SSRE with a sonic nozzle over the entire range of pressure ratios. Note, however, that a straight-tube PDRE in general does not compare favorably to a SSRE fitted with an optimized de Laval supersonic nozzle, particularly at the high pressure ratios typical for boost or in-space rocket applications. However, the calculations also show that if a dynamically optimized, supersonic de Laval nozzle could be could be fitted to a PDRE, then the specific impulse of the device would exceed that of a comparable SSRE

  17. Atlas Centaur Rocket With Reusable Booster Engines

    NASA Technical Reports Server (NTRS)

    Martin, James A.

    1993-01-01

    Proposed modification of Atlas Centaur enables reuse of booster engines. Includes replacement of current booster engines with engine of new design in which hydrogen used for both cooling and generation of power. Use of hydrogen in new engine eliminates coking and clogging and improves performance significantly. Primary advantages: reduction of cost; increased reliability; and increased payload.

  18. The development of Soviet rocket engines (For strategic missiles)

    SciTech Connect

    Bolonkin, A.

    1991-01-01

    A first-hand account of developments in the Soviet rocket industry is presented. The organization and leadership of the rocket and missile industry are traced from its beginning in the 1920s. The development of the Glushko Experimental Design Bureau, where the majority of Soviet rocket engines were created, is related. The evolution of Soviet rocket engines is traced in regard to both their technical improvement and their application in missiles and space vehicles. Improved Glushko engines and specialized Isaev and Kosberg engines are discussed. The difficulties faced by the Soviet missile and space program, such as the pre-Sputnik failures, the oscillation problem of 1965/1966, which exposed a weakness in Soviet ICBM missiles, and the Nedelin disaster of 1960, which cost the lives of more than 200 scientists and engineers, as well as the Commander-in-Chief of the Strategic Rocket Forces, Marshall Nedelin, are examined. 122 refs.

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

    SciTech Connect

    David L. Black

    2000-06-04

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

  20. J-2X Rocket Engine, 40-Second Test

    NASA Video Gallery

    NASA conducted a 40-second test of the J-2X rocket engine Sept. 28, the most recent in a series of tests of the next-generation engine selected as part of the Space Launch System architecture that ...

  1. The Prediction of Nozzle Performance and Heat Transfer in Hydrogen/Oxygen Rocket Engines with Transpiration Cooling, Film Cooling, and High Area Ratios

    NASA Technical Reports Server (NTRS)

    Kacynski, Kenneth J.; Hoffman, Joe D.

    1994-01-01

    An advanced engineering computational model has been developed to aid in the analysis of chemical rocket engines. The complete multispecies, chemically reacting and diffusing Navier-Stokes equations are modelled, including the Soret thermal diffusion and Dufour energy transfer terms. Demonstration cases are presented for a 1030:1 area ratio nozzle, a 25 lbf film-cooled nozzle, and a transpiration-cooled plug-and-spool rocket engine. The results indicate that the thrust coefficient predictions of the 1030:1 nozzle and the film-cooled nozzle are within 0.2 to 0.5 percent, respectively, of experimental measurements. Further, the model's predictions agree very well with the heat transfer measurements made in all of the nozzle test cases. It is demonstrated that thermal diffusion has a significant effect on the predicted mass fraction of hydrogen along the wall of the nozzle and was shown to represent a significant fraction of the diffusion fluxes occurring in the transpiration-cooled rocket engine.

  2. Modular Rocket Engine Control Software (MRECS)

    NASA Technical Reports Server (NTRS)

    Tarrant, Charlie; Crook, Jerry

    1997-01-01

    The Modular Rocket Engine Control Software (MRECS) Program is a technology demonstration effort designed to advance the state-of-the-art in launch vehicle propulsion systems. Its emphasis is on developing and demonstrating a modular software architecture for a generic, advanced engine control system that will result in lower software maintenance (operations) costs. It effectively accommodates software requirements changes that occur due to hardware. technology upgrades and engine development testing. Ground rules directed by MSFC were to optimize modularity and implement the software in the Ada programming language. MRECS system software and the software development environment utilize Commercial-Off-the-Shelf (COTS) products. This paper presents the objectives and benefits of the program. The software architecture, design, and development environment are described. MRECS tasks are defined and timing relationships given. Major accomplishment are listed. MRECS offers benefits to a wide variety of advanced technology programs in the areas of modular software, architecture, reuse software, and reduced software reverification time related to software changes. Currently, the program is focused on supporting MSFC in accomplishing a Space Shuttle Main Engine (SSME) hot-fire test at Stennis Space Center and the Low Cost Boost Technology (LCBT) Program.

