WorldWideScience

Sample records for model rocket engines

  1. Measuring Model Rocket Engine Thrust Curves

    Science.gov (United States)

    Penn, Kim; Slaton, William V.

    2010-01-01

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

  2. Measuring Model Rocket Engine Thrust Curves

    Science.gov (United States)

    Penn, Kim; Slaton, William V.

    2010-01-01

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

  3. Rocketdyne/Westinghouse nuclear thermal rocket engine modeling

    Science.gov (United States)

    Glass, James F.

    1993-01-01

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

  4. Axisymmetric Numerical Modeling of Pulse Detonation Rocket Engines

    Science.gov (United States)

    Morris, Christopher I.

    2005-01-01

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

  5. Introduction to rocket science and engineering

    CERN Document Server

    Taylor, Travis S

    2009-01-01

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

  6. Liquid rocket engine injectors

    Science.gov (United States)

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

    1976-01-01

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

  7. Liquid Rocket Engine Testing

    Science.gov (United States)

    2016-10-21

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

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

    Directory of Open Access Journals (Sweden)

    Zhukov Ilya S.

    2016-01-01

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

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

    OpenAIRE

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

    2016-01-01

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

  10. Double time lag combustion instability model for bipropellant rocket engines

    Science.gov (United States)

    Liu, C. K.

    1973-01-01

    A bipropellant stability model is presented in which feed system inertance and capacitance are treated along with injection pressure drop and distinctly different propellant time lags. The model is essentially an extension of Crocco's and Cheng's monopropellant model to the bipropellant case assuming that the feed system inertance and capacitance along with the resistance are located at the injector. The neutral stability boundaries are computed in terms of these parameters to demonstrate the interaction among them.

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

    Directory of Open Access Journals (Sweden)

    Qiang WEI

    2017-08-01

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

  12. Experimental Investigation on Performance of Pulse Detonation Rocket Engine Model

    Institute of Scientific and Technical Information of China (English)

    LI Qiang; FAN Wei; YAN Chuan-jun; HU Cheng-qi; YE Bin

    2007-01-01

    The PDRE test model used in these experiments utilized kerosene as the fuel, oxygen as oxidizer, and nitrogen as purge gas. The solenoid valves were employed to control intermittent supplies of kerosene, oxygen and purge gas. PDRE test model was 50 mm in inner diameter by 1.2 m long. The DDT (defiagration to detonation transition) enhancement device Shchelkin spiral was used in the test model.The effects of detonation frequency on its time-averaged thrust and specific impulse were experimentally investigated. The obtained results showes that the time-averaged thrust of PDRE test model was approximately proportional to the detonation frequency. For the detonation frequency 20 Hz, the time-averaged thrust was around 107 N, and the specific impulse was around 125 s. The nozzle experiments were conducted using PDRE test model with three traditional nozzles. The experimental results obtained demonstrated that all of those nozzles could augment the thrust and specific impulse. Among those three nozzles, the convergent nozzle had the largest increased augmentation, which was approximately 18%, under the specific condition of the experiment.

  13. Nuclear Rocket Engine Reactor

    CERN Document Server

    Lanin, Anatoly

    2013-01-01

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

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

    Science.gov (United States)

    2015-03-01

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

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

    Science.gov (United States)

    Casiano, Matthew J.

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

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

    Energy Technology Data Exchange (ETDEWEB)

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

    2015-05-15

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

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

    Science.gov (United States)

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

    2015-05-01

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

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

    Science.gov (United States)

    Kassoy, David R.; Norris, Adam

    2016-11-01

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

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

    Science.gov (United States)

    Dooling, Thomas A.

    2007-01-01

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

  20. Regenerative Cooling for Liquid Rocket Engines

    Institute of Scientific and Technical Information of China (English)

    QiFeng

    1995-01-01

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

  1. Summarization on variable liquid thrust rocket engines

    Institute of Scientific and Technical Information of China (English)

    2009-01-01

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

  2. Model Rockets and Microchips.

    Science.gov (United States)

    Fitzsimmons, Charles P.

    1986-01-01

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

  3. Unique nuclear thermal rocket engine

    Energy Technology Data Exchange (ETDEWEB)

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

    1993-01-15

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

  4. Unique nuclear thermal rocket engine

    Science.gov (United States)

    Culver, Donald W.; Rochow, Richard

    1993-06-01

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

  5. Reusable rocket engine optical condition monitoring

    Science.gov (United States)

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

    1987-01-01

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

  6. Computational modeling of nuclear thermal rockets

    Science.gov (United States)

    Peery, Steven D.

    1993-01-01

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

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

    Science.gov (United States)

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

    1973-01-01

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

  8. Software for Collaborative Engineering of Launch Rockets

    Science.gov (United States)

    Stanley, Thomas Troy

    2003-01-01

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

  9. Nitrous Oxide/Paraffin Hybrid Rocket Engines

    Science.gov (United States)

    Zubrin, Robert; Snyder, Gary

    2010-01-01

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

  10. Advanced Vortex Hybrid Rocket Engine (AVHRE) Project

    Data.gov (United States)

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

  11. Advanced Vortex Hybrid Rocket Engine (AVHRE) Project

    Data.gov (United States)

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

  12. Electrodynamic actuators for rocket engine valves

    Science.gov (United States)

    Fiet, O.; Doshi, D.

    1972-01-01

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

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

    Directory of Open Access Journals (Sweden)

    A. V. Rudinskii

    2015-01-01

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

  14. Safety Analysis of Liquid Rocket Engine Using Bayesian Networks

    Institute of Scientific and Technical Information of China (English)

    WANG Hua-wei; YAN Zhi-qiang

    2007-01-01

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

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

    Science.gov (United States)

    Ross, M.

    2009-12-01

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

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

    Science.gov (United States)

    Sutton, George P.

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

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

    Directory of Open Access Journals (Sweden)

    A. V. Aliev

    2015-01-01

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

  18. Additive Manufacturing for Affordable Rocket Engines

    Science.gov (United States)

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

    2016-01-01

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

  19. CFD Simulation of Liquid Rocket Engine Injectors

    Science.gov (United States)

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

    2001-01-01

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

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

    Science.gov (United States)

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

    2014-01-01

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

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

    Data.gov (United States)

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

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

    Science.gov (United States)

    Keba, John E.

    1999-01-01

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

  3. Investigations of Rocket Engine Combustion Emissions During ACCENT

    Science.gov (United States)

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

    2001-12-01

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

  4. The next generation rocket engines

    Science.gov (United States)

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

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

  5. Nuclear thermal rocket engine operation and control

    Science.gov (United States)

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

    1993-06-01

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

  6. MHD thrust vectoring of a rocket engine

    Science.gov (United States)

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

    2016-09-01

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

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

    Science.gov (United States)

    Bond, Ryan Bomar

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

  8. Analysis of a Radioisotope Thermal Rocket Engine

    Science.gov (United States)

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

    2017-01-01

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

  9. Numerical investigation of high-pressure combustion in rocket engines using Flamelet/Progress-variable models

    CERN Document Server

    Coclite, A; De Palma, P; Pascazio, G

    2015-01-01

    The present paper deals with the numerical study of high pressure LOx/H2 or LOx/hydrocarbon combustion for propulsion systems. The present research effort is driven by the continued interest in achieving low cost, reliable access to space and more recently, by the renewed interest in hypersonic transportation systems capable of reducing time-to-destination. Moreover, combustion at high pressure has been assumed as a key issue to achieve better propulsive performance and lower environmental impact, as long as the replacement of hydrogen with a hydrocarbon, to reduce the costs related to ground operations and increase flexibility. The current work provides a model for the numerical simulation of high- pressure turbulent combustion employing detailed chemistry description, embedded in a RANS equations solver with a Low Reynolds number k-omega turbulence model. The model used to study such a combustion phenomenon is an extension of the standard flamelet-progress-variable (FPV) turbulent combustion model combined ...

  10. Rocket Engine Innovations Advance Clean Energy

    Science.gov (United States)

    2012-01-01

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

  11. Parallelization of Rocket Engine System Software (Press)

    Science.gov (United States)

    Cezzar, Ruknet

    1996-01-01

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

  12. Liquid fuel injection elements for rocket engines

    Science.gov (United States)

    Cox, George B., Jr. (Inventor)

    1993-01-01

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

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

    Science.gov (United States)

    Davidian, Kenneth O.

    1990-01-01

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

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

    OpenAIRE

    2015-01-01

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

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

    Science.gov (United States)

    Blotzer, Michael J.; Woods, Jody L.

    2009-01-01

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

  16. Additive Manufacturing a Liquid Hydrogen Rocket Engine

    Science.gov (United States)

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

    2016-01-01

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

  17. Injector for liquid fueled rocket engine

    Science.gov (United States)

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

    2000-01-01

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

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

    Science.gov (United States)

    d'Agostino, L.

    2013-12-01

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

  19. Propellant Flow Actuated Piezoelectric Rocket Engine Igniter Project

    Data.gov (United States)

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

  20. Distributed Rocket Engine Testing Health Monitoring System Project

    Data.gov (United States)

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

  1. Distributed Rocket Engine Testing Health Monitoring System Project

    Data.gov (United States)

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

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

    Science.gov (United States)

    Ellis, David L.

    2016-01-01

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

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

    Science.gov (United States)

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

    2016-10-01

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

  4. Software for Estimating Costs of Testing Rocket Engines

    Science.gov (United States)

    Hines, Merlon M.

    2004-01-01

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

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

    Science.gov (United States)

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

    1987-01-01

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

  6. The Chameleon Solid Rocket Propulsion Model

    Science.gov (United States)

    Robertson, Glen A.

    2010-01-01

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

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

    Directory of Open Access Journals (Sweden)

    Kalomba Mboyi

    2015-04-01

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

  8. Analysis of rocket engine injection combustion processes

    Science.gov (United States)

    Salmon, J. W.

    1976-01-01

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

  9. Linear quadratic servo control of a reusable rocket engine

    Science.gov (United States)

    Musgrave, Jeffrey L.

    1991-01-01

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

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

    Science.gov (United States)

    Lehman, Matthew Kurt

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

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

    Science.gov (United States)

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

    2001-01-01

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

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

    Science.gov (United States)

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

    1998-01-01

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

  13. A Design Tool for Liquid Rocket Engine Injectors

    Science.gov (United States)

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

    2000-01-01

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

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

    Science.gov (United States)

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

    2010-01-01

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

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

    Science.gov (United States)

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

    2004-10-01

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

  16. Grooved Fuel Rings for Nuclear Thermal Rocket Engines

    Science.gov (United States)

    Emrich, William

    2009-01-01

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

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

    CERN Document Server

    Zhang, Wei

    2016-01-01

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

  18. Romanian MRE Rocket Engines Program - An Early Endeavor

    Science.gov (United States)

    Rugescu, R. E.

    2002-01-01

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

  19. Heat transfer in rocket engine combustion chambers and nozzles

    Science.gov (United States)

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

    1993-01-01

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

  20. Ablative Rocket Deflector Testing and Computational Modeling

    Science.gov (United States)

    Allgood, Daniel C.; Lott, Jeffrey W.; Raines, Nickey

    2010-01-01

    A deflector risk mitigation program was recently conducted at the NASA Stennis Space Center. The primary objective was to develop a database that characterizes the behavior of industry-grade refractory materials subjected to rocket plume impingement conditions commonly experienced on static test stands. The program consisted of short and long duration engine tests where the supersonic exhaust flow from the engine impinged on an ablative panel. Quasi time-dependent erosion depths and patterns generated by the plume impingement were recorded for a variety of different ablative materials. The erosion behavior was found to be highly dependent on the material s composition and corresponding thermal properties. For example, in the case of the HP CAST 93Z ablative material, the erosion rate actually decreased under continued thermal heating conditions due to the formation of a low thermal conductivity "crystallization" layer. The "crystallization" layer produced near the surface of the material provided an effective insulation from the hot rocket exhaust plume. To gain further insight into the complex interaction of the plume with the ablative deflector, computational fluid dynamic modeling was performed in parallel to the ablative panel testing. The results from the current study demonstrated that locally high heating occurred due to shock reflections. These localized regions of shock-induced heat flux resulted in non-uniform erosion of the ablative panels. In turn, it was observed that the non-uniform erosion exacerbated the localized shock heating causing eventual plume separation and reversed flow for long duration tests under certain conditions. Overall, the flow simulations compared very well with the available experimental data obtained during this project.

  1. Vacuum plasma spray applications on liquid fuel rocket engines

    Science.gov (United States)

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

    1992-01-01

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

  2. Near-term lunar nuclear thermal rocket engine options

    Science.gov (United States)

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

    1991-01-01

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

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

    Science.gov (United States)

    St. Germain, Brad David

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

  4. Reusable rocket engine turbopump condition monitoring

    Science.gov (United States)

    Hampson, M. E.; Barkhoudarian, S.

    1985-01-01

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

  5. Developing Avionics Hardware and Software for Rocket Engine Testing

    Science.gov (United States)

    Aberg, Bryce Robert

    2014-01-01

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

  6. Real-time diagnostics for a reusable rocket engine

    Science.gov (United States)

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

    1992-01-01

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

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

    Science.gov (United States)

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

    2015-09-01

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

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

    Science.gov (United States)

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

    2016-12-01

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

  9. Dual-fuel, dual-mode rocket engine

    Science.gov (United States)

    Martin, James A. (Inventor)

    1989-01-01

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

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

    Science.gov (United States)

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

    1998-01-01

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

  11. Method of fabricating a rocket engine combustion chamber

    Science.gov (United States)

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

    1993-01-01

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

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

    OpenAIRE

    Dario Pastrone

    2012-01-01

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

  13. Application of Chaboche Model in Rocket Thrust Chamber Analysis

    Science.gov (United States)

    Asraff, Ahmedul Kabir; Suresh Babu, Sheela; Babu, Aneena; Eapen, Reeba

    2017-06-01

    Liquid Propellant Rocket Engines are commonly used in space technology. Thrust chamber is one of the most important subsystems of a rocket engine. The thrust chamber generates propulsive thrust force for flight of the rocket by ejection of combustion products at supersonic speeds. Often double walled construction is employed for these chambers. The thrust chamber investigated here has its hot inner wall fabricated out of a high thermal conductive material like copper alloy and outer wall made of stainless steel. Inner wall is subjected to high thermal and pressure loads during operation of engine due to which it will be in the plastic regime. Main reasons for the failure of such chambers are fatigue in the plastic range (called as low cycle fatigue since the number of cycles to failure will be low in plastic range), creep and thermal ratcheting. Elasto plastic material models are required to simulate the above effects through a cyclic stress analysis. This paper gives the details of cyclic stress analysis carried out for the thrust chamber using different plasticity model combinations available in ANSYS (Version 15) FE code. The best model among the above is applied in the cyclic stress analysis of two dimensional (plane strain and axisymmetric) and three dimensional finite element models of thrust chamber. Cyclic life of the chamber is calculated from stress-strain graph obtained from above analyses.

  14. Oxidation of Copper Alloy Candidates for Rocket Engine Applications

    Science.gov (United States)

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

    2002-01-01

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

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

    Science.gov (United States)

    Greene, William D. (Inventor)

    2008-01-01

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

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

    Science.gov (United States)

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

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

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

    Science.gov (United States)

    Girault, J. PH.

    1992-01-01

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

  18. Standardization of Rocket Engine Pulse Time Parameters

    Science.gov (United States)

    Larin, Max E.; Lumpkin, Forrest E.; Rauer, Scott J.

    2001-01-01

    Plumes of bipropellant thrusters are a source of contamination. Small bipropellant thrusters are often used for spacecraft attitude control and orbit correction. Such thrusters typically operate in a pulse mode, at various pulse lengths. Quantifying their contamination effects onto spacecraft external surfaces is especially important for long-term complex-geometry vehicles, e.g. International Space Station. Plume contamination tests indicated the presence of liquid phase contaminant in the form of droplets. Their origin is attributed to incomplete combustion. Most of liquid-phase contaminant is generated during the startup and shutdown (unsteady) periods of thruster pulse. These periods are relatively short (typically 10-50 ms), and the amount of contaminant is determined by the thruster design (propellant valve response, combustion chamber size, thruster mass flow rate, film cooling percentage, dribble volume, etc.) and combustion process organization. Steady-state period of pulse is characterized by much lower contamination rates, but may be lengthy enough to significantly conh'ibute to the overall contamination effect. Because there was no standard methodology for thruster pulse time division, plume contamination tests were conducted at various pulse durations, and their results do not allow quantifying contaminant amounts from each portion of the pulse. At present, the ISS plume contamination model uses an assumption that all thrusters operate in a pulse mode with the pulse length being 100 ms. This assumption may lead to a large difference between the actual amounts of contaminant produced by the thruster and the model predictions. This paper suggests a way to standardize thruster startup and shutdown period definitions, and shows the usefulness of this approach to better quantify thruster plume contamination. Use of the suggested thruster pulse time-division technique will ensure methodological consistency of future thruster plume contamination test programs

  19. Rocket

    Directory of Open Access Journals (Sweden)

    K. Karmarkar

    1952-09-01

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

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

    Directory of Open Access Journals (Sweden)

    Luís Antonio Silva

    2011-05-01

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

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

    Science.gov (United States)

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

    2005-01-01

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

  2. Multiple dopant injection system for small rocket engines

    Science.gov (United States)

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

    1992-07-01

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

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

    Science.gov (United States)

    Betts, Erin M.; Hardin, Andy

    2011-01-01

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

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

    Science.gov (United States)

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

    2000-01-01

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

  5. Effects of rocket engines on laser during lunar landing

    Energy Technology Data Exchange (ETDEWEB)

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

    2013-11-15

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

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

    Science.gov (United States)

    Ray, Asok; Dai, Xiaowen

    1995-01-01

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

  7. More-Accurate Model of Flows in Rocket Injectors

    Science.gov (United States)

    Hosangadi, Ashvin; Chenoweth, James; Brinckman, Kevin; Dash, Sanford

    2011-01-01

    An improved computational model for simulating flows in liquid-propellant injectors in rocket engines has been developed. Models like this one are needed for predicting fluxes of heat in, and performances of, the engines. An important part of predicting performance is predicting fluctuations of temperature, fluctuations of concentrations of chemical species, and effects of turbulence on diffusion of heat and chemical species. Customarily, diffusion effects are represented by parameters known in the art as the Prandtl and Schmidt numbers. Prior formulations include ad hoc assumptions of constant values of these parameters, but these assumptions and, hence, the formulations, are inaccurate for complex flows. In the improved model, these parameters are neither constant nor specified in advance: instead, they are variables obtained as part of the solution. Consequently, this model represents the effects of turbulence on diffusion of heat and chemical species more accurately than prior formulations do, and may enable more-accurate prediction of mixing and flows of heat in rocket-engine combustion chambers. The model has been implemented within CRUNCH CFD, a proprietary computational fluid dynamics (CFD) computer program, and has been tested within that program. The model could also be implemented within other CFD programs.

  8. 3-D thermal analysis using finite difference technique with finite element model for improved design of components of rocket engine turbomachines for Space Shuttle Main Engine SSME

    Science.gov (United States)

    Sohn, Kiho D.; Ip, Shek-Se P.

    1988-01-01

    Three-dimensional finite element models were generated and transferred into three-dimensional finite difference models to perform transient thermal analyses for the SSME high pressure fuel turbopump's first stage nozzles and rotor blades. STANCOOL was chosen to calculate the heat transfer characteristics (HTCs) around the airfoils, and endwall effects were included at the intersections of the airfoils and platforms for the steady-state boundary conditions. Free and forced convection due to rotation effects were also considered in hollow cores. Transient HTCs were calculated by taking ratios of the steady-state values based on the flow rates and fluid properties calculated at each time slice. Results are presented for both transient plots and three-dimensional color contour isotherm plots; they were also converted into universal files to be used for FEM stress analyses.