  3. Finite area combustor theoretical rocket performance

    NASA Technical Reports Server (NTRS)

    Gordon, Sanford; Mcbride, Bonnie J.

    1988-01-01

    Previous to this report, the computer program of NASA SP-273 and NASA TM-86885 was capable of calculating theoretical rocket performance based only on the assumption of an infinite area combustion chamber (IAC). An option was added to this program which now also permits the calculation of rocket performance based on the assumption of a finite area combustion chamber (FAC). In the FAC model, the combustion process in the cylindrical chamber is assumed to be adiabatic, but nonisentropic. This results in a stagnation pressure drop from the injector face to the end of the chamber and a lower calculated performance for the FAC model than the IAC model.

  4. Evaluation of undeveloped rocket engine cycle applications to advanced transportation

    NASA Technical Reports Server (NTRS)

    1990-01-01

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

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

    NASA Technical Reports Server (NTRS)

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

    2002-01-01

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

  6. Parallelization of Rocket Engine Simulator Software (PRESS)

    NASA Technical Reports Server (NTRS)

    Cezzar, Ruknet

    1997-01-01

    Parallelization of Rocket Engine System Software (PRESS) project is part of a collaborative effort with Southern University at Baton Rouge (SUBR), University of West Florida (UWF), and Jackson State University (JSU). The second-year funding, which supports two graduate students enrolled in our new Master's program in Computer Science at Hampton University and the principal investigator, have been obtained for the period from October 19, 1996 through October 18, 1997. The key part of the interim report was new directions for the second year funding. This came about from discussions during Rocket Engine Numeric Simulator (RENS) project meeting in Pensacola on January 17-18, 1997. At that time, a software agreement between Hampton University and NASA Lewis Research Center had already been concluded. That agreement concerns off-NASA-site experimentation with PUMPDES/TURBDES software. Before this agreement, during the first year of the project, another large-scale FORTRAN-based software, Two-Dimensional Kinetics (TDK), was being used for translation to an object-oriented language and parallelization experiments. However, that package proved to be too complex and lacking sufficient documentation for effective translation effort to the object-oriented C + + source code. The focus, this time with better documented and more manageable PUMPDES/TURBDES package, was still on translation to C + + with design improvements. At the RENS Meeting, however, the new impetus for the RENS projects in general, and PRESS in particular, has shifted in two important ways. One was closer alignment with the work on Numerical Propulsion System Simulator (NPSS) through cooperation and collaboration with LERC ACLU organization. The other was to see whether and how NASA's various rocket design software can be run over local and intra nets without any radical efforts for redesign and translation into object-oriented source code. There were also suggestions that the Fortran based code be

  7. Axisymmetric Numerical Modeling of Pulse Detonation Rocket Engines

    NASA Technical Reports Server (NTRS)

    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

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

    NASA Technical Reports Server (NTRS)

    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.

  9. Electrically actuated regulation valves for rocket engines

    NASA Astrophysics Data System (ADS)

    Promper, Chr. F.

    2001-09-01

    From technical point of view, Liquid propellant Rocket Engine (LRE) control architectures have to be designed in such a way that: (i) complexity/weight for control and actuation power of valves is minimal, (ii) the engine working point is as far as possible adapted to actual requirements throughout the mission, and (iii) RAMS requirements are complied with. In this context, investigations related to proportional electrically actuated LRE valves are of particular interest in order to investigate the implications of "All Electric Engine control" or in-flight control of the LRE operating point, identify mechanism and electronic controller architectures which present a maximum potential for standardisation and growth, and provide a hardware demonstration of a representative electric valve which can be used for a LRE firing test. As part of the ESA-GSTP2 program, Techspace Aero conducted a market survey identifying the needs for proportional actuation. Based on these considerations, a Technical Specification has been established and a valve demonstrator has been designed and successfully tested (Alcatel ETCA for electronic aspects).

  10. Software Estimates Costs of Testing Rocket Engines

    NASA Technical Reports Server (NTRS)

    2002-01-01

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

  11. Parametric Model of an Aerospike Rocket Engine

    NASA Technical Reports Server (NTRS)

    Korte, J. J.

    2000-01-01

    A suite of computer codes was assembled to simulate the performance of an aerospike engine and to generate the engine input for the Program to Optimize Simulated Trajectories. First an engine simulator module was developed that predicts the aerospike engine performance for a given mixture ratio, power level, thrust vectoring level, and altitude. This module was then used to rapidly generate the aerospike engine performance tables for axial thrust, normal thrust, pitching moment, and specific thrust. Parametric engine geometry was defined for use with the engine simulator module. The parametric model was also integrated into the iSIGHTI multidisciplinary framework so that alternate designs could be determined. The computer codes were used to support in-house conceptual studies of reusable launch vehicle designs.

  12. Parametric Model of an Aerospike Rocket Engine

    NASA Technical Reports Server (NTRS)

    Korte, J. J.

    2000-01-01

    A suite of computer codes was assembled to simulate the performance of an aerospike engine and to generate the engine input for the Program to Optimize Simulated Trajectories. First an engine simulator module was developed that predicts the aerospike engine performance for a given mixture ratio, power level, thrust vectoring level, and altitude. This module was then used to rapidly generate the aerospike engine performance tables for axial thrust, normal thrust, pitching moment, and specific thrust. Parametric engine geometry was defined for use with the engine simulator module. The parametric model was also integrated into the iSIGHT multidisciplinary framework so that alternate designs could be determined. The computer codes were used to support in-house conceptual studies of reusable launch vehicle designs.

  13. A hybrid rocket engine design for simple low cost sounding rocket use

    NASA Astrophysics Data System (ADS)

    Grubelich, Mark; Rowland, John; Reese, Larry

    1993-06-01

    Preliminary test results on a nitrous oxide/HTPB hybrid rocket engine suitable for powering a small sounding rocket to altitudes of 50-100 K/ft are presented. It is concluded that the advantage of the N2O hybrid engine over conventional solid propellant rocket motors is the ability to obtain long burn times with core burning geometries due to the low regression rate of the fuel. Long burn times make it possible to reduce terminal velocity to minimize air drag losses.

  14. Estimates of the radiation environment for a nuclear rocket engine

    SciTech Connect

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

    1992-12-31

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

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

    SciTech Connect

    Bruce G. Schnitzler; Stanley K. Borowski

    2010-07-01

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

  16. Fiber reinforced superalloys for rocket engines

    NASA Technical Reports Server (NTRS)

    Petrasek, Donald W.; Stephens, Joseph R.

    1989-01-01

    High pressure turbopumps for advanced reusable liquid propellant rocket engines such as that for the Space Shuttle Main Engine (SSME) require turbine blade materials that operate under extreme conditions of temperature, hydrogen environment, high-cycle fatigue loading, thermal fatigue and thermal shock. Such requirements tax the capabilities of current blade materials. Based on projections of properties for tungsten fiber reinforced superalloy (FRS) composites, it was concluded that FRS turbine blades offer the potential of a several fold increase in life and over a 200 C increase in temperature capability over the current SSME blade material. FRS composites were evaluated with respect to mechanical property requirements for SSME blade applications. Compared to the current blade material, the thermal shock resistance of FRS materials is excellent, two to nine times better, and their thermal fatigue resistance is equal to or higher than the current blade material. FRS materials had excellent low and high-cycle fatigue strengths, and thermal shock-induced surface microcracks had no influence on their fatigue strength. The material also exhibited negligible embrittlement when exposed to a hydrogen environment.

  17. Fiber reinforced superalloys for rocket engines

    NASA Technical Reports Server (NTRS)

    Petrasek, Donald W.; Stephens, Joseph R.

    1988-01-01

    High-pressure turbopumps for advanced reusable liquid-propellant rocket engines such as that for the Space Shuttle Main Engine (SSME) require turbine blade materials that operate under extreme conditions of temperature, hydrogen environment, high-cycle fatigue loading, thermal fatigue and thermal shock. Such requirements tax the capabilities of current blade materials. Based on projections of properties for tungsten fiber reinforced superalloy (FRS) composites, it was concluded that FRS turbine blades offer the potential of a several-fold increase in life and over a 200C increase in temperature capability over current SSME blade material. FRS composites were evaluated with respect to mechanical property requirements for SSME blade applications. Compared to the current blade material, the thermal shock resistance of FRS materials is excellent, two to nine times better, and their thermal fatigue resistance is equal to or higher than the current blade material. FRS materials had excellent low and high-cycle fatigue strengths, and thermal shock-induced surface microcracks had no influence on their fatigue strength. The material also exhibited negligible embrittlement when exposed to a hydrogen environment.