  9. 3-D thermal analysis using finite difference technique with finite element model for improved design of components of rocket engine turbomachines for Space Shuttle Main Engine SSME

    Science.gov (United States)

    Sohn, Kiho D.; Ip, Shek-Se P.

    1988-01-01

    Three-dimensional finite element models were generated and transferred into three-dimensional finite difference models to perform transient thermal analyses for the SSME high pressure fuel turbopump's first stage nozzles and rotor blades. STANCOOL was chosen to calculate the heat transfer characteristics (HTCs) around the airfoils, and endwall effects were included at the intersections of the airfoils and platforms for the steady-state boundary conditions. Free and forced convection due to rotation effects were also considered in hollow cores. Transient HTCs were calculated by taking ratios of the steady-state values based on the flow rates and fluid properties calculated at each time slice. Results are presented for both transient plots and three-dimensional color contour isotherm plots; they were also converted into universal files to be used for FEM stress analyses.

  10. Distributed Health Monitoring System for Reusable Liquid Rocket Engines

    Science.gov (United States)

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

    2009-01-01

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

  11. Russian Nuclear Rocket Engine Design for Mars Exploration

    Institute of Scientific and Technical Information of China (English)

    Vadim Zakirov; Vladimir Pavshook

    2007-01-01

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

  12. Hyperthermal Environments Simulator for Nuclear Rocket Engine Development

    Science.gov (United States)

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

    2011-01-01

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

  13. Boiler and Pressure Balls Monopropellant Thermal Rocket Engine

    Science.gov (United States)

    Greene, William D. (Inventor)

    2009-01-01

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

  14. Using Innovative Technologies for Manufacturing Rocket Engine Hardware

    Science.gov (United States)

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

    2011-01-01

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

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

    Science.gov (United States)

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

    1990-01-01

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

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

    Science.gov (United States)

    Tejwani, Gopal D.

    2010-01-01

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

  17. Using Innovative Techniques for Manufacturing Rocket Engine Hardware

    Science.gov (United States)

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

    2011-01-01

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

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

    Science.gov (United States)

    Wua, Jianjun; Tanb, Songlin

    2002-01-01

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

  19. Software for Preprocessing Data from Rocket-Engine Tests

    Science.gov (United States)

    Cheng, Chiu-Fu

    2004-01-01

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

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

    Science.gov (United States)

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

    1971-01-01

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

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

    Science.gov (United States)

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

    2010-01-01

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

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

    Directory of Open Access Journals (Sweden)

    Dario Pastrone

    2012-01-01

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

  3. Low loss injector for liquid propellant rocket engines

    Science.gov (United States)

    Vonpragenau, G. L. (Inventor)

    1986-01-01

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

  4. Materials for advanced rocket engine turbopump turbine blades

    Science.gov (United States)

    Chandler, W. T.

    1985-01-01

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

  5. Performance Increase Verification for a Bipropellant Rocket Engine

    Science.gov (United States)

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

    2008-01-01

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

  6. Integrated model of a composite propellant rocket

    Science.gov (United States)

    Miccio, Francesco

    2016-12-01

    The combustion of composite solid propellants was investigated and an available numerical model was improved for taking into account the change of pressure, when the process occurs in a confined environment, as inside a rocket. The pressure increase upon ignition is correctly described by the improved model for both sandwich and dispersed particles propellants. In the latter case, self-induced fluctuations in the pressure and in all other computed variables occur, as consequence of the periodic rise and depletion of oxidizer particles from the binder matrix. The comparison with the results of the constant pressure model shows a different fluctuating profile of gas velocity, with a possible second order effect induced by the pressure fluctuations.

  7. Plume spectrometry for liquid rocket engine health monitoring

    Science.gov (United States)

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

    1988-01-01

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

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

    Directory of Open Access Journals (Sweden)

    Moritz Schulze

    2016-09-01

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

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

    Science.gov (United States)

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

    1993-01-01

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

  10. Studies of Fission Fragment Rocket Engine Propelled Spacecraft

    Science.gov (United States)

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

    2014-01-01

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

  11. Assumed PDF modeling in rocket combustor simulations

    Science.gov (United States)

    Lempke, M.; Gerlinger, P.; Aigner, M.

    2013-03-01

    In order to account for the interaction between turbulence and chemistry, a multivariate assumed PDF (Probability Density Function) approach is used to simulate a model rocket combustor with finite-rate chemistry. The reported test case is the PennState preburner combustor with a single shear coaxial injector. Experimental data for the wall heat flux is available for this configuration. Unsteady RANS (Reynolds-averaged Navier-Stokes) simulation results with and without the assumed PDF approach are analyzed and compared with the experimental data. Both calculations show a good agreement with the experimental wall heat flux data. Significant changes due to the utilization of the assumed PDF approach can be observed in the radicals, e. g., the OH mass fraction distribution, while the effect on the wall heat flux is insignificant.

  12. Hyperthermal Environments Simulator for Nuclear Rocket Engine Development

    Science.gov (United States)

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

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

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

    Science.gov (United States)

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

    2010-09-01

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

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

    Directory of Open Access Journals (Sweden)

    Zolotorev Nikolay

    2017-01-01

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

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

    Data.gov (United States)

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

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

    Data.gov (United States)

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

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

    Science.gov (United States)

    Thorpe, Douglas G.

    1991-01-01

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

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

    Science.gov (United States)

    1993-01-01

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

  19. Integrated propulsion and power modeling for bimodal nuclear thermal rockets

    Science.gov (United States)

    Clough, Joshua

    Bimodal nuclear thermal rocket (BNTR) engines have been shown to reduce the weight of space vehicles to the Moon, Mars, and beyond by utilizing a common reactor for propulsion and power generation. These savings lead to reduced launch vehicle costs and/or increased mission safety and capability. Experimental work of the Rover/NERVA program demonstrated the feasibility of NTR systems for trajectories to Mars. Numerous recent studies have demonstrated the economic and performance benefits of BNTR operation. Relatively little, however, is known about the reactor-level operation of a BNTR engine. The objective of this dissertation is to develop a numerical BNTR engine model in order to study the feasibility and component-level impact of utilizing a NERVA-derived reactor as a heat source for both propulsion and power. The primary contribution is to provide the first-of-its-kind model and analysis of a NERVA-derived BNTR engine. Numerical component models have been modified and created for the NERVA reactor fuel elements and tie tubes, including 1-D coolant thermodynamics and radial thermal conduction with heat generation. A BNTR engine system model has been created in order to design and analyze an engine employing an expander-cycle nuclear rocket and Brayton cycle power generator using the same reactor. Design point results show that a 316 MWt reactor produces a thrust and specific impulse of 66.6 kN and 917 s, respectively. The same reactor can be run at 73.8 kWt to produce the necessary 16.7 kW electric power with a Brayton cycle generator. This demonstrates the feasibility of BNTR operation with a NERVA-derived reactor but also indicates that the reactor control system must be able to operate with precision across a wide power range, and that the transient analysis of reactor decay heat merits future investigation. Results also identify a significant reactor pressure-drop limitation during propulsion and power-generation operation that is caused by poor tie tube

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

    Science.gov (United States)

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

    2016-11-01

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

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

    Science.gov (United States)

    2010-09-28

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

  2. Tie Tube Heat Transfer Modeling for Bimodal Nuclear Thermal Rockets

    Science.gov (United States)

    Clough, Joshua A.; Starkey, Ryan P.; Lewis, Mark J.; Lavelle, Thomas M.

    2007-01-01

    Bimodal nuclear thermal rocket systems have been shown to reduce the weight and cost of space vehicles to Mars and beyond by utilizing the reactor for power generation in the relatively long duration between burns in an interplanetary trajectory. No information, however, is available regarding engine and reactor-level operation of such bimodal systems. The purpose of this project is to generate engine and reactor models with sufficient fidelity and flexibility to accurately study the component-level effects of operating a propulsion-designed reactor at power generation levels. Previous development of a 1-D reactor and tie tube model found that ignoring heat generation inside of the tie tube leads to under-prediction of the temperature change and over-prediction of pressure change across the tie tube. This paper will present the development and results of a tie tube model that has been extended to account for heat generation, specifically in the moderator layer. This model is based on a 1-D distribution of power in the fuel elements and tie tubes, as a precursor to an eventual neutron-driven reactor model.

  3. ACTIVE MODEL ROCKET STABILIZATION VIA COLD GAS THRUSTERS

    OpenAIRE

    Malyuta, Danylo; Collaud, Xavier; Martins Gaspar, Mikael; Rouaze, Gautier Marie Pierre; Pictet, Raimondo; Ivanov, Anton; Mullin, Nickolay

    2015-01-01

    This paper describes the development and testing of a reaction control system (RCS) for a model rocket named FALCO-4. The rocket uses cold gas jets to keep itself perfectly vertical at low speeds. We first describe the mechanical layout of FALCO-4 and the characteristics of the cold gas propulsion system. We then propose a dynamical model of the rocket and a control scheme based on decoupled PID regulators for roll, pitch and yaw. The control scheme is then evaluated based on MATLAB simulatio...

  4. Dynamical Model of Rocket Propellant Loading with Liquid Hydrogen

    Data.gov (United States)

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

  5. Modeling and Diagnostic Software for Liquefying-Fuel Rockets

    Science.gov (United States)

    Poll, Scott; Iverson, David; Ou, Jeremy; Sanderfer, Dwight; Patterson-Hine, Ann

    2005-01-01

    A report presents a study of five modeling and diagnostic computer programs considered for use in an integrated vehicle health management (IVHM) system during testing of liquefying-fuel hybrid rocket engines in the Hybrid Combustion Facility (HCF) at NASA Ames Research Center. Three of the programs -- TEAMS, L2, and RODON -- are model-based reasoning (or diagnostic) programs. The other two programs -- ICS and IMS -- do not attempt to isolate the causes of failures but can be used for detecting faults. In the study, qualitative models (in TEAMS and L2) and quantitative models (in RODON) having varying scope and completeness were created. Each of the models captured the structure and behavior of the HCF as a physical system. It was noted that in the cases of the qualitative models, the temporal aspects of the behavior of the HCF and the abstraction of sensor data are handled outside of the models, and it is necessary to develop additional code for this purpose. A need for additional code was also noted in the case of the quantitative model, though the amount of development effort needed was found to be less than that for the qualitative models.

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

    Science.gov (United States)

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

    1981-01-01

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

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

    Science.gov (United States)

    Gradl, Paul R.

    2016-01-01

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

  8. Burn Rate Modelling of Solid Rocket Propellants (Short Communication

    Directory of Open Access Journals (Sweden)

    A.R. Kulkarni

    1998-01-01

    Full Text Available A generalised model of burning of a solid rocket propellant based on kinetics of propellant hasbeen developed. A complete set of variables has been formed after examining the existing models.Buckingham theorem provides the functional form of the model, such that the existing models are thesubcases of this generalised model. This proposed model has been validated by an experimental data.

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

    Science.gov (United States)

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

    2010-01-01

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

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

    Science.gov (United States)

    Hickox, C. E.

    1975-01-01

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

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

    Directory of Open Access Journals (Sweden)

    ARSHAD A.

    2012-07-01

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

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

    Science.gov (United States)

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

    2004-10-01

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

  13. Storable Hypergolic Solid Fuel for Hybrid Rocket Engines

    Directory of Open Access Journals (Sweden)

    R. V. Singh

    1976-07-01

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

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

    Science.gov (United States)

    2017-07-26

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

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

    Science.gov (United States)

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

    2013-01-01

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

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

    Science.gov (United States)

    Schoenman, L.

    1981-01-01

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

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

    Science.gov (United States)

    Schumaker, Stephen Alexander

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

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

    Science.gov (United States)

    Myers, David E.; Kosareo, Daniel N.

    2006-01-01

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

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

    Energy Technology Data Exchange (ETDEWEB)

    Friesen, H.N.

    1995-06-01

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

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

    Data.gov (United States)

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

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

    Science.gov (United States)

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

    2017-01-01

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

  2. Effects of turbulent and spray models on combustion process simultion of LOX/GH2 rocket engine%湍流、喷雾模型对氢氧火箭发动机燃烧仿真的影响

    Institute of Scientific and Technical Information of China (English)

    程钰锋; 聂万胜; 丰松江

    2011-01-01

    Based on the improved PISO algorithm, the numerical simulation for the combustion instability of a LOX/GH2 rocket engine was conducted by changing the turbulent and spay models of κ-ε equations. The compared results of the theoretical analysis and numerical simulation show that in the two-dimensional situation, both droplet collision model and TAB droplet breakup model are not suitable for the numerical simulation of LOX/GH2 combustion instability; the pressure oscillation in the combustion chamber can be simulated by combining the TVB droplet breakup model with the turbulent models of κ-ε equations, but the oscillation frequency can not be simulated; if the turbulent models of Realizable κ-ε equations are adopted without consideration of the droplet spray models, both the pressure oscillation in the combustion chamber and the distribution of the oscillation frequency can be simulated.%基于完善的压力隐式算子分裂(PISO)算法,通过改变κ-ε两方程湍流模型和喷雾模型,对氢氧火箭发动机不稳定燃烧进行数值仿真。比较理论分析和数值仿真的结果得出,在二维情况下,液滴碰撞模型和TAB液滴破碎模型不适于模拟氢氧火箭发动机不稳定燃烧;TVB液滴破碎模型与κ-ε两方程湍流模型的组合情况能够捕捉到燃烧室中的压力振荡,但不能体现出振荡频率;而采用Realizableκ-ε湍流模型不考虑液滴雾化模型时不但能够捕捉燃烧室内压力振荡情况,还能够很好地得出振荡频率的分布情况。

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

    Science.gov (United States)

    Adams, A.

    1973-01-01

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

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

    OpenAIRE

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

    2010-01-01

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

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

    Science.gov (United States)

    Shi, John J.

    2005-01-01

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

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

    Energy Technology Data Exchange (ETDEWEB)

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

    2015-06-15

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

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

    NARCIS (Netherlands)

    Sliphorst, M.

    2011-01-01

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

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

    NARCIS (Netherlands)

    Sliphorst, M.

    2011-01-01

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

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

    Science.gov (United States)

    Zhang, Yu-Lin

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

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

    OpenAIRE

    Sliphorst, M.

    2011-01-01

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

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

    Science.gov (United States)

    Wang, Ten-See; Cornelison, Joni

    1996-01-01

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

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

    Science.gov (United States)

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

    1992-07-01

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

  13. Lightweight Exit Cone for Liquid Rocket Engines Project

    Data.gov (United States)

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

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

    Science.gov (United States)

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

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

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

    Directory of Open Access Journals (Sweden)

    Luiza Fabrino Favato

    2015-04-01

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

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

    Science.gov (United States)

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

    2003-01-01

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

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

    Science.gov (United States)

    Ali, Moonis; Gupta, U. K.

    1990-01-01

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

  18. State Machine Modeling of the Space Launch System Solid Rocket Boosters

    Science.gov (United States)

    Harris, Joshua A.; Patterson-Hine, Ann

    2013-01-01

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

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

    Science.gov (United States)

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

    2012-01-01

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

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

    Science.gov (United States)

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

    2012-01-01

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

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

    Science.gov (United States)

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

    2000-01-01

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

  2. Health monitoring of rocket engines using image processing

    Science.gov (United States)

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

    1991-07-01

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

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

    Science.gov (United States)

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

    1989-01-01

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

  4. Design Considerations for Human Rating of Liquid Rocket Engines

    Science.gov (United States)

    Parkinson, Douglas

    2010-01-01

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

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

    Science.gov (United States)

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

    1992-01-01

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

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

    Science.gov (United States)

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

    1991-01-01

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

  7. Application of advanced coating techniques to rocket engine components

    Science.gov (United States)

    Verma, S. K.

    1988-01-01

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

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

    Science.gov (United States)

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

    2016-02-01

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

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

    Data.gov (United States)

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

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

    Science.gov (United States)

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

    2011-01-01

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

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

    Science.gov (United States)

    Ozawa, K.; Shimada, T.

    2016-07-01

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

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

    Science.gov (United States)

    Kobayashi, Hiroaki; Sato, Tetsuya; Tanatsugu, Nobuhiro

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

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

    Science.gov (United States)

    Ruiz Haro, Mercedes; Navarro Vásquez, Ricardo M.

    2002-01-01

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

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

    Science.gov (United States)

    Mazzetti, Alessandro; Merotto, Laura; Pinarello, Giordano

    2016-09-01

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

  15. SRM (Solid Rocket Motor) propellant and polymer materials structural modeling

    Science.gov (United States)

    Moore, Carleton J.

    1988-01-01

    The following investigation reviews and evaluates the use of stress relaxation test data for the structural analysis of Solid Rocket Motor (SRM) propellants and other polymer materials used for liners, insulators, inhibitors, and seals. The stress relaxation data is examined and a new mathematical structural model is proposed. This model has potentially wide application to structural analysis of polymer materials and other materials generally characterized as being made of viscoelastic materials. A dynamic modulus is derived from the new model for stress relaxation modulus and is compared to the old viscoelastic model and experimental data.

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

    Science.gov (United States)

    La Botz, R. J.

    1984-01-01

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

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

    Science.gov (United States)

    Liou, Larry C.

    2008-01-01

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

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

    Science.gov (United States)

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

    2017-04-01

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

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

    Science.gov (United States)

    Huppi, Hal; Tobias, Mark; Seiler, James

    2003-01-01

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

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

    Science.gov (United States)

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

    2010-07-01

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

  1. Development of Detonation Modeling Capabilities for Rocket Test Facilities: Hydrogen-Oxygen-Nitrogen Mixtures

    Science.gov (United States)

    Allgood, Daniel C.

    2016-01-01

    The objective of the presented work was to develop validated computational fluid dynamics (CFD) based methodologies for predicting propellant detonations and their associated blast environments. Applications of interest were scenarios relevant to rocket propulsion test and launch facilities. All model development was conducted within the framework of the Loci/CHEM CFD tool due to its reliability and robustness in predicting high-speed combusting flow-fields associated with rocket engines and plumes. During the course of the project, verification and validation studies were completed for hydrogen-fueled detonation phenomena such as shock-induced combustion, confined detonation waves, vapor cloud explosions, and deflagration-to-detonation transition (DDT) processes. The DDT validation cases included predicting flame acceleration mechanisms associated with turbulent flame-jets and flow-obstacles. Excellent comparison between test data and model predictions were observed. The proposed CFD methodology was then successfully applied to model a detonation event that occurred during liquid oxygen/gaseous hydrogen rocket diffuser testing at NASA Stennis Space Center.