  18. Software Estimates Costs of Testing Rocket Engines

    NASA Technical Reports Server (NTRS)

    Smith, C. L.

    2003-01-01

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

  19. NASA Orbit Transfer Rocket Engine Technology Program

    NASA Technical Reports Server (NTRS)

    1984-01-01

    The advanced expander cycle engine with a 15,000 lb thrust level and a 6:1 mixture ratio and optimized performance was used as the baseline for a design study of the hydrogen/oxgyen propulsion system for the orbit transfer vehicle. The critical components of this engine are the thrust chamber, the turbomachinery, the extendible nozzle system, and the engine throttling system. Turbomachinery technology is examined for gears, bearing, seals, and rapid solidification rate turbopump shafts. Continuous throttling concepts are discussed. Components of the OTV engine described include the thrust chamber/nozzle assembly design, nozzles, the hydrogen regenerator, the gaseous oxygen heat exchanger, turbopumps, and the engine control valves.

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

    NASA Astrophysics Data System (ADS)

    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.

  1. Developing a gas rocket performance prediction technique

    NASA Technical Reports Server (NTRS)

    Morgenthaler, J. H.; Moon, L. F.; Stepien, W. R.

    1974-01-01

    A simple, semi-empirical performance correlation/prediction technique applicable to gaseous and liquid propellant rocket engines is presented. Excellent correlations were attained for over 100 test firings by adjusting the computation of the gaseous mixing of an unreactive, coaxial jet using a correlation factor, F, which resulted in prediction of the experimental combustion efficiency for each firing. Static pressure, mean velocity and turbulence intensity in the developing region of non-reactive coaxial jets, typical of those of coaxial injector elements were determined. Detailed profiles were obtained at twelve axial locations (extending from the nozzle exit for a distance of five diameters) downstream from a single element of the Bell Aerospace H2/O2 19-element coaxial injector. These data are compared with analytical predictions made using both eddy viscosity and turbulence kinetic energy mixing models and available computer codes. Comparisons were disappointing, demonstrating the necessity of developing improved turbulence models and computational techniques before detailed predictions of practical coaxial free jet flows are attempted.

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

    NASA Technical Reports Server (NTRS)

    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.

  3. Nuclear rockets: High-performance propulsion for Mars

    SciTech Connect

    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.

  4. Controls concepts for next generation reuseable rocket engines

    NASA Technical Reports Server (NTRS)

    Lorenzo, Carl F.; Merrill, Walter C.; Musgrave, Jefferey L.; Ray, Asok

    1995-01-01

    Three primary issues will drive the design and control used in next generation reuseable rocket engines. In addition to steady-state and dynamic performance, the requirements for increased durability, reliability and operability (with faults) will dictate which new controls and design technologies and features will be brought to bear. An array of concepts which have been brought forward will be tested against the measures of cost and benefit as reflected in the above 'ilities'. This paper examines some of the new concepts and looks for metrics to judge their value.

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

    NASA Technical Reports Server (NTRS)

    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.

  6. Telemetry Boards Interpret Rocket, Airplane Engine Data

    NASA Technical Reports Server (NTRS)

    2009-01-01

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

  7. 29. SATURN ROCKET ENGINE LOCATED ON NORTH SIDE OF STATIC ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    29. SATURN ROCKET ENGINE LOCATED ON NORTH SIDE OF STATIC TEST STAND - DETAILS OF THE EXPANSION NOZZLE. - Marshall Space Flight Center, Saturn Propulsion & Structural Test Facility, East Test Area, Huntsville, Madison County, AL

  8. Integrated Design Methodology for Highly Reliable Liquid Rocket Engine

    NASA Astrophysics Data System (ADS)

    Kuratani, Naoshi; Aoki, Hiroshi; Yasui, Masaaki; Kure, Hirotaka; Masuya, Goro

    The Integrated Design Methodology is strongly required at the conceptual design phase to achieve the highly reliable space transportation systems, especially the propulsion systems, not only in Japan but also all over the world in these days. Because in the past some catastrophic failures caused some losses of mission and vehicle (LOM/LOV) at the operational phase, moreover did affect severely the schedule delays and cost overrun at the later development phase. Design methodology for highly reliable liquid rocket engine is being preliminarily established and investigated in this study. The sensitivity analysis is systematically performed to demonstrate the effectiveness of this methodology, and to clarify and especially to focus on the correlation between the combustion chamber, turbopump and main valve as main components. This study describes the essential issues to understand the stated correlations, the need to apply this methodology to the remaining critical failure modes in the whole engine system, and the perspective on the engine development in the future.