  2. Facility for cold flow testing of solid rocket motor models

    Science.gov (United States)

    Bacchus, D. L.; Hill, O. E.; Whitesides, R. Harold

    1992-02-01

    A new cold flow test facility was designed and constructed at NASA Marshall Space Flight Center for the purpose of characterizing the flow field in the port and nozzle of solid propellant rocket motors (SRM's). A National Advisory Committee was established to include representatives from industry, government agencies, and universities to guide the establishment of design and instrumentation requirements for the new facility. This facility design includes the basic components of air storage tanks, heater, submicron filter, quiet control valve, venturi, model inlet plenum chamber, solid rocket motor (SRM) model, exhaust diffuser, and exhaust silencer. The facility was designed to accommodate a wide range of motor types and sizes from small tactical motors to large space launch boosters. This facility has the unique capability of testing ten percent scale models of large boosters such as the new Advanced Solid Rocket Motor (ASRM), at full scale motor Reynolds numbers. Previous investigators have established the validity of studying basic features of solid rocket motor development programs include the acquisition of data to (1) directly evaluate and optimize the design configuration of the propellant grain, insulation, and nozzle; and (2) provide data for validation of the computational fluid dynamics, (CFD), analysis codes and the performance analysis codes. A facility checkout model was designed, constructed, and utilized to evaluate the performance characteristics of the new facility. This model consists of a cylindrical chamber and converging/diverging nozzle with appropriate manifolding to connect it to the facility air supply. It was designed using chamber and nozzle dimensions to simulate the flow in a 10 percent scale model of the ASRM. The checkout model was recently tested over the entire range of facility flow conditions which include flow rates from 9.07 to 145 kg/sec (20 to 320 Ibm/sec) and supply pressure from 5.17 x 10 exp 5 to 8.27 x 10 exp 6 Pa. The

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

    Energy Technology Data Exchange (ETDEWEB)

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

    1992-09-01

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

  4. Development of Efficient Real-Fluid Model in Simulating Liquid Rocket Injector Flows

    Science.gov (United States)

    Cheng, Gary; Farmer, Richard

    2003-01-01

    The characteristics of propellant mixing near the injector have a profound effect on the liquid rocket engine performance. However, the flow features near the injector of liquid rocket engines are extremely complicated, for example supercritical-pressure spray, turbulent mixing, and chemical reactions are present. Previously, a homogeneous spray approach with a real-fluid property model was developed to account for the compressibility and evaporation effects such that thermodynamics properties of a mixture at a wide range of pressures and temperatures can be properly calculated, including liquid-phase, gas- phase, two-phase, and dense fluid regions. The developed homogeneous spray model demonstrated a good success in simulating uni- element shear coaxial injector spray combustion flows. However, the real-fluid model suffered a computational deficiency when applied to a pressure-based computational fluid dynamics (CFD) code. The deficiency is caused by the pressure and enthalpy being the independent variables in the solution procedure of a pressure-based code, whereas the real-fluid model utilizes density and temperature as independent variables. The objective of the present research work is to improve the computational efficiency of the real-fluid property model in computing thermal properties. The proposed approach is called an efficient real-fluid model, and the improvement of computational efficiency is achieved by using a combination of a liquid species and a gaseous species to represent a real-fluid species.

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

    Science.gov (United States)

    Gradl, Paul

    2016-01-01

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

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

    Institute of Scientific and Technical Information of China (English)

    庄逢辰; 聂万胜; 邹勤; 张中光

    2001-01-01

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

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

    Science.gov (United States)

    Zubrin, Robert M.

    1991-01-01

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

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

    CERN Document Server

    Kitsche, Wolfgang

    2010-01-01

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

  9. Viscoelastic Modelling of Solid Rocket Propellants using Maxwell Fluid Model

    Directory of Open Access Journals (Sweden)

    Himanshu Shekhar

    2010-07-01

    Full Text Available Maxwell fluid model consisting of a spring and a dashpot in series is applied for viscoelastic characterisation of solid rocket propellants. Suitable values of spring constant and damping coefficient wereemployed by least square variation of errors for generation of complete stress-strain curve in uniaxial tensile mode for case-bonded solid propellant formulations. Propellants from the same lot were tested at different strain rates. It was observed that change in spring constant, representing elastic part was very small with strain rate but damping constant varies significantly with variation in strain rate. For a typical propellant formulation, when strain rate was raised from 0.00037/s to 0.185/s, spring constant K changed from 5.5 MPato 7.9 MPa, but damping coefficient D was reduced from 1400 MPa-s to 4 MPa-s. For all strain rates, stress-strain curve was generated using Maxwell model and close matching with actual test curve was observed.This indicates validity of Maxwell fluid model for uniaxial tensile testing curves of case-bonded solid propellant formulations. It was established that at higher strain rate, damping coefficient becomes negligible as compared to spring constant. It was also observed that variation of spring constant is logarithmic with strain rate and that of damping coefficient follows power law. The correlation coefficients were introduced to ascertain spring constants and damping coefficients at any strain rate from that at a reference strain rate. Correlationfor spring constant needs a coefficient H, which is function of propellant formulation alone and not of test conditions and the equation developeds K2 = K1 + H ´ ln{(de2/dt/(de1/dt}. Similarly for damping coefficient D also another constant S is introduced and prediction formula is given by D2 = D1 ´ {(de2/dt/(de1/dt}S.Evaluating constants H and S at different strain rates validate this mathematical formulation for differentpropellant formulations

  10. On the Elastic Vibration Model for High Length-Diameter Ratio Rocket with Attitude Control System

    Institute of Scientific and Technical Information of China (English)

    朱伯立; 杨树兴

    2003-01-01

    An elastic vibration model for high length-diameter ratio spinning rocket with attitude control system which can be used for trajectory simulation is established. The basic theory of elastic dynamics and vibration dynamics were both used to set up the elastic vibration model of rocket body. In order to study the problem more conveniently, the rocket's body was simplified to be an even beam with two free ends. The model was validated by simulation results and the test data.

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

    Science.gov (United States)

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

    2013-11-01

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

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

    Science.gov (United States)

    Ross, M. N.; Toohey, D.

    2008-12-01

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

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

    Science.gov (United States)

    Verma, Satyajit

    2006-01-01

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

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

    Science.gov (United States)

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

    2006-01-01

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

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

    Science.gov (United States)

    Messing, Wesley E.

    1959-01-01

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

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

    Science.gov (United States)

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

    2014-01-01

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

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

    Science.gov (United States)

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

    2017-10-01

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

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

    Science.gov (United States)

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

    2013-01-01

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

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

    Science.gov (United States)

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

    2013-01-01

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

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

    Science.gov (United States)

    Woodward, Roger D.

    1993-01-01

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

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

    Science.gov (United States)

    Kiris, Cetin; Williams, Robert

    2000-01-01

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

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

    Institute of Scientific and Technical Information of China (English)

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

    2011-01-01

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

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

    Science.gov (United States)

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

    2003-01-01

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

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

    Science.gov (United States)

    Trujillo, Abraham Gerardo

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

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

    Science.gov (United States)

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

    1993-06-01

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

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

    Science.gov (United States)

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

    2015-01-01

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

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

    Science.gov (United States)

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

    2014-01-01

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

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

    Science.gov (United States)

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

    2016-11-01

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

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

    Science.gov (United States)

    Holzmann, Wilfried A.; Hayden, Warren R.

    1988-01-01

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

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

    Science.gov (United States)

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

    2017-01-01

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

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

    Science.gov (United States)

    Meyer, Michael L.; Giuliani, James E.

    1993-11-01

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

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

    Directory of Open Access Journals (Sweden)

    Jun Qian

    2014-04-01

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

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

    Science.gov (United States)

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

    2010-01-01

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

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

    Science.gov (United States)

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

    2000-01-01

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

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

    Science.gov (United States)

    Dougherty, N. S.

    2010-01-01

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

  16. Internal Flow Simulation of High-Performance Solid Rockets using a k-ωTurbulence Model

    Institute of Scientific and Technical Information of China (English)

    V.R. SANAL KUMAR; H.D. KIM; B.N. RAGHUNANDAN; T. SETOGUCHI; S. RAGHUNATHAN

    2005-01-01

    @@ For technological reasons many high-performance solid rocket motors are made from segmented propellant grains with non-uniform port geometry. In this paper parametric studies have been carried out to examine the geometric dependence of transient flow features in solid rockets with non-uniform ports. Numerical computations have been carried out in an inert simulator of solid propellant rocket motor with the aid of a standard k-ω turbulence model. It was seen that the damping of the temperature fluctuation is faster in solid rocket with convergent port than with divergent port geometry. We inferred that the damping of the flow fluctuations using the port geometry is a meaningful objective for the suppression and control of the instability and/or pressure/thrust oscillations during the starting transient of solid rockets.

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

    Science.gov (United States)

    Barkhoudarian, Sarkis; Kittinger, Scott

    2006-01-01

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

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

    Energy Technology Data Exchange (ETDEWEB)

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

    2003-11-01

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

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

    Science.gov (United States)

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

    2003-08-01

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

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

    Science.gov (United States)

    Fittje, James E.; Buehrle, Robert J.

    2006-01-01

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

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

    Science.gov (United States)

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

    2002-01-01

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

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

    Science.gov (United States)

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

    2017-02-01

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

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

    Science.gov (United States)

    Kan, Brandon K.

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

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

    Science.gov (United States)

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

    2017-07-01

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

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

    Science.gov (United States)

    Bhat, Biliyar; Greene, Sandra

    2015-01-01

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

  6. Ultrasonic inspection of rocket fuel model using laminated transducer and multi-channel step pulser

    Science.gov (United States)

    Mihara, T.; Hamajima, T.; Tashiro, H.; Sato, A.

    2013-01-01

    For the ultrasonic inspection for the packing of solid fuel in a rocket booster, an industrial inspection is difficult. Because the signal to noise ratio in ultrasonic inspection of rocket fuel become worse due to the large attenuation even using lower frequency ultrasound. For the improvement of this problem, we tried to applied the two techniques in ultrasonic inspection, one was the step function pulser system with the super wideband frequency properties and the other was the laminated element transducer. By combining these two techniques, we developed the new ultrasonic measurement system and demonstrated the advantages in ultrasonic inspection of rocket fuel model specimen.

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

    Data.gov (United States)

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

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

    Science.gov (United States)

    1984-01-01

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

  9. Two phase flow combustion modelling of a ducted rocket

    NARCIS (Netherlands)

    Stowe, R.A.; Dubois, C.; Harris, P.G.; Mayer, A.E.H.J.; Champlain, A. de; Ringuette, S.

    2001-01-01

    Under a co-operative program, the Defence Research Establishment Valcartier and Université Laval in Canada and the TNO Prins Maurits Laboratory in the Netherlands have studied the use of a ducted rocket for missile propulsion. Hot-flow direct-connect combustion experiments using both simulated and s

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

    Science.gov (United States)

    Wang, Ten-See; Chen, Yen-Sen

    1990-07-01

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

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

    Science.gov (United States)

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

    2007-01-01

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

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

    Science.gov (United States)

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

    2007-01-01

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

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

    Science.gov (United States)

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

    2016-01-01

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

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

    Science.gov (United States)

    Brown, T. M.; Smith, Norm

    1999-01-01

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

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

    Science.gov (United States)

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

    2017-03-01

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

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

    Science.gov (United States)

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

    2014-11-01

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

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

    Science.gov (United States)

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

    2010-10-01

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

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

    Science.gov (United States)

    Bennewitz, John William

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

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

    Science.gov (United States)

    Ellis, David L.

    2013-01-01

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

  20. Principles of models based engineering

    Energy Technology Data Exchange (ETDEWEB)

    Dolin, R.M.; Hefele, J.

    1996-11-01

    This report describes a Models Based Engineering (MBE) philosophy and implementation strategy that has been developed at Los Alamos National Laboratory`s Center for Advanced Engineering Technology. A major theme in this discussion is that models based engineering is an information management technology enabling the development of information driven engineering. Unlike other information management technologies, models based engineering encompasses the breadth of engineering information, from design intent through product definition to consumer application.

  1. System Modeling and Diagnostics for Liquefying-Fuel Hybrid Rockets

    Science.gov (United States)

    Poll, Scott; Iverson, David; Ou, Jeremy; Sanderfer, Dwight; Patterson-Hine, Ann

    2003-01-01

    A Hybrid Combustion Facility (HCF) was recently built at NASA Ames Research Center to study the combustion properties of a new fuel formulation that burns approximately three times faster than conventional hybrid fuels. Researchers at Ames working in the area of Integrated Vehicle Health Management recognized a good opportunity to apply IVHM techniques to a candidate technology for next generation launch systems. Five tools were selected to examine various IVHM techniques for the HCF. Three of the tools, TEAMS (Testability Engineering and Maintenance System), L2 (Livingstone2), and RODON, are model-based reasoning (or diagnostic) systems. Two other tools in this study, ICS (Interval Constraint Simulator) and IMS (Inductive Monitoring System) do not attempt to isolate the cause of the failure but may be used for fault detection. Models of varying scope and completeness were created, both qualitative and quantitative. In each of the models, the structure and behavior of the physical system are captured. In the qualitative models, the temporal aspects of the system behavior and the abstraction of sensor data are handled outside of the model and require the development of additional code. In the quantitative model, less extensive processing code is also necessary. Examples of fault diagnoses are given.

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

    Science.gov (United States)

    Trushlyakov, V.; Shatrov, Ya.

    2017-09-01

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

  3. A Mathematical and Numerical Model for the Analysis of Hybrid Rocket Motors

    Directory of Open Access Journals (Sweden)

    Florin MINGIREANU

    2011-12-01

    Full Text Available The hybrid rocket motors (HRM use a two-phase propellant system. This offers some remarkable advantages but also arises some difficulties like the neutralization of their instabilities. The non-acoustic combustion instabilities are high-amplitude pressure oscillations that have too low frequencies to be associated with acoustics. Acoustic type combustion instabilities are self-excited oscillations generated by the interaction between acoustic waves and combustion. The goal of the present work is to develop a simplified model of the coupling of the hybrid combustion process with the complete unsteady flow, starting from the combustion port and ending with the nozzle. This model must be useful for transient and stability analysis and also for scaling of HRMs. The numerical results obtained with our model show a good agreement with published experimental and numerical results. The computational and stability analysis models developed in this work are simple, computationally efficient and offer the advantage of taking into account a large number of functional and constructive parameters that are used by the engineers.

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

    Institute of Scientific and Technical Information of China (English)

    徐峰; 刘英元; 陈海峰

    2012-01-01

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

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

    Science.gov (United States)

    Nagatomo, Makoto; Kaya, Nobuyuki; Matsumoto, Hiroshi

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

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

    Science.gov (United States)

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

    2001-01-01

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

  7. Low-thrust rocket trajectories

    Energy Technology Data Exchange (ETDEWEB)

    Keaton, P.W.

    1986-01-01

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

  8. Low-thrust rocket trajectories

    Energy Technology Data Exchange (ETDEWEB)

    Keaton, P.W.

    1987-03-01

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

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

    Directory of Open Access Journals (Sweden)

    N. S. Dorosh

    2014-01-01

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

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

    Science.gov (United States)

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

    1950-01-01

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

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

    Science.gov (United States)

    Csomor, A.

    1974-01-01

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

  12. Fuel/Oxidizer Injector Modeling in Sub- and Super-Critical Regimes for Deep Throttling Cryogenic Engines Project

    Data.gov (United States)

    National Aeronautics and Space Administration — Accurate CFD modeling of fuel/oxidizer injection and combustion is needed to design and analyze liquid rocket engines. Currently, however, there is no mature...

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

    Science.gov (United States)

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

    2009-11-01

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

  14. Mathematical Modelling of In-Chamber Processes in Hydrocombined Propellant Solid Rocket Motors

    Directory of Open Access Journals (Sweden)

    Nikolai A. Obukhov

    1998-10-01

    Full Text Available The special conditions of employment of commercial rockets in the sea environment has opened up new possibilities of improving motor performance. The interesting method suggests supplying water into the running motor. This paper reports the calculations and experiments carried out with solid propellant model setups. The results prove the validity of the proposed method and allow the refinement of calculation techniques for the prediction of solid rocket motor performance characteristics. The serviceability of the solid propellant charges working in combination with water is demonstrated. A mathematical model is proposed for the operation of a hydrocombined propellant motor with water and powdered additives applied to the combustion chamber."

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

    Science.gov (United States)

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

    2016-01-01

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

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

    Institute of Scientific and Technical Information of China (English)

    张黎辉; 李伟; 段娜

    2011-01-01

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

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

    Science.gov (United States)

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

    2016-01-01

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

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

    Science.gov (United States)

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

    2016-01-01

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

  19. Rocket Combustor Validation Data for Advanced Combustion Models Project

    Data.gov (United States)

    National Aeronautics and Space Administration — The pace and cost of developing an engine system for future explorations is strongly influenced by the inadequacies of design tools and the supporting databases. The...

  20. One Dimensional Analysis Model of a Condensing Spray Chamber Including Rocket Exhaust Using SINDA/FLUINT and CEA

    Science.gov (United States)

    Sakowski, Barbara; Edwards, Daryl; Dickens, Kevin

    2014-01-01

    Modeling droplet condensation via CFD codes can be very tedious, time consuming, and inaccurate. CFD codes may be tedious and time consuming in terms of using Lagrangian particle tracking approaches or particle sizing bins. Also since many codes ignore conduction through the droplet and or the degradating effect of heat and mass transfer if noncondensible species are present, the solutions may be inaccurate. The modeling of a condensing spray chamber where the significant size of the water droplets and the time and distance these droplets take to fall, can make the effect of droplet conduction a physical factor that needs to be considered in the model. Furthermore the presence of even a relatively small amount of noncondensible has been shown to reduce the amount of condensation [Ref 1]. It is desirable then to create a modeling tool that addresses these issues. The path taken to create such a tool is illustrated. The application of this tool and subsequent results are based on the spray chamber in the Spacecraft Propulsion Research Facility (B2) located at NASA's Plum Brook Station that tested an RL-10 engine. The platform upon which the condensation physics is modeled is SINDAFLUINT. The use of SINDAFLUINT enables the ability to model various aspects of the entire testing facility, including the rocket exhaust duct flow and heat transfer to the exhaust duct wall. The ejector pumping system of the spray chamber is also easily implemented via SINDAFLUINT. The goal is to create a transient one dimensional flow and heat transfer model beginning at the rocket, continuing through the condensing spray chamber, and finally ending with the ejector pumping system. However the model of the condensing spray chamber may be run independently of the rocket and ejector systems detail, with only appropriate mass flow boundary conditions placed at the entrance and exit of the condensing spray chamber model. The model of the condensing spray chamber takes into account droplet

  1. Engine Modelling for Control Applications

    DEFF Research Database (Denmark)

    Hendricks, Elbert

    1997-01-01

    In earlier work published by the author and co-authors, a dynamic engine model called a Mean Value Engine Model (MVEM) was developed. This model is physically based and is intended mainly for control applications. In its newer form, it is easy to fit to many different engines and requires little...... engine data for this purpose. It is especially well suited to embedded model applications in engine controllers, such as nonlinear observer based air/fuel ratio and advanced idle speed control. After a brief review of this model, it will be compared with other similar models which can be found...

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

    Science.gov (United States)

    Kudrin, O. I.

    1993-10-01

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

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

    Science.gov (United States)

    Weiss, Jonathan M.; Merkle, Charles L.

    1993-01-01

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

  4. Rocket University at KSC

    Science.gov (United States)

    Sullivan, Steven J.

    2014-01-01

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

  5. Modeling Liquid Rocket Engine Atomization and Swirl/Coaxial Injectors

    Science.gov (United States)

    2008-02-27

    High-Speed Jets, Computers and Fluids, V35, pp. 1033-1045, 2006. 11. Corpening, J. H., Heister, S. D, and Anderson, W.A., On the Thermal Decomposition...Additional study is required to further illuminate this issue. Finally, a series of BEM computations were conduted to assess the effect of perturbation

  6. A DYNAMIC MODEL FOR ROCKET LAUNCHER WITH COUPLED RIGID AND FLEXIBLW MOTION

    Institute of Scientific and Technical Information of China (English)

    ZHANG Ding-guo; XIAO Jian-qiang

    2005-01-01

    The dynamics of a coupled rigid-flexible rocket launcher is reported. The coupled rigid-flexible rocket launcher is divided into two subsystems, one is a system of rigid bodies,the other a flexible launch tube which can undergo large overall motions spatially. First, the mathematical models for these two subsystems were established respectively. Then the dynamic model for the whole system was obtained by considering the coupling effect between these two subsystems. The approach, which divides a complex system into several simple subsystems first and then obtains the dynamic model for the whole system via combining the existing dynamic models for simple subsystems, can make the modeling procedure efficient and convenient.