  9. Safety issues in SSTO spacecraft powered by antimatter rocket engines

    NASA Astrophysics Data System (ADS)

    Tarpley, Christopher; Lewis, Mark J.; Kothari, Ajay

    1990-07-01

    Safety issues related to operating an antimatter-powered spacecraft in the atmosphere are examined. Radiation danger from the high-energy gamma radiation produced by the annihilation reaction, radioactivity induced in the rocket engine because of photonuclear interactions between gamma radiation and shield nuclei, and antimatter containment failure are discussed in detail. The dangers to the public and the crew and payload are then quantified. An engine design reducing the risks associated with operation in the atmosphere is proposed, and its performance is assessed by calculating the temperature profile and the heat transfer to hydrogen as it flows through the tungsten shell structure of the engine. It is noted that the requirement for shielding and the state of magnet technology dictate a radiation shield of enormous mass and, as designed, the booster would not be reusable and the mass of tungsten lost.

  10. Rocketdyne/Westinghouse nuclear thermal rocket engine modeling

    NASA Technical Reports Server (NTRS)

    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.

  11. Injector for liquid fueled rocket engine

    NASA Technical Reports Server (NTRS)

    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.

  12. Thrust augmentation nozzle (TAN) concept for rocket engine booster applications

    NASA Astrophysics Data System (ADS)

    Forde, Scott; Bulman, Mel; Neill, Todd

    2006-07-01

    Aerojet used the patented thrust augmented nozzle (TAN) concept to validate a unique means of increasing sea-level thrust in a liquid rocket booster engine. We have used knowledge gained from hypersonic Scramjet research to inject propellants into the supersonic region of the rocket engine nozzle to significantly increase sea-level thrust without significantly impacting specific impulse. The TAN concept overcomes conventional engine limitations by injecting propellants and combusting in an annular region in the divergent section of the nozzle. This injection of propellants at moderate pressures allows for obtaining high thrust at takeoff without overexpansion thrust losses. The main chamber is operated at a constant pressure while maintaining a constant head rise and flow rate of the main propellant pumps. Recent hot-fire tests have validated the design approach and thrust augmentation ratios. Calculations of nozzle performance and wall pressures were made using computational fluid dynamics analyses with and without thrust augmentation flow, resulting in good agreement between calculated and measured quantities including augmentation thrust. This paper describes the TAN concept, the test setup, test results, and calculation results.

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

    NASA Technical Reports Server (NTRS)

    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.

  14. Spray Diagnostics in Rocket Engines Using Phase Doppler Analyzer

    NASA Technical Reports Server (NTRS)

    Isaac, Kakkattukuzhy M.

    1996-01-01

    Characteristics such as drop velocity, drop size, number density, volume flow rate, volume flux, and evaporation rate of the fuel spray in a rocket engine are directly related to engine performance. Several studies of shear coaxial atomization have been done in the past. However, additional work related to sprays at supercritical and transcritical conditions would be useful. The author undertook a study of the feasibility of using a phase Doppler particle analyzer (PDPA) for spray measurements in rocket engines as a part of the Summer Faculty Fellowship Program at the NASA Marshall Space Flight Center. The PDPA is a single particle counter (SPC) system based on light scattering from spherical particles. The PDPA instrument is based on refractive and reflective scattering as opposed to other instruments based on diffractive scattering. The main advantage of the PDPA instrument is its ability to provide point-wise information. The following sections describe the principle of operation of the PDPA system and the data obtained using a PDPA instrument in a spray formed by a commercially available fuel injection nozzle.

  15. Parallel Unsteady Turbopump Simulations for Liquid Rocket Engines

    NASA Technical Reports Server (NTRS)

    Kiris, Cetin C.; Kwak, Dochan; Chan, William

    2000-01-01

    This paper reports the progress being made towards complete turbo-pump simulation capability for liquid rocket engines. Space Shuttle Main Engine (SSME) turbo-pump impeller is used as a test case for the performance evaluation of the MPI and hybrid MPI/Open-MP versions of the INS3D code. Then, a computational model of a turbo-pump has been developed for the shuttle upgrade program. Relative motion of the grid system for rotor-stator interaction was obtained by employing overset grid techniques. Time-accuracy of the scheme has been evaluated by using simple test cases. Unsteady computations for SSME turbo-pump, which contains 136 zones with 35 Million grid points, are currently underway on Origin 2000 systems at NASA Ames Research Center. Results from time-accurate simulations with moving boundary capability, and the performance of the parallel versions of the code will be presented in the final paper.

  16. Supercomputer modeling of hydrogen combustion in rocket engines

    NASA Astrophysics Data System (ADS)

    Betelin, V. B.; Nikitin, V. F.; Altukhov, D. I.; Dushin, V. R.; Koo, Jaye

    2013-08-01

    Hydrogen being an ecological fuel is very attractive now for rocket engines designers. However, peculiarities of hydrogen combustion kinetics, the presence of zones of inverse dependence of reaction rate on pressure, etc. prevents from using hydrogen engines in all stages not being supported by other types of engines, which often brings the ecological gains back to zero from using hydrogen. Computer aided design of new effective and clean hydrogen engines needs mathematical tools for supercomputer modeling of hydrogen-oxygen components mixing and combustion in rocket engines. The paper presents the results of developing verification and validation of mathematical model making it possible to simulate unsteady processes of ignition and combustion in rocket engines.

  17. The method of solid rocket motors firings environmental engineering model

    NASA Astrophysics Data System (ADS)

    Pang, Baojun; Xu, Ke; Peng, Keke; Mi, Yaoqi

    The solid rocket motors firings is one main source of space debris, the solid rocket motors firings model is a part of space debris engineering model. In this paper, researching the NASA and ESA model to achieve an appropriate firing model, using the discrete method to model the solid rocket motors firings; application of the long-term approximation orbit evolution algorithm to calculate the evolution of firings generated by a single solid rocket motors ignition event in space; finally, application of space debris environment space debris density algorithm to calculate the distribution of firings generated by a single solid rocket motors ignition event in space, analysing the influence on the space environment and spacecraft.

  18. Rocket U: Engineers Push New Limits

    NASA Video Gallery

    Rocket University participants at NASA's Kennedy Space Center in Florida created an instrumented capsule and dropped it from a high-altitude balloon during a recent test to find out how the aerodyn...

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

    NASA Technical Reports Server (NTRS)

    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.

  20. Hyperthermal Environments Simulator for Nuclear Rocket Engine Development

    NASA Technical Reports Server (NTRS)

    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.

  1. Development of GRCop-84 for Rocket Engine Applications

    NASA Technical Reports Server (NTRS)

    Ellis, David L.; Nathal, Michael V.

    2007-01-01

    GRCop-84 (Cu-8 at.% Cr-4 at.% Nb) has been under development at the NASA Glenn Research Center for several years. The alloy possesses a unique combination of good thermal conductivity, high elevated temperature strength, long creep lives and long low cycle fatigue lives. The alloy is also more oxidation resistant than pure copper and most competitive alloys. The combination of properties has attracted attention from major rocket engine manufacturers who are interested in the alloy for the combustion chamber liner in their next generation of regeneratively cooled engines. It is also a strong candidate for various components in hypersonic vehicles, which also experience very high heat flux conditions. This presentation will discuss the alloy design strategies used to develop GRCop-84 and the various processing routes available for manufacturing components. The microstructure and mechanical properties of the alloy will be reviewed, as well as the results of actual hot fire testing of subscale rocket combustion chambers. The use of environmental and thermal barrier coatings to extend the performance to even higher levels will also be discussed.

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

    NASA Astrophysics Data System (ADS)

    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

  3. Propellant-Flow-Actuated Rocket Engine Igniter

    NASA Technical Reports Server (NTRS)

    Wollen, Mark

    2013-01-01

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

  4. Unsteady Analyses of Valve Systems in Rocket Engine Testing Environments

    NASA Technical Reports Server (NTRS)

    Shipman, Jeremy; Hosangadi, Ashvin; Ahuja, Vineet

    2004-01-01

    This paper discusses simulation technology used to support the testing of rocket propulsion systems by performing high fidelity analyses of feed system components. A generalized multi-element framework has been used to perform simulations of control valve systems. This framework provides the flexibility to resolve the structural and functional complexities typically associated with valve-based high pressure feed systems that are difficult to deal with using traditional Computational Fluid Dynamics (CFD) methods. In order to validate this framework for control valve systems, results are presented for simulations of a cryogenic control valve at various plug settings and compared to both experimental data and simulation results obtained at NASA Stennis Space Center. A detailed unsteady analysis has also been performed for a pressure regulator type control valve used to support rocket engine and component testing at Stennis Space Center. The transient simulation captures the onset of a modal instability that has been observed in the operation of the valve. A discussion of the flow physics responsible for the instability and a prediction of the dominant modes associated with the fluctuations is presented.