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

    Science.gov (United States)

    Weiss, Jonathan Mitchell

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

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

    Science.gov (United States)

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

    2013-01-01

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

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

    Science.gov (United States)

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

    2017-03-01

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

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

    Science.gov (United States)

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

    1972-01-01

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

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

    Science.gov (United States)

    Lohr, Jonathan C.; Trinh, Huu P.

    1992-01-01

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

  12. Rocket injector anomalies study. Volume 1: Description of the mathematical model and solution procedure

    Science.gov (United States)

    Przekwas, A. J.; Singhal, A. K.; Tam, L. T.

    1984-01-01

    The capability of simulating three dimensional two phase reactive flows with combustion in the liquid fuelled rocket engines is demonstrated. This was accomplished by modifying an existing three dimensional computer program (REFLAN3D) with Eulerian Lagrangian approach to simulate two phase spray flow, evaporation and combustion. The modified code is referred as REFLAN3D-SPRAY. The mathematical formulation of the fluid flow, heat transfer, combustion and two phase flow interaction of the numerical solution procedure, boundary conditions and their treatment are described.

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

    Directory of Open Access Journals (Sweden)

    V. S. Zarubin

    2016-01-01

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

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

    Directory of Open Access Journals (Sweden)

    Trushlyakov Valeriy I.

    2017-01-01

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

  15. Business Engineering Model

    OpenAIRE

    Österle, Hubert; Blessing, Dieter

    2000-01-01

    Business Engineering bedeutet systematische Entwicklung neuer Geschäftslösungen. Business Engineering zerlegt die Transformation von Unternehmen in beherrschbare Schritte, gibt Anleitung zur Bearbeitung dieser Schritte und verbindet diese in Vorgehensmodellen für Projekte. Der Aufsatz erklärt die Arbeitsweise des Business Engineering anhand eines durchgängigen Beispiels und beschreibt die Grundprinzipien dieser neuen Disziplin.

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

    Directory of Open Access Journals (Sweden)

    Tarakanova S.Y.

    2014-12-01

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

  17. An Otto Engine Dynamic Model

    OpenAIRE

    Florian Ion Tiberiu Petrescu; Relly Victoria Virgil Petrescu

    2016-01-01

    Otto engine dynamics are similar in almost all common internal combustion engines. We can speak so about dynamics of engines: Lenoir, Otto, and Diesel. The dynamic presented model is simple and original. The first thing necessary in the calculation of Otto engine dynamics, is to determine the inertial mass reduced at the piston. One uses then the Lagrange equation. Kinetic energy conservation shows angular speed variation (from the shaft) with inertial masses. One uses and elastic constant of...

  18. Engineering model for body armor

    NARCIS (Netherlands)

    Roebroeks, G.H.J.J.; Carton, E.P.

    2014-01-01

    TNO has developed an engineering model for flexible body armor, as one of their energy based engineering models that describe the physics of projectile to target interactions (weaves, metals, ceramics). These models form the basis for exploring the possibilities for protection improvement. This

  19. Mechanical Slosh Models for Rocket-Propelled Spacecraft

    Science.gov (United States)

    Jang, Jiann-Woei; Alaniz, Abram; Yang, Lee; Powers. Joseph; Hall, Charles

    2013-01-01

    Several analytical mechanical slosh models for a cylindrical tank with flat bottom are reviewed. Even though spacecrafts use cylinder shaped tanks, most of those tanks usually have elliptical domes. To extend the application of the analytical models for a cylindrical tank with elliptical domes, the modified slosh parameter models are proposed in this report by mapping an elliptical dome cylindrical tank to a flat top/bottom cylindrical tank while maintaining the equivalent liquid volume. For the low Bond number case, the low-g slosh models were also studied. Those low-g models can be used for Bond number > 10. The current low-g slosh models were also modified to extend their applications for the case that liquid height is smaller than the tank radius. All modified slosh models are implemented in MATLAB m-functions and are collected in the developed MST (Mechanical Slosh Toolbox).

  20. Space engineering modeling and optimization with case studies

    CERN Document Server

    Pintér, János

    2016-01-01

    This book presents a selection of advanced case studies that cover a substantial range of issues and real-world challenges and applications in space engineering. Vital mathematical modeling, optimization methodologies and numerical solution aspects of each application case study are presented in detail, with discussions of a range of advanced model development and solution techniques and tools. Space engineering challenges are discussed in the following contexts: •Advanced Space Vehicle Design •Computation of Optimal Low Thrust Transfers •Indirect Optimization of Spacecraft Trajectories •Resource-Constrained Scheduling, •Packing Problems in Space •Design of Complex Interplanetary Trajectories •Satellite Constellation Image Acquisition •Re-entry Test Vehicle Configuration Selection •Collision Risk Assessment on Perturbed Orbits •Optimal Robust Design of Hybrid Rocket Engines •Nonlinear Regression Analysis in Space Engineering< •Regression-Based Sensitivity Analysis and Robust Design ...

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

    Science.gov (United States)

    Pavli, A. J.

    1979-01-01

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

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

    Science.gov (United States)

    Pavli, A. J.

    1979-01-01

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

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

    Directory of Open Access Journals (Sweden)

    O. A. Vorozheeva

    2014-01-01

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

  4. Validation of a Solid Rocket Motor Internal Environment Model

    Science.gov (United States)

    Martin, Heath T.

    2017-01-01

    In a prior effort, a thermal/fluid model of the interior of Penn State University's laboratory-scale Insulation Test Motor (ITM) was constructed to predict both the convective and radiative heat transfer to the interior walls of the ITM with a minimum of empiricism. These predictions were then compared to values of total and radiative heat flux measured in a previous series of ITM test firings to assess the capabilities and shortcomings of the chosen modeling approach. Though the calculated fluxes reasonably agreed with those measured during testing, this exercise revealed means of improving the fidelity of the model to, in the case of the thermal radiation, enable direct comparison of the measured and calculated fluxes and, for the total heat flux, compute a value indicative of the average measured condition. By replacing the P1-Approximation with the discrete ordinates (DO) model for the solution of the gray radiative transfer equation, the radiation intensity field in the optically thin region near the radiometer is accurately estimated, allowing the thermal radiation flux to be calculated on the heat-flux sensor itself, which was then compared directly to the measured values. Though the fully coupling the wall thermal response with the flow model was not attempted due to the excessive computational time required, a separate wall thermal response model was used to better estimate the average temperature of the graphite surfaces upstream of the heat flux gauges and improve the accuracy of both the total and radiative heat flux computations. The success of this modeling approach increases confidence in the ability of state-of-the-art thermal and fluid modeling to accurately predict SRM internal environments, offers corrections to older methods, and supplies a tool for further studies of the dynamics of SRM interiors.

  5. Gas Turbine Engine Behavioral Modeling

    OpenAIRE

    Meyer, Richard T; DeCarlo, Raymond A.; Pekarek, Steve; Doktorcik, Chris

    2014-01-01

    This paper develops and validates a power flow behavioral model of a gas tur- bine engine with a gas generator and free power turbine. “Simple” mathematical expressions to describe the engine’s power flow are derived from an understand- ing of basic thermodynamic and mechanical interactions taking place within the engine. The engine behavioral model presented is suitable for developing a supervisory level controller of an electrical power system that contains the en- gine connected to a gener...

  6. High-Fidelity Gas and Granular Flow Physics Models for Rocket Exhaust Interaction with Lunar Soil Project

    Data.gov (United States)

    National Aeronautics and Space Administration — Current modeling of Lunar and Martian soil erosion and debris transport caused by rocket plume impingement lacks essential physics from the peculiar granular...

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

    Science.gov (United States)

    1972-01-01

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

  8. Analytical model for liquid rocket propellant feedline dynamics

    Science.gov (United States)

    Holster, J. L.; Astleford, W. J.

    1974-01-01

    A generalized analytical model and computer program have been developed to predict the frequency response of arbitrary liquid propellant feedline designs. The analytical model is based on an extension of an existing distributed parameter representation of a viscous fluid transmission line with laminar flow which was modified to include the effects of a turbulent mean flow. The effects of dissolved ullage gases, wall elasticity, localized gas or vapor bubbles, bellows, forced changes in length due to structural excitation, complex side branches, and structural mounting stiffness are also included. Each line component is written as a four-terminal, pressure-flow relationship in matrix form in the Laplace domain; the transfer function relating the pressure response at the line terminal (inducer inlet) to the external excitation is obtained in the computer program by sequential matrix substitution.

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

    Science.gov (United States)

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

    2010-01-01

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

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

    Science.gov (United States)

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

    2014-01-01

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

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

    Science.gov (United States)

    Liu, Chung-Chiun

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

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

    Institute of Scientific and Technical Information of China (English)

    黄延平; 高俊国

    2012-01-01

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

  13. Coupled simulation of CFD-flight-mechanics with a two-species-gas-model for the hot rocket staging

    Science.gov (United States)

    Li, Yi; Reimann, Bodo; Eggers, Thino

    2016-11-01

    The hot rocket staging is to separate the lowest stage by directly ignite the continuing-stage-motor. During the hot staging, the rocket stages move in a harsh dynamic environment. In this work, the hot staging dynamics of a multistage rocket is studied using the coupled simulation of Computational Fluid Dynamics and Flight Mechanics. Plume modeling is crucial for a coupled simulation with high fidelity. A 2-species-gas model is proposed to simulate the flow system of the rocket during the staging: the free-stream is modeled as "cold air" and the exhausted plume from the continuing-stage-motor is modeled with an equivalent calorically-perfect-gas that approximates the properties of the plume at the nozzle exit. This gas model can well comprise between the computation accuracy and efficiency. In the coupled simulations, the Navier-Stokes equations are time-accurately solved in moving system, with which the Flight Mechanics equations can be fully coupled. The Chimera mesh technique is utilized to deal with the relative motions of the separated stages. A few representative staging cases with different initial flight conditions of the rocket are studied with the coupled simulation. The torque led by the plume-induced-flow-separation at the aft-wall of the continuing-stage is captured during the staging, which can assist the design of the controller of the rocket. With the increasing of the initial angle-of-attack of the rocket, the staging quality becomes evidently poorer, but the separated stages are generally stable when the initial angle-of-attack of the rocket is small.

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

    Science.gov (United States)

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

    2008-01-01

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

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

    Science.gov (United States)

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

    2008-01-01

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

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

    Science.gov (United States)

    Kirk, Daniel R.

    2005-01-01

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

  17. Analysis and modeling of infrasound from a four-stage rocket launch.

    Science.gov (United States)

    Blom, Philip; Marcillo, Omar; Arrowsmith, Stephen

    2016-06-01

    Infrasound from a four-stage sounding rocket was recorded by several arrays within 100 km of the launch pad. Propagation modeling methods have been applied to the known trajectory to predict infrasonic signals at the ground in order to identify what information might be obtained from such observations. There is good agreement between modeled and observed back azimuths, and predicted arrival times for motor ignition signals match those observed. The signal due to the high-altitude stage ignition is found to be low amplitude, despite predictions of weak attenuation. This lack of signal is possibly due to inefficient aeroacoustic coupling in the rarefied upper atmosphere.

  18. ADDITIVE-MULTIPLICATIVE MODEL FOR RISK ESTIMATION IN THE PRODUCTION OF ROCKET AND SPACE TECHNICS

    Directory of Open Access Journals (Sweden)

    Orlov A. I.

    2014-10-01

    Full Text Available For the first time we have developed a general additive-multiplicative model of the risk estimation (to estimate the probabilities of risk events. In the two-level system in the lower level the risk estimates are combined additively, on the top – in a multiplicative way. Additive-multiplicative model was used for risk estimation for (1 implementation of innovative projects at universities (with external partners, (2 the production of new innovative products, (3 the projects for creation of rocket and space equipmen

  19. Computational Modeling in Tissue Engineering

    CERN Document Server

    2013-01-01

    One of the major challenges in tissue engineering is the translation of biological knowledge on complex cell and tissue behavior into a predictive and robust engineering process. Mastering this complexity is an essential step towards clinical applications of tissue engineering. This volume discusses computational modeling tools that allow studying the biological complexity in a more quantitative way. More specifically, computational tools can help in:  (i) quantifying and optimizing the tissue engineering product, e.g. by adapting scaffold design to optimize micro-environmental signals or by adapting selection criteria to improve homogeneity of the selected cell population; (ii) quantifying and optimizing the tissue engineering process, e.g. by adapting bioreactor design to improve quality and quantity of the final product; and (iii) assessing the influence of the in vivo environment on the behavior of the tissue engineering product, e.g. by investigating vascular ingrowth. The book presents examples of each...

  20. Model-based Software Engineering

    DEFF Research Database (Denmark)

    2010-01-01

    The vision of model-based software engineering is to make models the main focus of software development and to automatically generate software from these models. Part of that idea works already today. But, there are still difficulties when it comes to behaviour. Actually, there is no lack in models...

  1. Model-based Software Engineering

    DEFF Research Database (Denmark)

    2010-01-01

    The vision of model-based software engineering is to make models the main focus of software development and to automatically generate software from these models. Part of that idea works already today. But, there are still difficulties when it comes to behaviour. Actually, there is no lack in models...

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

    Science.gov (United States)

    Valenti, Elizabeth L.

    1992-10-01

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

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

    Science.gov (United States)

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

    2010-01-01

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

  4. High-speed schlieren imaging of rocket exhaust plumes

    Science.gov (United States)

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

    2016-11-01

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

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

    Science.gov (United States)

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

    2016-10-01

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

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

    Science.gov (United States)

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

    2015-06-01

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

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

    Science.gov (United States)

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

    1992-07-01

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

  8. Rocket Flight.

    Science.gov (United States)

    Van Evera, Bill; Sterling, Donna R.

    2002-01-01

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

  9. Rocket exhaust effluent modeling for tropospheric air quality and environmental assessments

    Science.gov (United States)

    Stephens, J. B.; Stewart, R. B.

    1977-01-01

    The various techniques for diffusion predictions to support air quality predictions and environmental assessments for aerospace applications are discussed in terms of limitations imposed by atmospheric data. This affords an introduction to the rationale behind the selection of the National Aeronautics and Space Administration (NASA)/Marshall Space Flight Center (MSFC) Rocket Exhaust Effluent Diffusion (REED) program. The models utilized in the NASA/MSFC REED program are explained. This program is then evaluated in terms of some results from a joint MSFC/Langley Research Center/Kennedy Space Center Titan Exhaust Effluent Prediction and Monitoring Program.

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

    Science.gov (United States)

    Brown, Andrew M.

    2014-01-01

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

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

    Science.gov (United States)

    Ashline, George; Ellis-Monaghan, Joanna

    2006-01-01

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

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

    Science.gov (United States)

    Ashline, George; Ellis-Monaghan, Joanna

    2006-01-01

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

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

    Science.gov (United States)

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

    1992-01-01

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

  14. Some typical solid propellant rocket motors

    NARCIS (Netherlands)

    Zandbergen, B.T.C.

    2013-01-01

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

  15. Some typical solid propellant rocket motors

    NARCIS (Netherlands)

    Zandbergen, B.T.C.

    2013-01-01

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

  16. Engine System Model Development for Nuclear Thermal Propulsion

    Science.gov (United States)

    Nelson, Karl W.; Simpson, Steven P.

    2006-01-01

    In order to design, analyze, and evaluate conceptual Nuclear Thermal Propulsion (NTP) engine systems, an improved NTP design and analysis tool has been developed. The NTP tool utilizes the Rocket Engine Transient Simulation (ROCETS) system tool and many of the routines from the Enabler reactor model found in Nuclear Engine System Simulation (NESS). Improved non-nuclear component models and an external shield model were added to the tool. With the addition of a nearly complete system reliability model, the tool will provide performance, sizing, and reliability data for NERVA-Derived NTP engine systems. A new detailed reactor model is also being developed and will replace Enabler. The new model will allow more flexibility in reactor geometry and include detailed thermal hydraulics and neutronics models. A description of the reactor, component, and reliability models is provided. Another key feature of the modeling process is the use of comprehensive spreadsheets for each engine case. The spreadsheets include individual worksheets for each subsystem with data, plots, and scaled figures, making the output very useful to each engineering discipline. Sample performance and sizing results with the Enabler reactor model are provided including sensitivities. Before selecting an engine design, all figures of merit must be considered including the overall impacts on the vehicle and mission. Evaluations based on key figures of merit of these results and results with the new reactor model will be performed. The impacts of clustering and external shielding will also be addressed. Over time, the reactor model will be upgraded to design and analyze other NTP concepts with CERMET and carbide fuel cores.

  17. Advanced Multi-Phase Flow CFD Model Development for Solid Rocket Motor Flowfield Analysis

    Science.gov (United States)

    Liaw, Paul; Chen, Y. S.; Shang, H. M.; Doran, Denise

    1993-01-01

    It is known that the simulations of solid rocket motor internal flow field with AL-based propellants require complex multi-phase turbulent flow model. The objective of this study is to develop an advanced particulate multi-phase flow model which includes the effects of particle dynamics, chemical reaction and hot gas flow turbulence. The inclusion of particle agglomeration, particle/gas reaction and mass transfer, particle collision, coalescence and breakup mechanisms in modeling the particle dynamics will allow the proposed model to realistically simulate the flowfield inside a solid rocket motor. The Finite Difference Navier-Stokes numerical code FDNS is used to simulate the steady-state multi-phase particulate flow field for a 3-zone 2-D axisymmetric ASRM model and a 6-zone 3-D ASRM model at launch conditions. The 2-D model includes aft-end cavity and submerged nozzle. The 3-D model represents the whole ASRM geometry, including additional grain port area in the gas cavity and two inhibitors. FDNS is a pressure based finite difference Navier-Stokes flow solver with time-accurate adaptive second-order upwind schemes, standard and extended k-epsilon models with compressibility corrections, multi zone body-fitted formulations, and turbulence particle interaction model. Eulerian/Lagrangian multi-phase solution method is applied for multi-zone mesh. To simulate the chemical reaction, penalty function corrected efficient finite-rate chemistry integration method is used in FDNS. For the AL particle combustion rate, the Hermsen correlation is employed. To simulate the turbulent dispersion of particles, the Gaussian probability distribution with standard deviation equal to (2k/3)(exp 1/2) is used for the random turbulent velocity components. The computational results reveal that the flow field near the juncture of aft-end cavity and the submerged nozzle is very complex. The effects of the turbulent particles affect the flow field significantly and provide better

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

    Science.gov (United States)

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

    1992-01-01

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

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

    Science.gov (United States)

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

    2017-01-01

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

  20. Conceptual Models for Search Engines

    Science.gov (United States)

    Hendry, D. G.; Efthimiadis, E. N.

    Search engines have entered popular culture. They touch people in diverse private and public settings and thus heighten the importance of such important social matters as information privacy and control, censorship, and equitable access. To fully benefit from search engines and to participate in debate about their merits, people necessarily appeal to their understandings for how they function. In this chapter we examine the conceptual understandings that people have of search engines by performing a content analysis on the sketches that 200 undergraduate and graduate students drew when asked to draw a sketch of how a search engine works. Analysis of the sketches reveals a diverse range of conceptual approaches, metaphors, representations, and misconceptions. On the whole, the conceptual models articulated by these students are simplistic. However, students with higher levels of academic achievement sketched more complete models. This research calls attention to the importance of improving students' technical knowledge of how search engines work so they can be better equipped to develop and advocate policies for how search engines should be embedded in, and restricted from, various private and public information settings.

  1. Fuel Chemistry And Combustion Distribution Effects On Rocket Engine Combustion Stability

    Science.gov (United States)

    2015-11-19

    model, the D 2 law rate constant for ideal combustion, k0, of a droplet is dependent on the thermal properties of the fuel and oxidizer and is...remaining increase in the D 2 law regression rate constant is caused by the non- ideal conditions of the experiment. Natural convection is present, as...is the gap-averaged pressure. Considering incompressible gas flow (ρ is constant ), two governing equations can be solved for the pressure in the gap

  2. Mean Value Engine Modelling of an SI Engine with EGR

    DEFF Research Database (Denmark)

    Føns, Michael; Müller, Martin; Chevalier, Alain

    1999-01-01

    Mean Value Engine Models (MVEMs) are simplified, dynamic engine models what are physically based. Such models are useful for control studies, for engine control system analysis and for model based engine control systems. Very few published MVEMs have included the effects of Exhaust Gas...... Recirculation (EGR). The purpose of this paper is to present a modified MVEM which includes EGR in a physical way. It has been tested using newly developed, very fast manifold pressure, manifold temperature, port and EGR mass flow sensors. Reasonable agreement has been obtained on an experimental engine...

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

    Science.gov (United States)

    Huff, Vearl N.; Kerrebrock, Jack

    1954-01-01

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

  4. Nonlinear theory of combustion stability in liquid rocket engine based on chemistry dynamics

    Institute of Scientific and Technical Information of China (English)

    黄玉辉; 王振国; 周进

    2002-01-01

    Detailed models of combustion instability based on chemistry dynamics are developed. The results show that large activation energy goes against the combustion stability. The heat transfer coefficient between the wall and the combust gas is an important bifurcation parameter for the combustion instability. The acoustics modes of the chamber are in competition and cooperation with each other for limited vibration energy. Thermodynamics criterion of combustion stability can be deduced from the nonlinear thermodynamics. Correlations of the theoretical results and historical experiments indicate that chemical kinetics play a critical role in the combustion instability.

  5. Stirling Engine Dynamic System Modeling

    Science.gov (United States)

    Nakis, Christopher G.

    2004-01-01

    The Thermo-Mechanical systems branch at the Glenn Research Center focuses a large amount time on Stirling engines. These engines will be used on missions where solar power is inefficient, especially in deep space. I work with Tim Regan and Ed Lewandowski who are currently developing and validating a mathematical model for the Stirling engines. This model incorporates all aspects of the system including, mechanical, electrical and thermodynamic components. Modeling is done through Simplorer, a program capable of running simulations of the model. Once created and then proven to be accurate, a model is used for developing new ideas for engine design. My largest specific project involves varying key parameters in the model and quantifying the results. This can all be done relatively trouble-free with the help of Simplorer. Once the model is complete, Simplorer will do all the necessary calculations. The more complicated part of this project is determining which parameters to vary. Finding key parameters depends on the potential for a value to be independently altered in the design. For example, a change in one dimension may lead to a proportional change to the rest of the model, and no real progress is made. Also, the ability for a changed value to have a substantial impact on the outputs of the system is important. Results will be condensed into graphs and tables with the purpose of better communication and understanding of the data. With the changing of these parameters, a more optimal design can be created without having to purchase or build any models. Also, hours and hours of results can be simulated in minutes. In the long run, using mathematical models can save time and money. Along with this project, I have many other smaller assignments throughout the summer. My main goal is to assist in the processes of model development, validation and testing.

  6. An Otto Engine Dynamic Model

    Directory of Open Access Journals (Sweden)

    Florian Ion Tiberiu Petrescu

    2016-03-01

    Full Text Available Otto engine dynamics are similar in almost all common internal combustion engines. We can speak so about dynamics of engines: Lenoir, Otto, and Diesel. The dynamic presented model is simple and original. The first thing necessary in the calculation of Otto engine dynamics, is to determine the inertial mass reduced at the piston. One uses then the Lagrange equation. Kinetic energy conservation shows angular speed variation (from the shaft with inertial masses. One uses and elastic constant of the crank shaft, k. Calculations should be made for an engine with a single cylinder. Finally it makes a dynamic analysis of the mechanism with discussion and conclusions. The ratio between the crank length r and the length of the connecting-rod l is noted with landa. When landa increases the mechanism dynamics is deteriorating. For a proper operation is necessary the reduction of the ratio landa, especially if we want to increase the engine speed. We can reduce the acceleration values by reducing the dimensions r and l.

  7. Multiscale modeling in food engineering

    NARCIS (Netherlands)

    Ho, Q.T.; Carmeliet, J.; Datta, A.K.; Defraeye, T.; Delele, M.A.; Herremans, E.; Opara, L.; Ramon, H.; Tijskens, E.; Sman, van der R.G.M.; Liedekerke, Van P.; Verboven, P.; Nicolai, B.M.

    2013-01-01

    Since many years food engineers have attempted to describe physical phenomena such as heat and mass transfer that occur in food during unit operations by means of mathematical models. Foods are hierarchically structured and have features that extend from the molecular scale to the food plant scale.

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

    Science.gov (United States)

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

    2014-01-01

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

  9. A modular ducted rocket missile model for threat and performance assessment

    NARCIS (Netherlands)

    Mayer, A.E.H.J.; Halswijk, W.H.C.; Komduur, H.J.; Lauzon, M.; Stowe, R.A.

    2005-01-01

    A model was developed to predict the thrust of throttled ramjet propelled missiles. The model is called DRCORE and fulfils the growing need to predict the performance of air breathing missiles. Each subsystem of the propulsion unit of this model is coded by using engineering formulae and enables the

  10. Model-Driven Useware Engineering

    Science.gov (United States)

    Meixner, Gerrit; Seissler, Marc; Breiner, Kai

    User-oriented hardware and software development relies on a systematic development process based on a comprehensive analysis focusing on the users' requirements and preferences. Such a development process calls for the integration of numerous disciplines, from psychology and ergonomics to computer sciences and mechanical engineering. Hence, a correspondingly interdisciplinary team must be equipped with suitable software tools to allow it to handle the complexity of a multimodal and multi-device user interface development approach. An abstract, model-based development approach seems to be adequate for handling this complexity. This approach comprises different levels of abstraction requiring adequate tool support. Thus, in this chapter, we present the current state of our model-based software tool chain. We introduce the use model as the core model of our model-based process, transformation processes, and a model-based architecture, and we present different software tools that provide support for creating and maintaining the models or performing the necessary model transformations.

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

    Institute of Scientific and Technical Information of China (English)

    窦唯

    2011-01-01

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

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

    Science.gov (United States)

    Macgregor, C.; Csomor, A.

    1974-01-01

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

  13. Coupled Solid Rocket Motor Ballistics and Trajectory Modeling for Higher Fidelity Launch Vehicle Design

    Science.gov (United States)

    Ables, Brett

    2014-01-01

    Multi-stage launch vehicles with solid rocket motors (SRMs) face design optimization challenges, especially when the mission scope changes frequently. Significant performance benefits can be realized if the solid rocket motors are optimized to the changing requirements. While SRMs represent a fixed performance at launch, rapid design iterations enable flexibility at design time, yielding significant performance gains. The streamlining and integration of SRM design and analysis can be achieved with improved analysis tools. While powerful and versatile, the Solid Performance Program (SPP) is not conducive to rapid design iteration. Performing a design iteration with SPP and a trajectory solver is a labor intensive process. To enable a better workflow, SPP, the Program to Optimize Simulated Trajectories (POST), and the interfaces between them have been improved and automated, and a graphical user interface (GUI) has been developed. The GUI enables real-time visual feedback of grain and nozzle design inputs, enforces parameter dependencies, removes redundancies, and simplifies manipulation of SPP and POST's numerous options. Automating the analysis also simplifies batch analyses and trade studies. Finally, the GUI provides post-processing, visualization, and comparison of results. Wrapping legacy high-fidelity analysis codes with modern software provides the improved interface necessary to enable rapid coupled SRM ballistics and vehicle trajectory analysis. Low cost trade studies demonstrate the sensitivities of flight performance metrics to propulsion characteristics. Incorporating high fidelity analysis from SPP into vehicle design reduces performance margins and improves reliability. By flying an SRM designed with the same assumptions as the rest of the vehicle, accurate comparisons can be made between competing architectures. In summary, this flexible workflow is a critical component to designing a versatile launch vehicle model that can accommodate a volatile

  14. Solid propellant rocket motor

    Science.gov (United States)

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

    1973-01-01

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

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

    Science.gov (United States)

    Dawson, Joshua

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

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

    Science.gov (United States)

    2013-12-01

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

  17. Engineered Swine Models of Cancer

    Directory of Open Access Journals (Sweden)

    Adrienne L. Watson

    2016-05-01

    Full Text Available Over the past decade, the technology to engineer genetically modified swine has seen many advancements, and because their physiology is remarkably similar to that of humans, swine models of cancer may be extremely valuable for preclinical safety studies as well as toxicity testing of pharmaceuticals prior to the start of human clinical trials. Hence, the benefits of using swine as a large animal model in cancer research and the potential applications and future opportunities of utilizing pigs in cancer modeling are immense. In this review, we discuss how pigs have been and can be used as a biomedical models for cancer research, with an emphasis on current technologies. We have focused on applications of precision genetics that can provide models that mimic human cancer predisposition syndromes. In particular, we describe the advantages of targeted gene-editing using custom endonucleases, specifically TALENs and CRISPRs, and transposon systems, to make novel pig models of cancer with broad preclinical applications.

  18. Fundamentals of aircraft and rocket propulsion

    CERN Document Server

    El-Sayed, Ahmed F

    2016-01-01

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

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

    Institute of Scientific and Technical Information of China (English)

    王永青; 刘海波; 李护林; 贾振元

    2012-01-01

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

  20. Engineering graphic modelling a workbook for design engineers

    CERN Document Server

    Tjalve, E; Frackmann Schmidt, F

    2013-01-01

    Engineering Graphic Modelling: A Practical Guide to Drawing and Design covers how engineering drawing relates to the design activity. The book describes modeled properties, such as the function, structure, form, material, dimension, and surface, as well as the coordinates, symbols, and types of projection of the drawing code. The text provides drawing techniques, such as freehand sketching, bold freehand drawing, drawing with a straightedge, a draughting machine or a plotter, and use of templates, and then describes the types of drawing. Graphic designers, design engineers, mechanical engine

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

    Science.gov (United States)

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

    1959-01-01

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

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

    Institute of Scientific and Technical Information of China (English)

    段祥军

    2014-01-01

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

  3. Nuclear Thermal Rocket Simulation in NPSS

    Science.gov (United States)

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

    2013-01-01

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

  4. Simulation of 3-D Ultraviolet Radiation from Liquid Rocket Engine Plume%液体火箭发动机羽烟三维紫外辐射仿真研究

    Institute of Scientific and Technical Information of China (English)

    国爱燕; 唐义; 白廷柱; 黄刚

    2012-01-01

    针对液体火箭发动机羽烟紫外辐射的空间分布问题,建立三维数值计算模型.该模型采用标准κ-ε湍流模型和PDF模型仿真羽烟流场的状态参数,根据HITRAN数据库计算流场内吸收系数分布,并利用离散坐标法求解辐射传输方程,计算三维空间的紫外辐射分布.测试结果表明:三维紫外辐射模型的计算结果与实验数据一致,能够反应不同视角下液体火箭发动机羽烟紫外辐射强度的空间变化.%To analyze the spacial distribution of UV radiation from liquid rocket engine plume, a 3-D numerical model has been built. The standard κ -e model and probability density function (PDF) model were adopted to compute the flow-field properties of the plume. Combined with HITRAN database, the distribution of absorption coefficient was calculated, and the radiation transfer equation could be solved by the discrete ordinate method. Test results show that the 3-D computation model can provide numerical data that agree well with measured experimental data. It also could reflect the spacial distribution of UV radiation from the hydrogen-oxygen rocket engine plume at different angles of view.

  5. Scaled Rocket Testing in Hypersonic Flow

    Science.gov (United States)

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

    2015-01-01

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

  6. Slip-model Performance for Underexpanded Micro-scale Rocket Nozzle Flows

    Institute of Scientific and Technical Information of China (English)

    José A. Morí(n)igo; José Hermida Quesada; Francisco Caballero Requena

    2007-01-01

    In aerospace Micro-ElectroMechanical Systems (MEMS), the characteristic length scale of the flow approaches the molecular mean free path, thus invalidating the continuum description and enforcing the use of particle methods, like the Direct Simulation Monte Carlo (DSMC), to deal with the non-equilibrium regions. Within the slip-regime (0.01<Kn<~0.1) both approaches, continuum and particle-based, seem to behave well in terms of accuracy. The present study summarizes the implementation and results obtained with a 2nd-order slip boundary condition in a Navier-Stokes solver to address the rarefaction near the nozzle walls. Its assessment and application to a cold-gas micro-scale conical nozzle of 300μm throat diameter, discharging into the low-pressure freestream,constitutes the major aim of the work. The slip-model incorporates the velocity slip with thermal creep and temperature jump, thus permitting to deal with non-isothermal flows as well. Results show that the gas experiences an intense rarefaction in the lip vicinity, pointing to the limits of model validity. Furthermore, a strong Mach deceleration is observed, attributed to the rather thick subsonic boundary layer and supersonic bulk heating caused by the viscous dissipation, in contrast with the expansion to occur in large rocket nozzles during underexpanded operation.

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

    CERN Document Server

    Fargion, D

    2007-01-01

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

  8. Rocket noise - A review

    Science.gov (United States)

    McInerny, S. A.

    1990-10-01

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

  9. Modeling the internal combustion engine

    Science.gov (United States)

    Zeleznik, F. J.; Mcbride, B. J.

    1985-01-01

    A flexible and computationally economical model of the internal combustion engine was developed for use on large digital computer systems. It is based on a system of ordinary differential equations for cylinder-averaged properties. The computer program is capable of multicycle calculations, with some parameters varying from cycle to cycle, and has restart capabilities. It can accommodate a broad spectrum of reactants, permits changes in physical properties, and offers a wide selection of alternative modeling functions without any reprogramming. It readily adapts to the amount of information available in a particular case because the model is in fact a hierarchy of five models. The models range from a simple model requiring only thermodynamic properties to a complex model demanding full combustion kinetics, transport properties, and poppet valve flow characteristics. Among its many features the model includes heat transfer, valve timing, supercharging, motoring, finite burning rates, cycle-to-cycle variations in air-fuel ratio, humid air, residual and recirculated exhaust gas, and full combustion kinetics.

  10. Numerically Modeling the Erosion of Lunar Soil by Rocket Exhaust Plumes

    Science.gov (United States)

    2008-01-01

    In preparation for the Apollo program, Leonard Roberts of the NASA Langley Research Center developed a remarkable analytical theory that predicts the blowing of lunar soil and dust beneath a rocket exhaust plume. Roberts assumed that the erosion rate was determined by the excess shear stress in the gas (the amount of shear stress greater than what causes grains to roll). The acceleration of particles to their final velocity in the gas consumes a portion of the shear stress. The erosion rate continues to increase until the excess shear stress is exactly consumed, thus determining the erosion rate. Roberts calculated the largest and smallest particles that could be eroded based on forces at the particle scale, but the erosion rate equation assumed that only one particle size existed in the soil. He assumed that particle ejection angles were determined entirely by the shape of the terrain, which acts like a ballistic ramp, with the particle aerodynamics being negligible. The predicted erosion rate and the upper limit of particle size appeared to be within an order of magnitude of small-scale terrestrial experiments but could not be tested more quantitatively at the time. The lower limit of particle size and the predictions of ejection angle were not tested. We observed in the Apollo landing videos that the ejection angles of particles streaming out from individual craters were time-varying and correlated to the Lunar Module thrust, thus implying that particle aerodynamics dominate. We modified Roberts theory in two ways. First, we used ad hoc the ejection angles measured in the Apollo landing videos, in lieu of developing a more sophisticated method. Second, we integrated Roberts equations over the lunar-particle size distribution and obtained a compact expression that could be implemented in a numerical code. We also added a material damage model that predicts the number and size of divots which the impinging particles will cause in hardware surrounding the landing

  11. Reusable Solid Rocket Motor - V(RSRMV)Nozzle Forward Nose Ring Thermo-Structural Modeling

    Science.gov (United States)

    Clayton, J. Louie

    2012-01-01

    During the developmental static fire program for NASAs Reusable Solid Rocket Motor-V (RSRMV), an anomalous erosion condition appeared on the nozzle Carbon Cloth Phenolic nose ring that had not been observed in the space shuttle RSRM program. There were regions of augmented erosion located on the bottom of the forward nose ring (FNR) that measured nine tenths of an inch deeper than the surrounding material. Estimates of heating conditions for the RSRMV nozzle based on limited char and erosion data indicate that the total heat loading into the FNR, for the new five segment motor, is about 40-50% higher than the baseline shuttle RSRM nozzle FNR. Fault tree analysis of the augmented erosion condition has lead to a focus on a thermomechanical response of the material that is outside the existing experience base of shuttle CCP materials for this application. This paper provides a sensitivity study of the CCP material thermo-structural response subject to the design constraints and heating conditions unique to the RSRMV Forward Nose Ring application. Modeling techniques are based on 1-D thermal and porous media calculations where in-depth interlaminar loading conditions are calculated and compared to known capabilities at elevated temperatures. Parameters such as heat rate, in-depth pressures and temperature, degree of char, associated with initiation of the mechanical removal process are quantified and compared to a baseline thermo-chemical material removal mode. Conclusions regarding postulated material loss mechanisms are offered.

  12. The Effect of Resistance on Rocket Injector Acoustics

    Science.gov (United States)

    Morgan, C. J.

    2015-01-01

    Combustion instability, where unsteady heat release couples with acoustic modes, has long been an area of concern in liquid rocket engines. Accurate modeling of the acoustic normal modes of the combustion chamber is important to understanding and preventing combustion instability. The injector resistance can have a significant influence on the chamber normal mode shape, and hence on the system stability.

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

    Directory of Open Access Journals (Sweden)

    A. V. Zolin

    2015-01-01

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

  14. Computational simulation of liquid rocket injector anomalies

    Science.gov (United States)

    Przekwas, A. J.; Singhal, A. K.; Tam, L. T.; Davidian, K.

    1986-01-01

    A computer model has been developed to analyze the three-dimensional two-phase reactive flows in liquid fueled rocket combustors. The model is designed to study the influence of liquid propellant injection nonuniformities on the flow pattern, combustion and heat transfer within the combustor. The Eulerian-Lagrangian approach for simulating polidisperse spray flow, evaporation and combustion has been used. Full coupling between the phases is accounted for. A nonorthogonal, body fitted coordinate system along with a conservative control volume formulation is employed. The physical models built into the model include a kappa-epsilon turbulence model, a two-step chemical reaction, and the six-flux radiation model. Semiempirical models are used to describe all interphase coupling terms as well as chemical reaction rates. The purpose of this study was to demonstrate an analytical capability to predict the effects of reactant injection nonuniformities (injection anomalies) on combustion and heat transfer within the rocket combustion chamber. The results show promising application of the model to comprehensive modeling of liquid propellant rocket engines.