  5. Blood Pump Development Using Rocket Engine Flow Simulation Technology

    NASA Technical Reports Server (NTRS)

    Kiris, Cetin C.; Kwak, Dochan

    2002-01-01

    This viewgraph presentation provides information on the transfer of rocket engine flow simulation technology to work involving the development of blood pumps. Details are offered regarding the design and requirements of mechanical heart assist devices, or VADs (ventricular assist device). There are various computational fluid dynamics issues involved in the visualization of flow in such devices, and these are highlighted and compared to those of rocket turbopumps.

  6. 6. "EXPERIMENTAL ROCKET ENGINE TEST STATION AT AFFTC." A low ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    6. "EXPERIMENTAL ROCKET ENGINE TEST STATION AT AFFTC." A low oblique aerial view of Test Area 1-115, looking south, showing Test Stand 1-3 at left, Instrumentation and Control building 8668 at center, and Test Stand 15 at right. The test area is under construction; no evidence of railroad line in photo. - Edwards Air Force Base, Air Force Rocket Propulsion Laboratory, Leuhman Ridge near Highways 58 & 395, Boron, Kern County, CA

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

    NASA Technical Reports Server (NTRS)

    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. An Object Model for a Rocket Engine Numerical Simulator

    NASA Technical Reports Server (NTRS)

    Mitra, D.; Bhalla, P. N.; Pratap, V.; Reddy, P.

    1998-01-01

    Rocket Engine Numerical Simulator (RENS) is a packet of software which numerically simulates the behavior of a rocket engine. Different parameters of the components of an engine is the input to these programs. Depending on these given parameters the programs output the behaviors of those components. These behavioral values are then used to guide the design of or to diagnose a model of a rocket engine "built" by a composition of these programs simulating different components of the engine system. In order to use this software package effectively one needs to have a flexible model of a rocket engine. These programs simulating different components then should be plugged into this modular representation. Our project is to develop an object based model of such an engine system. We are following an iterative and incremental approach in developing the model, as is the standard practice in the area of object oriented design and analysis of softwares. This process involves three stages: object modeling to represent the components and sub-components of a rocket engine, dynamic modeling to capture the temporal and behavioral aspects of the system, and functional modeling to represent the transformational aspects. This article reports on the first phase of our activity under a grant (RENS) from the NASA Lewis Research center. We have utilized Rambaugh's object modeling technique and the tool UML for this purpose. The classes of a rocket engine propulsion system are developed and some of them are presented in this report. The next step, developing a dynamic model for RENS, is also touched upon here. In this paper we will also discuss the advantages of using object-based modeling for developing this type of an integrated simulator over other tools like an expert systems shell or a procedural language, e.g., FORTRAN. Attempts have been made in the past to use such techniques.

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

    NASA Technical Reports Server (NTRS)

    Finseth, J. L.

    1991-01-01

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

  10. The Viking Orbiter 1975 beryllium INTEREGEN rocket engine assembly.

    NASA Technical Reports Server (NTRS)

    Martinez, R. S.; Mcfarland, B. L.; Fischler, S.

    1972-01-01

    Description of the conversion of the Mariner 9 rocket engine for Viking Orbiter use. Engine conversion consists of replacing the 40:1 expansion area ratio nozzle with a 60:1 nozzle of the internal regeneratively (INTEREGEN) cooled rocket engine. Five converted engines using nitrogen tetroxide and monomethylhydrazine demonstrated thermal stability during the nominal 2730-sec burn, but experienced difficulty at operating extremes. The thermal stability characteristic was treated in two ways. The first treatment consisted of mapping the operating regime of the engine to determine its safest operating boundaries as regards thermal equilibrium. Six engines were used for this purpose. Two of the six engines were then modified to effect the second approach - i.e., extend the operating regime. The engines were modified by permitting fuel injection into the acoustic cavity.