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

    Science.gov (United States)

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

    2000-01-01

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

  16. 轨姿控液体火箭发动机水击仿真模拟%Simulation of water hammer in liquid rocket engine of orbit and attitude control system

    Institute of Scientific and Technical Information of China (English)

    张峥岳; 康乃全

    2012-01-01

    Taking the liquid rocket engine of orbit and attitude control system as the study object, an emulator was established with AMESim according to the modular modeling idea. The simulation computation of water hammer pressure in the pipeline while the engine system was working was per- formed. The results show that the running of orbit control engine is a major factor creating high water hammer. The compared result of theoretical calculation and test data indicate that the simulation mod- els can give reasonable descriptions for generative process of water hammer. The measure to reduce the amount of water hammer is introduced.%以轨姿控液体火箭发动机为研究对象,根据模块化思想,利用AMESim建立了仿真平台,仿真计算了发动机系统工作中管路的水击压力。结果表明:轨控发动机的工作是引起大水击的主要因素。通过与理论计算和试验数据的对比表明,仿真模型较好地描述了管路水击的生成过程。介绍了减小系统水击量的措施。

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

    Science.gov (United States)

    1973-01-01

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

  18. Rocket Flight Path

    Directory of Open Access Journals (Sweden)

    Jamie Waters

    2014-09-01

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

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

    Science.gov (United States)

    Bellman, Donald R; Humphrey, Jack C

    1948-01-01

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

  20. Model-Based Systems Engineering in Concurrent Engineering Centers

    Science.gov (United States)

    Iwata, Curtis; Infeld, Samantha; Bracken, Jennifer Medlin; McGuire, Melissa; McQuirk, Christina; Kisdi, Aron; Murphy, Jonathan; Cole, Bjorn; Zarifian, Pezhman

    2015-01-01

    Concurrent Engineering Centers (CECs) are specialized facilities with a goal of generating and maturing engineering designs by enabling rapid design iterations. This is accomplished by co-locating a team of experts (either physically or virtually) in a room with a narrow design goal and a limited timeline of a week or less. The systems engineer uses a model of the system to capture the relevant interfaces and manage the overall architecture. A single model that integrates other design information and modeling allows the entire team to visualize the concurrent activity and identify conflicts more efficiently, potentially resulting in a systems model that will continue to be used throughout the project lifecycle. Performing systems engineering using such a system model is the definition of model-based systems engineering (MBSE); therefore, CECs evolving their approach to incorporate advances in MBSE are more successful in reducing time and cost needed to meet study goals. This paper surveys space mission CECs that are in the middle of this evolution, and the authors share their experiences in order to promote discussion within the community.

  1. Transforming System Engineering through Model-Centric Engineering

    Science.gov (United States)

    2015-01-31

    can be addressed through “engineering,” and NAVAIR is making some headway on this item. As for item #3, this topic relates to a question posed by...through “engineering,” and NAVAIR is making some headway on this item. The third topic relates to a question posed by our sponsor after our review of the...ACAT Acquisition Category AFT Architecture Framework Tool of NASA/JPL AGI Analytical Graphics, Inc. AGM Acquisition Guidance Model ANSI American

  2. Transforming System Engineering through Model-Centric Engineering

    Science.gov (United States)

    2015-11-18

    Resilient System Conceptual Representation of Environment [63] - Enhanced.... 26 Figure 13. Measurement Collection Instrument...Model- Based Enterprise [81], which brings in more focus on manufacturability. The concept characterized as Digital Thread2 envisions a frameworks ...that merges physics- based models generated by the discipline engineers during the detailed design process with MBSE’s conceptual and top-level

  3. Modeling of the filling and cooling processes of hot fuel mains in Liquid Fuel Rocket Power Plant (LFRPP)

    Science.gov (United States)

    Prisnyakov, V. F.; Pokrishkin, V. V.; Serebryansky, V. N.

    A mathematical model of heat and mass exchange processes during filling and cooling of hot fuel mains of the Liquid Fuel Rocket Power Plant (LFRPP), which allows to define a mass consumption and distribution of two-phase flow parameters by the length of pipeline. Results of calculations are compared with experimental data, taken during filling of the main with a supply of liquid oxygen from the tank into the combustion chamber. Also, the results of modeling of hydrogen main dynamic characteristics of LFRPP in the same conditions are given.

  4. State Estimation for the VASIMR Plasma Engine

    OpenAIRE

    2008-01-01

    This paper presents work on the application of virtual metrology techniques to the VAriable Specific Impulse Magnetoplasma Rocket (VASMIR) engine. The work concentrates on the estimation of internal temperatures of the rocket using state space models and Optical Emission Spectroscopy (OES). These estimations are useful as direct thermal measurements will not be available in the final system design.

  5. Rocket Tablet,

    Science.gov (United States)

    1984-09-12

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

  6. Mean Value Modelling of Turbocharged SI Engines

    DEFF Research Database (Denmark)

    Müller, Martin; Hendricks, Elbert; Sorenson, Spencer C.

    1998-01-01

    The development of a computer simulation to predict the performance of a turbocharged spark ignition engine during transient operation. New models have been developed for the turbocharged and the intercooling system. An adiabatic model for the intake manifold is presented....

  7. Rocket launchers as passive controllers

    Science.gov (United States)

    Cochran, J. E., Jr.; Gunnels, R. T.; McCutchen, R. K., Jr.

    1981-12-01

    A concept is advanced for using the motion of launchers of a free-flight launcher/rocket system which is caused by random imperfections of the rockets launched from it to reduce the total error caused by the imperfections. This concept is called 'passive launcher control' because no feedback is generated by an active energy source after an error is sensed; only the feedback inherent in the launcher/rocket interaction is used. Relatively simple launcher models with two degrees of freedom, pitch and yaw, were used in conjunction with a more detailed, variable-mass model in a digital simulation code to obtain rocket trajectories with and without thrust misalignment and dynamic imbalance. Angular deviations of rocket velocities and linear deviations of the positions of rocket centers of mass at burnout were computed for cases in which the launcher was allowed to move ('flexible' launcher) and was constrained so that it did not rotate ('rigid' launcher) and ratios of flexible to rigid deviations were determined. Curves of these error ratios versus launcher frequency are presented. These show that a launcher which has a transverse moment of inertia about its pivot point of the same magnitude as that of the centroidal transverse moments of inertia of the rockets launched from it can be tuned to passively reduce the errors caused by rocket imperfections.

  8. This "Is" Rocket Science!

    Science.gov (United States)

    Keith, Wayne; Martin, Cynthia; Veltkamp, Pamela

    2013-01-01

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

  9. Final Progress Report for the NASA Inductrack Model Rocket Launcher at the Lawrence Livermore National Laboratory

    Energy Technology Data Exchange (ETDEWEB)

    Tung, L S; Post, R F; Martinez-Frias, J

    2001-06-27

    The Inductrack magnetic levitation system, developed at the Lawrence Livermore National Laboratory, was studied for its possible use for launching rockets. Under NASA sponsorship, a small model system was constructed at the Laboratory to pursue key technical aspects of this proposed application. The Inductrack is a passive magnetic levitation system employing special arrays of high-field permanent magnets (Halbach arrays) on the levitating cradle, moving above a ''track'' consisting of a close-packed array of shorted coils with which are interleaved with special drive coils. Halbach arrays produce a strong spatially periodic magnetic field on the front surface of the arrays, while canceling the field on their back surface. Relative motion between the Halbach arrays and the track coils induces currents in those coils. These currents levitate the cradle by interacting with the horizontal component of the magnetic field. Pulsed currents in the drive coils, synchronized with the motion of the carrier, interact with the vertical component of the magnetic field to provide acceleration forces. Motional stability, including resistance to both vertical and lateral aerodynamic forces, is provided by having Halbach arrays that interact with both the upper and the lower sides of the track coils. At present, a 7.8 meter track composed of drive and levitation coils has been built and the electronic drive circuitry performs as designed. A 9 kg cradle that carries the Halbach array of permanent magnets has been built. A mechanical launcher is nearly complete which will provide an initial cradle velocity of 9 m/s into the electronic drive section. We have found that the drag forces from the levitation coils were higher than in our original design. However, measurements of drag force at velocities less than 1 m/s are exactly as predicted by theory. Provided here are recommended design changes to improve the track's performance so that a final velocity of 40

  10. Chemical Kinetic Models for Advanced Engine Combustion

    Energy Technology Data Exchange (ETDEWEB)

    Pitz, William J. [Lawrence Livermore National Lab. (LLNL), Livermore, CA (United States); Mehl, Marco [Lawrence Livermore National Lab. (LLNL), Livermore, CA (United States); Westbrook, Charles K. [Lawrence Livermore National Lab. (LLNL), Livermore, CA (United States)

    2014-10-22

    The objectives for this project are as follows: Develop detailed chemical kinetic models for fuel components used in surrogate fuels for compression ignition (CI), homogeneous charge compression ignition (HCCI) and reactivity-controlled compression-ignition (RCCI) engines; and Combine component models into surrogate fuel models to represent real transportation fuels. Use them to model low-temperature combustion strategies in HCCI, RCCI, and CI engines that lead to low emissions and high efficiency.

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

    Directory of Open Access Journals (Sweden)

    Victor Fernandez-Villace

    2013-02-01

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

  12. Hybrid Rocket Technology

    Directory of Open Access Journals (Sweden)

    Sankaran Venugopal

    2011-04-01

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

  13. Optimization in engineering models and algorithms

    CERN Document Server

    Sioshansi, Ramteen

    2017-01-01

    This textbook covers the fundamentals of optimization, including linear, mixed-integer linear, nonlinear, and dynamic optimization techniques, with a clear engineering focus. It carefully describes classical optimization models and algorithms using an engineering problem-solving perspective, and emphasizes modeling issues using many real-world examples related to a variety of application areas. Providing an appropriate blend of practical applications and optimization theory makes the text useful to both practitioners and students, and gives the reader a good sense of the power of optimization and the potential difficulties in applying optimization to modeling real-world systems. The book is intended for undergraduate and graduate-level teaching in industrial engineering and other engineering specialties. It is also of use to industry practitioners, due to the inclusion of real-world applications, opening the door to advanced courses on both modeling and algorithm development within the industrial engineering ...

  14. Mathematical model of the Amazon Stirling engine

    Energy Technology Data Exchange (ETDEWEB)

    Vidal Medina, Juan Ricardo [Universidad Autonoma de Occidente (Colombia)], e-mail: jrvidal@uao.edu.co; Cobasa, Vladimir Melian; Silva, Electo [Universidade Federal de Itajuba, MG (Brazil)], e-mail: vlad@unifei.edu.br

    2010-07-01

    The Excellency Group in Thermoelectric and Distributed Generation (NEST, for its acronym in Portuguese) at the Federal University of Itajuba, has designed a Stirling engine prototype to provide electricity to isolated regions of Brazil. The engine was designed to operate with residual biomass from timber process. This paper presents mathematical models of heat exchangers (hot, cold and regenerator) integrated into second order adiabatic models. The general model takes into account the pressure drop losses, hysteresis and internal losses. The results of power output, engine efficiency, optimal velocity of the exhaust gases and the influence of dead volume in engine efficiency are presented in this paper. The objective of this modeling is to propose improvements to the manufactured engine design. (author)

  15. 液氧煤油发动机稳态故障仿真分析%Steady-state fault simulation and effect analysis of LOX/kerosene rocket engine

    Institute of Scientific and Technical Information of China (English)

    党锋刚; 马红宇; 李春红; 宋春

    2012-01-01

    根据液氧煤油补燃循环发动机的特点,建立了稳态工作过程故障仿真数学模型,并针对比较典型的几种故障模式,进行了仿真计算与效应分析,最后进行了故障参数特征的初步提取。结果表明,选定的10个缓变热力参数,可对泄漏、堵塞及涡轮泵等典型故障模式进行有效识别和分离。%The simulation mathematical model of fault occurring in the steady-state working process is built and simulation software is designed according to the characteristics of LOX/kerosene staged combustion cycled rocket engine.Several typical failure modes are simulatively calculated.The effect of the fault is analyzed and characteristics of the fault modes are extracted.Ten slow variable thermal parameters were selected on the basis of the anaysis result to identify the faut modes of leakage and jamming in the fuel pipe of engine.

  16. Laser docking sensor engineering model

    Science.gov (United States)

    Dekome, Kent; Barr, Joseph M.

    NASA JSC has been involved in the development of Laser sensors for the past ten years in order to support future rendezvous and docking missions, both manned and unmanned. Although many candidate technologies have been breadboarded and evaluated, no sensor hardware designed specifically for rendezvous and docking applications has been demonstrated on-orbit. It has become apparent that representative sensors need to be flown and demonstrated as soon as possible, with minimal cost, to provide the capability of the technology in meeting NASA's future AR&C applications. Technology and commercial component reliability have progressed to where it is now feasible to fly hardware as a detailed test objective minimizing the overall cost and development time. This presentation will discuss the ongoing effort to convert an existing in-house developed breadboard to an engineering model configuration suitable for flight. The modifications include improving the ranger resolution and stability with an in-house design, replacing the rack mounted galvanometric scanner drivers with STD-bus cards, replacing the system controlling personal computer with a microcontroller, and repackaging the subsystems as appropriate. The sensor will use the performance parameters defined in previous JSC requirements working groups as design goals and be built to withstand the space environment where fiscally feasible. Testing of the in-house ranger design is expected to be completed in October. The results will be included in the presentation. Preliminary testing of the ranging circuitry indicates a range resolution of 4mm is possible. The sensor will be mounted in the payload bay on a shelf bracket and have command, control, and display capabilities using the payload general support computer via an RS422 data line.

  17. Nuclear-thermal rocket thrust transient effects on minimum-fuel lunar trajectories

    Science.gov (United States)

    Rivas, Matthew L.

    1995-01-01

    A technically viable option for low-cost minimum-fuel Lunar transfers with short trip times is the use of nuclear thermal rockets. However, little work has been done on the effects the associated thrust transients have on these optimal trajectories. The nominal thrust level of an engine is not immediately reached when the rocket is turned ``on.'' Similarly, when the engine is turned ``off'', the thrust and specific impulse levels decrease over a period of time which is directly related to both the flow effecs of the engine and cooling requirements. This paper presents an analysis of these effects on a typical optimal Lunar transfer. Several different models simulating the transient effects are used. They range from simple ``mass dumps'' to account for the extra required propellant to curve-fits of actual engine characteristics obtained from the NERVA nuclear rocket program.

  18. 基于工作流的液体火箭发动机虚拟试验流程管理%Workflow-based process management in virtual test of liquid rocket engine

    Institute of Scientific and Technical Information of China (English)

    段娜; 朱子环; 于海磊; 张黎辉

    2012-01-01

    以液体火箭发动机虚拟试验为对象,研究了数字化试验流程管理的解决方案。通过分析发动机虚拟试验流程,建立了基于工作流的试验管理系统过程仿真模型,提出了一个支撑整个试验管理系统的层次化体系架构。该架构为整个虚拟试验过程提供了统一的试验信息集成平台和应用服务环境。最后,以试验准备阶段为例,给出了该系统的一个集成应用,介绍了基于图形管理系统的实现过程,得到了过程中资源消耗优化的结论。所提出的液体火箭发动机虚拟试验流程管理体系为真实试验的资源优化提供了理论依据,在航天领域的数字化试验方面进行了探索性研究。%The digitalized test process management in the virtual test of liquid rocket engines is studied. By analyzing the virtual test process of such engines, the workflow-based process simulation model of test management system is established, and a hierarchical architecture is proposed to support the test process management system. The architecture provides a unified test information integrated platform and an application service environment for the entire virtual test process. Taking preparatory stage of the test as an example, the integrated application of the the management system is shown, and the implementation process based on graphics management system is introduced. A conclusion of optimized consumption of resources is achieved. The test process management system presented in this paper for the virtual test of liquid rocket engines can provide a theoretic foundation for resource utilization in real test. It conduces to farther research in digitalized test process management in the field of space technology.

  19. Model based development of engine control algorithms

    NARCIS (Netherlands)

    Dekker, H.J.; Sturm, W.L.

    1996-01-01

    Model based development of engine control systems has several advantages. The development time and costs are strongly reduced because much of the development and optimization work is carried out by simulating both engine and control system. After optimizing the control algorithm it can be executed b

  20. Review: Modeling Damping in Mechanical Engineering Structures

    Directory of Open Access Journals (Sweden)

    Michel Lalanne

    2000-01-01

    Full Text Available This paper is concerned with the introduction of damping effects in the analysis of mechanical engineering structures. Damping can be considered as being generated by concentrated elements, by distributed elements, or by several effects existing simultaneously. Modeling damping for different engineering situations is described and some applications are presented briefly.

  1. 氢氧火箭发动机射流仿真与试验台热防护%Thermal Protection and Plume Simulation for Hydrogen/Oxygen Rocket Engine Test Stage

    Institute of Scientific and Technical Information of China (English)

    李茂; 王占林

    2014-01-01

    The plume characteristics of the hydrogen/oxygen rocket engine were studied with the method of CFD ( computational fluid dynamics ) .The influences of the geometry model , the combus-tion model and the turbulence model on the characteristics of the combustion flowfield were ana-lyzed .The results from the numerical simulation were compared with those from the experiments qualitatively and quantitatively .Based on the distributions of the temperature from the numerical simulation results in the different tests , the armor plate was designed for the thermal protection and applied for the test stage .It is confirmed that the approach is effective and reliable .%采用计算流体方法获得氢氧火箭发动机地面试验射流特征,开展了几何模型、燃烧模型和湍流模型对射流场的影响分析以及与试验结果的定性和定量对比。依据不同试验模式下的射流场温度的数值分布,提出试验台钢板防护方案并进行防护,试验证明方案可靠有效。

  2. Simple-1: Development stage of the data transmission system for a solid propellant mid-power rocket model

    Science.gov (United States)

    Yarce, Andrés; Sebastián Rodríguez, Juan; Galvez, Julián; Gómez, Alejandro; García, Manuel J.

    2017-06-01

    This paper presents the development stage of a communication module for a solid propellant mid-power rocket model. The communication module was named. Simple-1 and this work considers its design, construction and testing. A rocket model Estes Ventris Series Pro II® was modified to introduce, on the top of the payload, several sensors in a CanSat form factor. The Printed Circuit Board (PCB) was designed and fabricated from Commercial Off The Shelf (COTS) components and assembled in a cylindrical rack structure similar to this small format satellite concept. The sensors data was processed using one Arduino Mini and transmitted using a radio module to a Software Defined Radio (SDR) HackRF based platform on the ground station. The Simple-1 was tested using a drone in successive releases, reaching altitudes from 200 to 300 meters. Different kind of data, in terms of altitude, position, atmospheric pressure and vehicle temperature were successfully measured, making possible the progress to a next stage of launching and analysis.

  3. Rocket + Science = Dialogue

    Science.gov (United States)

    Morris,Bruce; Sullivan, Greg; Burkey, Martin

    2010-01-01

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

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

    Science.gov (United States)

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

    2013-01-01

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

  5. Distributed simulation a model driven engineering approach

    CERN Document Server

    Topçu, Okan; Oğuztüzün, Halit; Yilmaz, Levent

    2016-01-01

    Backed by substantive case studies, the novel approach to software engineering for distributed simulation outlined in this text demonstrates the potent synergies between model-driven techniques, simulation, intelligent agents, and computer systems development.

  6. Numerical methods and modelling for engineering

    CERN Document Server

    Khoury, Richard

    2016-01-01

    This textbook provides a step-by-step approach to numerical methods in engineering modelling. The authors provide a consistent treatment of the topic, from the ground up, to reinforce for students that numerical methods are a set of mathematical modelling tools which allow engineers to represent real-world systems and compute features of these systems with a predictable error rate. Each method presented addresses a specific type of problem, namely root-finding, optimization, integral, derivative, initial value problem, or boundary value problem, and each one encompasses a set of algorithms to solve the problem given some information and to a known error bound. The authors demonstrate that after developing a proper model and understanding of the engineering situation they are working on, engineers can break down a model into a set of specific mathematical problems, and then implement the appropriate numerical methods to solve these problems. Uses a “building-block” approach, starting with simpler mathemati...

  7. Mathematical Model For Engineering Analysis And Optimization

    Science.gov (United States)

    Sobieski, Jaroslaw

    1992-01-01

    Computational support for engineering design process reveals behavior of designed system in response to external stimuli; and finds out how behavior modified by changing physical attributes of system. System-sensitivity analysis combined with extrapolation forms model of design complementary to model of behavior, capable of direct simulation of effects of changes in design variables. Algorithms developed for this method applicable to design of large engineering systems, especially those consisting of several subsystems involving many disciplines.

  8. Mathematical Model For Engineering Analysis And Optimization

    Science.gov (United States)

    Sobieski, Jaroslaw

    1992-01-01

    Computational support for engineering design process reveals behavior of designed system in response to external stimuli; and finds out how behavior modified by changing physical attributes of system. System-sensitivity analysis combined with extrapolation forms model of design complementary to model of behavior, capable of direct simulation of effects of changes in design variables. Algorithms developed for this method applicable to design of large engineering systems, especially those consisting of several subsystems involving many disciplines.

  9. Control of Stirling engine. Simplified, compressible model

    Science.gov (United States)

    Plotnikov, P. I.; Sokołowski, J.; Żochowski, A.

    2016-06-01

    A one-dimensional free boundary problem on a motion of a heavy piston in a tube filled with viscous gas is considered. The system of governing equations and boundary conditions is derived. The obtained system of differential equations can be regarded as a mathematical model of an exterior combustion engine. The existence of a weak solution to this model is proved. The problem of maximization of the total work of the engine is considered.

  10. Parametric studies with an atmospheric diffusion model that assesses toxic fuel hazards due to the ground clouds generated by rocket launches

    Science.gov (United States)

    Stewart, R. B.; Grose, W. L.

    1975-01-01

    Parametric studies were made with a multilayer atmospheric diffusion model to place quantitative limits on the uncertainty of predicting ground-level toxic rocket-fuel concentrations. Exhaust distributions in the ground cloud, cloud stabilized geometry, atmospheric coefficients, the effects of exhaust plume afterburning of carbon monoxide CO, assumed surface mixing-layer division in the model, and model sensitivity to different meteorological regimes were studied. Large-scale differences in ground-level predictions are quantitatively described. Cloud alongwind growth for several meteorological conditions is shown to be in error because of incorrect application of previous diffusion theory. In addition, rocket-plume calculations indicate that almost all of the rocket-motor carbon monoxide is afterburned to carbon dioxide CO2, thus reducing toxic hazards due to CO. The afterburning is also shown to have a significant effect on cloud stabilization height and on ground-level concentrations of exhaust products.

  11. Design of Turning Equipment of Niobium Ring-ceramic Matrix Composites Rocket Engine Thrusters%铌环-C/SiC复合材料喷管的车削工装设计

    Institute of Scientific and Technical Information of China (English)

    黎明; 顾力强; 郭洪勤

    2014-01-01

    A model of rocket engine adopts the ceramic matrix composites nozzle, on which a niobium ring was deposited to connect to the engine by means of electron beam welding. For turning the welding stage of niobium ring, this paper introduces a kind of process equipment design train of thought. By using combining process equipment to find new positioning base, it effectively solved the problem of the fracture of ceramic matrix composites. This technology completely satisfied the requirements of tolerance of welding stage by turning results. And it improves the performance of the engine much further.%某火箭发动机采用 C/SiC复合材料的喷管,并在喷管上沉积一个带焊接头的铌环,通过该铌环与发动机金属头部进行电子束焊接。现需要车削完成喷管铌环的焊接台阶等特征。本文介绍了一种工装设计的新思路,即通过组合车削工装寻找定位基准。通过实际加工证明:该工艺技术完全满足焊接台阶的形位公差要求,并有效地解决了 C/SiC 复合材料脆性断裂等难题,最终满足了发动机的性能指标。

  12. CFD Modelling of a Quadrupole Vortex Inside a Cylindrical Channel for Research into Advanced Hybrid Rocket Designs

    Science.gov (United States)

    Godfrey, B.; Majdalani, J.

    2014-11-01

    This study relies on computational fluid dynamics (CFD) tools to analyse a possible method for creating a stable quadrupole vortex within a simulated, circular-port, cylindrical rocket chamber. A model of the vortex generator is created in a SolidWorks CAD program and then the grid is generated using the Pointwise mesh generation software. The non-reactive flowfield is simulated using an open source computational program, Stanford University Unstructured (SU2). Subsequent analysis and visualization are performed using ParaView. The vortex generation approach that we employ consists of four tangentially injected monopole vortex generators that are arranged symmetrically with respect to the center of the chamber in such a way to produce a quadrupole vortex with a common downwash. The present investigation focuses on characterizing the flow dynamics so that future investigations can be undertaken with increasing levels of complexity. Our CFD simulations help to elucidate the onset of vortex filaments within the monopole tubes, and the evolution of quadrupole vortices downstream of the injection faceplate. Our results indicate that the quadrupole vortices produced using the present injection pattern can become quickly unstable to the extent of dissipating soon after being introduced into simulated rocket chamber. We conclude that a change in the geometrical configuration will be necessary to produce more stable quadrupoles.

  13. A Simple HCCI Engine Model for Control

    Energy Technology Data Exchange (ETDEWEB)

    Killingsworth, N; Aceves, S; Flowers, D; Krstic, M

    2006-06-29

    The homogeneous charge compression ignition (HCCI) engine is an attractive technology because of its high efficiency and low emissions. However, HCCI lacks a direct combustion trigger making control of combustion timing challenging, especially during transients. To aid in HCCI engine control we present a simple model of the HCCI combustion process valid over a range of intake pressures, intake temperatures, equivalence ratios, and engine speeds. The model provides an estimate of the combustion timing on a cycle-by-cycle basis. An ignition threshold, which is a function of the in-cylinder motored temperature and pressure is used to predict start of combustion. This model allows the synthesis of nonlinear control laws, which can be utilized for control of an HCCI engine during transients.

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

    Institute of Scientific and Technical Information of China (English)

    牛小莉

    2012-01-01

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

  15. Dynamic model for the internal combustion engine

    Energy Technology Data Exchange (ETDEWEB)

    Rizzoni, G.

    1986-01-01

    Over the last decade there has been increasing interest in the application of control theory to passenger vehicles: stringent governmental regulations constraining fuel consumption and exhaust emissions have required a shift to integrated electronics controls. Unfortunately, the lack of robust global models for the dynamics of the IC engine has limited the application of the tools of control theory in this areas. This dissertation is devoted to the formulation of a robust model for the dynamics of the IC engine. The engine is viewed as a system with inputs given by cylinder pressure and net engine torque, and output corresponding to crankshaft angular acceleration. The model is well suited to closed loop engine and transmission control applications. The deterministic model provides a powerful tool for estimating average and instantaneous net engine torque based on a noncontacting measurement of crankshaft acceleration. The stochastic model explains cyclic pressure variations by an additive Gaussian WSS vector noise process. Further, it demonstrates that by means of a suitable linear transformation-invariant with load and RPM-, the noise process may be expressed in terms of a three-dimensional uncorrelated vector random process.

  16. Unsteady Three-Dimensional Simulation of a Shear Coaxial GO2/GH2 Rocket Injector with RANS and Hybrid-RAN-LES/DES Using Flamelet Models

    Science.gov (United States)

    Westra, Doug G.; West, Jeffrey S.; Richardson, Brian R.

    2015-01-01

    Historically, the analysis and design of liquid rocket engines (LREs) has relied on full-scale testing and one-dimensional empirical tools. The testing is extremely expensive and the one-dimensional tools are not designed to capture the highly complex, and multi-dimensional features that are inherent to LREs. Recent advances in computational fluid dynamics (CFD) tools have made it possible to predict liquid rocket engine performance, stability, to assess the effect of complex flow features, and to evaluate injector-driven thermal environments, to mitigate the cost of testing. Extensive efforts to verify and validate these CFD tools have been conducted, to provide confidence for using them during the design cycle. Previous validation efforts have documented comparisons of predicted heat flux thermal environments with test data for a single element gaseous oxygen (GO2) and gaseous hydrogen (GH2) injector. The most notable validation effort was a comprehensive validation effort conducted by Tucker et al. [1], in which a number of different groups modeled a GO2/GH2 single element configuration by Pal et al [2]. The tools used for this validation comparison employed a range of algorithms, from both steady and unsteady Reynolds Averaged Navier-Stokes (U/RANS) calculations, large-eddy simulations (LES), detached eddy simulations (DES), and various combinations. A more recent effort by Thakur et al. [3] focused on using a state-of-the-art CFD simulation tool, Loci/STREAM, on a two-dimensional grid. Loci/STREAM was chosen because it has a unique, very efficient flamelet parameterization of combustion reactions that are too computationally expensive to simulate with conventional finite-rate chemistry calculations. The current effort focuses on further advancement of validation efforts, again using the Loci/STREAM tool with the flamelet parameterization, but this time with a three-dimensional grid. Comparisons to the Pal et al. heat flux data will be made for both RANS and

  17. Efficient Model-Based Diagnosis Engine

    Science.gov (United States)

    Fijany, Amir; Vatan, Farrokh; Barrett, Anthony; James, Mark; Mackey, Ryan; Williams, Colin

    2009-01-01

    An efficient diagnosis engine - a combination of mathematical models and algorithms - has been developed for identifying faulty components in a possibly complex engineering system. This model-based diagnosis engine embodies a twofold approach to reducing, relative to prior model-based diagnosis engines, the amount of computation needed to perform a thorough, accurate diagnosis. The first part of the approach involves a reconstruction of the general diagnostic engine to reduce the complexity of the mathematical-model calculations and of the software needed to perform them. The second part of the approach involves algorithms for computing a minimal diagnosis (the term "minimal diagnosis" is defined below). A somewhat lengthy background discussion is prerequisite to a meaningful summary of the innovative aspects of the present efficient model-based diagnosis engine. In model-based diagnosis, the function of each component and the relationships among all the components of the engineering system to be diagnosed are represented as a logical system denoted the system description (SD). Hence, the expected normal behavior of the engineering system is the set of logical consequences of the SD. Faulty components lead to inconsistencies between the observed behaviors of the system and the SD (see figure). Diagnosis - the task of finding faulty components - is reduced to finding those components, the abnormalities of which could explain all the inconsistencies. The solution of the diagnosis problem should be a minimal diagnosis, which is a minimal set of faulty components. A minimal diagnosis stands in contradistinction to the trivial solution, in which all components are deemed to be faulty, and which, therefore, always explains all inconsistencies.

  18. Engineering Abstractions in Model Checking and Testing

    DEFF Research Database (Denmark)

    Achenbach, Michael; Ostermann, Klaus

    2009-01-01

    Abstractions are used in model checking to tackle problems like state space explosion or modeling of IO. The application of these abstractions in real software development processes, however, lacks engineering support. This is one reason why model checking is not widely used in practice yet...... and testing is still state of the art in falsification. We show how user-defined abstractions can be integrated into a Java PathFinder setting with tools like AspectJ or Javassist and discuss implications of remaining weaknesses of these tools. We believe that a principled engineering approach to designing...... and implementing abstractions will improve the applicability of model checking in practice....

  19. Learning to Model in Engineering

    Science.gov (United States)

    Gainsburg, Julie

    2013-01-01

    Policymakers and education scholars recommend incorporating mathematical modeling into mathematics education. Limited implementation of modeling instruction in schools, however, has constrained research on how students learn to model, leaving unresolved debates about whether modeling should be reified and explicitly taught as a competence, whether…

  20. Coordinated control for regulation/protection mode-switching of ducted rockets

    Science.gov (United States)

    Qi, Yiwen; Bao, Wen; Zhao, Jun; Chang, Juntao

    2014-05-01

    This study is concerned with the coordinated control problem for regulation/protection mode-switching of a ducted rocket, in order to obtain the maximum system performance while ensuring safety. The proposed strategy has an inner/outer loop control structure which decomposes the contradiction between performance and safety into two modes of regulation and protection. Specifically, first, the mathematical model including the actuator (gas regulating system) and the plant (ducted rocket engine) is introduced. Second, taking the inlet buzz for instance, the ducted rocket coordinated control problem for thrust regulation and inlet buzz limit protection is formulated and discussed. Third, to solve the problem, based on the main inner loop, a limit protection controller (outer loop) design method is developed utilizing a linear quadratic optimal control technique, and a coordinated control logic is then presented. At last, the whole coordinated control strategy is applied to the ducted rocket control model, and simulation results demonstrate its effectiveness.

  1. Graph-based modelling in engineering

    CERN Document Server

    Rysiński, Jacek

    2017-01-01

    This book presents versatile, modern and creative applications of graph theory in mechanical engineering, robotics and computer networks. Topics related to mechanical engineering include e.g. machine and mechanism science, mechatronics, robotics, gearing and transmissions, design theory and production processes. The graphs treated are simple graphs, weighted and mixed graphs, bond graphs, Petri nets, logical trees etc. The authors represent several countries in Europe and America, and their contributions show how different, elegant, useful and fruitful the utilization of graphs in modelling of engineering systems can be. .

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

    Science.gov (United States)

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

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

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

    Science.gov (United States)

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

    2004-08-01

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

  4. Systems Engineering Model for ART Energy Conversion

    Energy Technology Data Exchange (ETDEWEB)

    Mendez Cruz, Carmen Margarita [Sandia National Lab. (SNL-NM), Albuquerque, NM (United States); Rochau, Gary E. [Sandia National Lab. (SNL-NM), Albuquerque, NM (United States); Wilson, Mollye C. [Sandia National Lab. (SNL-NM), Albuquerque, NM (United States)

    2017-02-01

    The near-term objective of the EC team is to establish an operating, commercially scalable Recompression Closed Brayton Cycle (RCBC) to be constructed for the NE - STEP demonstration system (demo) with the lowest risk possible. A systems engineering approach is recommended to ensure adequate requirements gathering, documentation, and mode ling that supports technology development relevant to advanced reactors while supporting crosscut interests in potential applications. A holistic systems engineering model was designed for the ART Energy Conversion program by leveraging Concurrent Engineering, Balance Model, Simplified V Model, and Project Management principles. The resulting model supports the identification and validation of lifecycle Brayton systems requirements, and allows designers to detail system-specific components relevant to the current stage in the lifecycle, while maintaining a holistic view of all system elements.

  5. Study of Photovoltaic Cells Engineering Mathematical Model

    Science.gov (United States)

    Zhou, Jun; Yu, Zhengping; Lu, Zhengyi; Li, Chenhui; Zhang, Ruilan

    2016-11-01

    The characteristic curve of photovoltaic cells is the theoretical basis of PV Power, which simplifies the existing mathematical model, eventually, obtains a mathematical model used in engineering. The characteristic curve of photovoltaic cells contains both exponential and logarithmic calculation. The exponential and logarithmic spread out through Taylor series, which includes only four arithmetic and use single chip microcontroller as the control center. The result shows that: the use of single chip microcontroller for calculating exponential and logarithmic functions, simplifies mathematical model of PV curve, also can meet the specific conditions’ requirement for engineering applications.

  6. Mathematical modeling a chemical engineer's perspective

    CERN Document Server

    Rutherford, Aris

    1999-01-01

    Mathematical modeling is the art and craft of building a system of equations that is both sufficiently complex to do justice to physical reality and sufficiently simple to give real insight into the situation. Mathematical Modeling: A Chemical Engineer's Perspective provides an elementary introduction to the craft by one of the century's most distinguished practitioners.Though the book is written from a chemical engineering viewpoint, the principles and pitfalls are common to all mathematical modeling of physical systems. Seventeen of the author's frequently cited papers are reprinted to illus

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

    Science.gov (United States)

    Wehrmeyer, Joseph A.; Trinh, Huu Phuoc

    2001-01-01

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

  8. Modeling and Simulation of Hydraulic Engine Mounts

    Institute of Scientific and Technical Information of China (English)

    DUAN Shanzhong; Marshall McNea

    2012-01-01

    Hydraulic engine mounts are widely used in automotive powertrains for vibration isolation.A lumped mechanical parameter model is a traditional approach to model and simulate such mounts.This paper presents a dynamical model of a passive hydraulic engine mount with a double-chamber,an inertia track,a decoupler,and a plunger.The model is developed based on analogy between electrical systems and mechanical-hydraulic systems.The model is established to capture both low and high frequency dynatmic behaviors of the hydraulic mount.The model will be further used to find the approximate pulse responses of the mounts in terms of the force transmission and top chamber pressure.The close form solution from the simplifiod linear model may provide some insight into the highly nonlinear behavior of the mounts.Based on the model,computer simulation has been carried out to study dynamic performance of the hydraulic mount.

  9. Performance Engineering in the Community Atmosphere Model

    Energy Technology Data Exchange (ETDEWEB)

    Worley, P; Mirin, A; Drake, J; Sawyer, W

    2006-05-30

    The Community Atmosphere Model (CAM) is the atmospheric component of the Community Climate System Model (CCSM) and is the primary consumer of computer resources in typical CCSM simulations. Performance engineering has been an important aspect of CAM development throughout its existence. This paper briefly summarizes these efforts and their impacts over the past five years.

  10. Software Engineering Tools for Scientific Models

    Science.gov (United States)

    Abrams, Marc; Saboo, Pallabi; Sonsini, Mike

    2013-01-01

    Software tools were constructed to address issues the NASA Fortran development community faces, and they were tested on real models currently in use at NASA. These proof-of-concept tools address the High-End Computing Program and the Modeling, Analysis, and Prediction Program. Two examples are the NASA Goddard Earth Observing System Model, Version 5 (GEOS-5) atmospheric model in Cell Fortran on the Cell Broadband Engine, and the Goddard Institute for Space Studies (GISS) coupled atmosphere- ocean model called ModelE, written in fixed format Fortran.

  11. Formation of vortex structures in channels with mass injection and their interaction with surfaces in solid-fuel rocket engines

    Science.gov (United States)

    Benderskiy, B. Ya.; Chernova, A. A.

    2015-03-01

    The topological features of the structure of combustion products flow in the flow paths with different shapes of channel cross sections at power installations are considered. The results of mathematical modeling of internal gas dynamics of the flow paths of power installations are compared with experimental data.

  12. Chain modeling for life cycle systems engineering

    Energy Technology Data Exchange (ETDEWEB)

    Rivera, J.J. [Sandia National Lab., Albuquerque, NM (United States); Shapiro, V. [Univ. of Wisconsin, Madison, WI (United States). Spatial Automation Lab.

    1997-12-01

    Throughout Sandia`s history, products have been represented by drawings. Solid modeling systems have recently replaced drawings as the preferred means for representing product geometry. These systems are used for product visualization, engineering analysis and manufacturing planning. Unfortunately, solid modeling technology is inadequate for life cycle systems engineering, which requires maintenance of technical history, efficient management of geometric and non-geometric data, and explicit representation of engineering and manufacturing characteristics. Such information is not part of the mathematical foundation of solid modeling. The current state-of-the-art in life cycle engineering is comprised of painstakingly created special purpose tools, which often are incompatible. New research on {open_quotes}chain modeling{close_quotes} provides a method of chaining the functionality of a part to the geometric representation. Chain modeling extends classical solid modeling to include physical, manufacturing, and procedural information required for life cycle engineering. In addition, chain modeling promises to provide the missing theoretical basis for Sandia`s parent/child product realization paradigm. In chain modeling, artifacts and systems are characterized in terms of their combinatorial properties: cell complexes, chains, and their operators. This approach is firmly rooted in algebraic topology and is a natural extension of current technology. The potential benefits of this approach include explicit hierarchical and combinatorial representation of physics, geometry, functionality, test, and legacy data in a common computational framework that supports a rational decision process and partial design automation. Chain modeling will have a significant impact on design preservation, system identification, parameterization, system reliability, and design simplification.

  13. PARAMETER ESTIMATION OF ENGINEERING TURBULENCE MODEL

    Institute of Scientific and Technical Information of China (English)

    钱炜祺; 蔡金狮

    2001-01-01

    A parameter estimation algorithm is introduced and used to determine the parameters in the standard k-ε two equation turbulence model (SKE). It can be found from the estimation results that although the parameter estimation method is an effective method to determine model parameters, it is difficult to obtain a set of parameters for SKE to suit all kinds of separated flow and a modification of the turbulence model structure should be considered. So, a new nonlinear k-ε two-equation model (NNKE) is put forward in this paper and the corresponding parameter estimation technique is applied to determine the model parameters. By implementing the NNKE to solve some engineering turbulent flows, it is shown that NNKE is more accurate and versatile than SKE. Thus, the success of NNKE implies that the parameter estimation technique may have a bright prospect in engineering turbulence model research.

  14. Computer code for single-point thermodynamic analysis of hydrogen/oxygen expander-cycle rocket engines

    Science.gov (United States)

    Glassman, Arthur J.; Jones, Scott M.

    1991-01-01

    This analysis and this computer code apply to full, split, and dual expander cycles. Heat regeneration from the turbine exhaust to the pump exhaust is allowed. The combustion process is modeled as one of chemical equilibrium in an infinite-area or a finite-area combustor. Gas composition in the nozzle may be either equilibrium or frozen during expansion. This report, which serves as a users guide for the computer code, describes the system, the analysis methodology, and the program input and output. Sample calculations are included to show effects of key variables such as nozzle area ratio and oxidizer-to-fuel mass ratio.

  15. Rocket injector anomalies study. Volume 2: Results of parametric studies

    Science.gov (United States)

    Przekwas, A. J.; Singhal, A. K.; Tam, L. T.

    1984-01-01

    The employment of a existing computer program to simulate three dimensional two phase gas spray flows in liquid propellant rocket engines. This was accomplished by modification of an existing three dimensional computer program (REFLAN3D) with Euler/Lagrange approach for simulating two phase spray flow, evaporation and combustion. The modified code is referred to as REFLAN3D-SPRAY. Computational studies of the model rocket engine combustion chamber are presented. The parametric studies of the two phase flow and combustion shows qualitatively correct response for variations in geometrical and physical parameters. The injection nonuniformity test with blocked central fuel injector holes shows significant changes in the central flame core and minor influence on the wall heat transfer fluxes.

  16. Engine modeling and control modeling and electronic management of internal combustion engines

    CERN Document Server

    Isermann, Rolf

    2014-01-01

    The increasing demands for internal combustion engines with regard to fuel consumption, emissions and driveability lead to more actuators, sensors and complex control functions. A systematic implementation of the electronic control systems requires mathematical models from basic design through simulation to calibration. The book treats physically-based as well as models based experimentally on test benches for gasoline (spark ignition) and diesel (compression ignition) engines and uses them for the design of the different control functions. The main topics are: - Development steps for engine control - Stationary and dynamic experimental modeling - Physical models of intake, combustion, mechanical system, turbocharger, exhaust, cooling, lubrication, drive train - Engine control structures, hardware, software, actuators, sensors, fuel supply, injection system, camshaft - Engine control methods, static and dynamic feedforward and feedback control, calibration and optimization, HiL, RCP, control software developm...

  17. Numerical modeling in materials science and engineering

    CERN Document Server

    Rappaz, Michel; Deville, Michel

    2003-01-01

    This book introduces the concepts and methodologies related to the modelling of the complex phenomena occurring in materials processing. After a short reminder of conservation laws and constitutive relationships, the authors introduce the main numerical methods: finite differences, finite volumes and finite elements. These techniques are developed in three main chapters of the book that tackle more specific problems: phase transformation, solid mechanics and fluid flow. The two last chapters treat inverse methods to obtain the boundary conditions or the material properties and stochastic methods for microstructural simulation. This book is intended for undergraduate and graduate students in materials science and engineering, mechanical engineering and physics and for engineering professionals or researchers who want to get acquainted with numerical simulation to model and compute materials processing.

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

    Institute of Scientific and Technical Information of China (English)

    丁兆波; 许晓勇; 乔桂玉; 陶瑞峰

    2013-01-01

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

  19. User Requirements and Domain Model Engineering

    NARCIS (Netherlands)

    Specht, Marcus; Glahn, Christian

    2006-01-01

    Specht, M., & Glahn, C. (2006). User requirements and domain model engineering. Presentation at International Workshop in Learning Networks for Lifelong Competence Development. March, 30-31, 2006. Sofia, Bulgaria: TENCompetence Conference. Retrieved June 30th, 2006, from http://dspace.learningnetwor

  20. User Requirements and Domain Model Engineering

    NARCIS (Netherlands)

    Specht, Marcus; Glahn, Christian

    2006-01-01

    Specht, M., & Glahn, C. (2006). User requirements and domain model engineering. Presentation at International Workshop in Learning Networks for Lifelong Competence Development. March, 30-31, 2006. Sofia, Bulgaria: TENCompetence Conference. Retrieved June 30th, 2006, from http://dspace.learningnetwor

  1. Mars Rocket Propulsion System

    Science.gov (United States)

    Zubrin, Robert; Harber, Dan; Nabors, Sammy

    2008-01-01

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

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

    Directory of Open Access Journals (Sweden)

    K. I. Denisova

    2016-01-01

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

  3. Multi dimentional modeling of a CI engine

    Energy Technology Data Exchange (ETDEWEB)

    Koten, Hasan; Yilmaz, Mustafa; Zafer Gul, M. [Marmara University Mechanical Engineering Department (Turkey)], E-mail: hasan.koten@marmara.edu.tr

    2011-07-01

    With the coming shortage of fossil fuels and rising concerns about the environment, it is important to develop new technologies that reduce both energy consumption and pollution at the same time. In the transportation sector, new combustion processes are under development to provide clean diesel combustion with no particulate and NOx emissions. However, these processes have issues such as limited power output, high levels of unburned hydrocarbons, and carbon monoxide emissions. The aim of this paper is to determine in-cylinder flow characteristics to improve combustion performance. Combustion modeling was performed using the ECFM-3Z combustion model and 1D dynamic model and calculations on the configuration of a direct injection diesel engine were made. This study showed that the new ECFM-3Z combustion model provides results in accordance with previous research but that further studies are needed to determine the optimum engine parameters.

  4. Rocket propulsion elements

    CERN Document Server

    Sutton, George P

    2011-01-01

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

  5. Mechanics, Models and Methods in Civil Engineering

    CERN Document Server

    Maceri, Franco

    2012-01-01

    „Mechanics, Models and Methods in Civil Engineering” collects leading papers dealing with actual Civil Engineering problems. The approach is in the line of the Italian-French school and therefore deeply couples mechanics and mathematics creating new predictive theories, enhancing clarity in understanding, and improving effectiveness in applications. The authors of the contributions collected here belong to the Lagrange Laboratory, an European Research Network active since many years. This book will be of a major interest for the reader aware of modern Civil Engineering.

  6. Engineering design of systems models and methods

    CERN Document Server

    Buede, Dennis M

    2009-01-01

    The ideal introduction to the engineering design of systems-now in a new edition. The Engineering Design of Systems, Second Edition compiles a wealth of information from diverse sources to provide a unique, one-stop reference to current methods for systems engineering. It takes a model-based approach to key systems engineering design activities and introduces methods and models used in the real world. Features new to this edition include: * The addition of Systems Modeling Language (SysML) to several of the chapters, as well as the introduction of new terminology * Additional material on partitioning functions and components * More descriptive material on usage scenarios based on literature from use case development * Updated homework assignments * The software product CORE (from Vitech Corporation) is used to generate the traditional SE figures and the software product MagicDraw UML with SysML plugins (from No Magic, Inc.) is used for the SysML figures This book is designed to be an introductory reference ...

  7. Genome-scale modeling for metabolic engineering

    Energy Technology Data Exchange (ETDEWEB)

    Simeonidis, E; Price, ND

    2015-01-13

    We focus on the application of constraint-based methodologies and, more specifically, flux balance analysis in the field of metabolic engineering, and enumerate recent developments and successes of the field. We also review computational frameworks that have been developed with the express purpose of automatically selecting optimal gene deletions for achieving improved production of a chemical of interest. The application of flux balance analysis methods in rational metabolic engineering requires a metabolic network reconstruction and a corresponding in silico metabolic model for the microorganism in question. For this reason, we additionally present a brief overview of automated reconstruction techniques. Finally, we emphasize the importance of integrating metabolic networks with regulatory information-an area which we expect will become increasingly important for metabolic engineering-and present recent developments in the field of metabolic and regulatory integration.

  8. Genome-scale modeling for metabolic engineering.

    Science.gov (United States)

    Simeonidis, Evangelos; Price, Nathan D

    2015-03-01

    We focus on the application of constraint-based methodologies and, more specifically, flux balance analysis in the field of metabolic engineering, and enumerate recent developments and successes of the field. We also review computational frameworks that have been developed with the express purpose of automatically selecting optimal gene deletions for achieving improved production of a chemical of interest. The application of flux balance analysis methods in rational metabolic engineering requires a metabolic network reconstruction and a corresponding in silico metabolic model for the microorganism in question. For this reason, we additionally present a brief overview of automated reconstruction techniques. Finally, we emphasize the importance of integrating metabolic networks with regulatory information-an area which we expect will become increasingly important for metabolic engineering-and present recent developments in the field of metabolic and regulatory integration.

  9. Information Technology Model for Product Lifecycle Engineering

    Directory of Open Access Journals (Sweden)

    Bhanumathi KS

    2013-02-01

    Full Text Available An aircraft is a complex, multi-disciplinary, system-engineered product that requires real-time global technical collaboration through its life-cycle. Engineering data and processes which form the backbone of the aircraft should be under strict Configuration Control (CC. It should be model-based and allow for 3D visualization and manipulation. This requires accurate, realtime collaboration and concurrent engineering-based business processes operating in an Integrated Digital Environment (IDE. The IDE uses lightweight, neutral Computer Aided Design (CAD Digital Mock-Up (DMU. The DMU deals with complex structural assemblies and systems of more than a hundred thousand parts created by engineers across the globe, each using diverse CAD, Computer Aided Engineering (CAE, Computer Aided Manufacturing (CAM, Computer Integrated Manufacturing (CIM, Enterprise Resource Planning (ERP, Supply Chain Management(SCM,Customer Relationship Management(CRM and Computer Aided Maintenance Management System (CAMMS systems. In this paper, a comprehensive approach to making such an environment a reality is presented.

  10. Reduced Basis and Stochastic Modeling of Liquid Propellant Rocket Engine as a Complex System

    Science.gov (United States)

    2015-07-02

    Schlichting H. and Gersten K. Boundary Layer Theory. Springer, 2000. [31] White F. Viscous Fluid Flow. Tata McGraw Hill, 2011. [32] Ogata K. Modern control ...display a currently valid OMB control number. PLEASE DO NOT RETURN YOUR FORM TO THE ABOVE ORGANIZATION. 1.  REPORT DATE (DD-MM-YYYY)      21-07-2015 2...and control has been seen. Discussion in the next four sections will provide some detail. Additional detail is provided through the addenda

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

    Institute of Scientific and Technical Information of China (English)

    李丹琳; 田原; 孙纪国

    2012-01-01

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

  12. Mini-Rocket User Guide

    Science.gov (United States)

    2007-08-01

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

  13. Engineering

    National Research Council Canada - National Science Library

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

  14. Cycle Engine Modelling Of Spark Ignition Engine Processes during Wide-Open Throttle (WOT) Engine Operation Running By Gasoline Fuel

    Science.gov (United States)

    Rahim, M. F. Abdul; Rahman, M. M.; Bakar, R. A.

    2012-09-01

    One-dimensional engine model is developed to simulate spark ignition engine processes in a 4-stroke, 4 cylinders gasoline engine. Physically, the baseline engine is inline cylinder engine with 3-valves per cylinder. Currently, the engine's mixture is formed by external mixture formation using piston-type carburettor. The model of the engine is based on one-dimensional equation of the gas exchange process, isentropic compression and expansion, progressive engine combustion process, and accounting for the heat transfer and frictional losses as well as the effect of valves overlapping. The model is tested for 2000, 3000 and 4000 rpm of engine speed and validated using experimental engine data. Results showed that the engine is able to simulate engine's combustion process and produce reasonable prediction. However, by comparing with experimental data, major discrepancy is noticeable especially on the 2000 and 4000 rpm prediction. At low and high engine speed, simulated cylinder pressures tend to under predict the measured data. Whereas the cylinder temperatures always tend to over predict the measured data at all engine speed. The most accurate prediction is obtained at medium engine speed of 3000 rpm. Appropriate wall heat transfer setup is vital for more precise calculation of cylinder pressure and temperature. More heat loss to the wall can lower cylinder temperature. On the hand, more heat converted to the useful work mean an increase in cylinder pressure. Thus, instead of wall heat transfer setup, the Wiebe combustion parameters are needed to be carefully evaluated for better results.

  15. Thrust Vector Control for Nuclear Thermal Rockets

    Science.gov (United States)

    Ensworth, Clinton B. F.

    2013-01-01

    Future space missions may use Nuclear Thermal Rocket (NTR) stages for human and cargo missions to Mars and other destinations. The vehicles are likely to require engine thrust vector control (TVC) to maintain desired flight trajectories. This paper explores requirements and concepts for TVC systems for representative NTR missions. Requirements for TVC systems were derived using 6 degree-of-freedom models of NTR vehicles. Various flight scenarios were evaluated to determine vehicle attitude control needs and to determine the applicability of TVC. Outputs from the models yielded key characteristics including engine gimbal angles, gimbal rates and gimbal actuator power. Additional factors such as engine thrust variability and engine thrust alignment errors were examined for impacts to gimbal requirements. Various technologies are surveyed for TVC systems for the NTR applications. A key factor in technology selection is the unique radiation environment present in NTR stages. Other considerations including mission duration and thermal environments influence the selection of optimal TVC technologies. Candidate technologies are compared to see which technologies, or combinations of technologies best fit the requirements for selected NTR missions. Representative TVC systems are proposed and key properties such as mass and power requirements are defined. The outputs from this effort can be used to refine NTR system sizing models, providing higher fidelity definition for TVC systems for future studies.

  16. Solar Thermal Rocket Propulsion

    Science.gov (United States)

    Sercel, J. C.

    1986-01-01

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

  17. Laser optogalvanic spectroscopy of neon in a discharge plasma and modeling and analysis of rocket plume RF-line emissions

    Science.gov (United States)

    Ogungbemi, Kayode I.

    The Optogalvanic Effect (OGE) of neon in a hollow cathode discharge lamp has been investigated both experimentally and theoretically. A tunable dye laser was tuned to several 1si -- 2pj neon transitions and the associated time--resolved optogalvanic (OG) spectral waveforms recorded corresponding to the DeltaJ = DeltaK = 0, +/-1 selection rules and modeled using a semi-empirical model. Decay rate constants, amplitudes and the instrumentation time constants were recorded following a good least-squares fit (between the experimental and the theoretical OG data) using the Monte Carlo technique and utilizing both the search and random walk methods. Dominant physical processes responsible for the optogalvanic effect have been analyzed, and the corresponding populations of the laser-excited level and collisional excited levels determined. The behavior of the optogalvanic signal waveform as a function of time, together with the decay rate constants as a function of the discharge current and the instrumentation time constant as a function of current have been studied in detail. The decay times of the OG signals and the population redistributions were also determined. Fairly linear relationships between the decay rate constant and the discharge current, as well as between the instrumental time constant and the discharge current, have been observed. The decay times and the electron collisional rate parameters of the 1s levels involved in the OG transitions have been obtained with accuracy. The excitation temperature of the discharge for neon transitions grouped with the same 1s level have been determined and found to be fairly constant for the neon transitions studied. The experimental optogalvanic effort in the visible region of the electromagnetic spectrum has been complemented by a computation-intensive modeling investigation of rocket plumes in the microwave region. Radio frequency lines of each of the plume species identified were archived utilizing the HITRAN and other

  18. 3D Modeling Engine Representation Summary Report

    Energy Technology Data Exchange (ETDEWEB)

    Steven Prescott; Ramprasad Sampath; Curtis Smith; Timothy Yang

    2014-09-01

    Computers have been used for 3D modeling and simulation, but only recently have computational resources been able to give realistic results in a reasonable time frame for large complex models. This summary report addressed the methods, techniques, and resources used to develop a 3D modeling engine to represent risk analysis simulation for advanced small modular reactor structures and components. The simulations done for this evaluation were focused on external events, specifically tsunami floods, for a hypothetical nuclear power facility on a coastline.

  19. Software-Engineering Process Simulation (SEPS) model

    Science.gov (United States)

    Lin, C. Y.; Abdel-Hamid, T.; Sherif, J. S.

    1992-01-01

    The Software Engineering Process Simulation (SEPS) model is described which was developed at JPL. SEPS is a dynamic simulation model of the software project development process. It uses the feedback principles of system dynamics to simulate the dynamic interactions among various software life cycle development activities and management decision making processes. The model is designed to be a planning tool to examine tradeoffs of cost, schedule, and functionality, and to test the implications of different managerial policies on a project's outcome. Furthermore, SEPS will enable software managers to gain a better understanding of the dynamics of software project development and perform postmodern assessments.

  20. Mean Value Modelling of an SI Engine with EGR

    DEFF Research Database (Denmark)

    Føns, Michael; Muller, Martin; Chevalier, Alain

    1999-01-01

    Mean Value Engine Models (MVEMs) are simplified, dynamic engine models which are physically based. Such models are useful for control studies, for engine control system analysis and for model based control systems. Very few published MVEMs have included the effects of Exhaust Gas Recirculation (E...

  1. Human Factors Engineering Program Review Model

    Science.gov (United States)

    2004-02-01

    AA NUREG -0711,Rev. 2 Human Factors Engineering Program Review Model 20081009191 I i m To] Bi U.S. Nuclear Regulatory Commission Office of...Material As of November 1999, you may electronically access NUREG -series publications and other NRC records at NRC’s Public Electronic Reading Room at...http://www.nrc.qov/readinq-rm.html. Publicly released records include, to name a few, NUREG -series publications; Federal Register notices; applicant

  2. Linguistic Model for Engine Power Loss

    Science.gov (United States)

    2011-11-27

    model, engine power loss, intelligent diagnostics 16. SECURITY CLASSIFICATION OF: 17. LIMITATION OF ABSTRACT Public Release 18. NUMBER OF PAGES...module via a wireless interface that is compliant with the IEEE 802.15 protocol. The wireless personal area network is implemented using the ZigBee ... secure short range wireless communications link, or via a wired Ethernet port to a mobile device (e.g., another laptop computer authorized for CBM

  3. Finite element modeling for materials engineers using Matlab

    CERN Document Server

    Oluwole, Oluleke

    2014-01-01

    Finite Element Modeling for Materials Engineers Using MATLAB® combines the finite element method with MATLAB to offer materials engineers a fast and code-free way of modeling for many materials processes.

  4. Rockets two classic papers

    CERN Document Server

    Goddard, Robert

    2002-01-01

    Rockets, in the primitive form of fireworks, have existed since the Chinese invented them around the thirteenth century. But it was the work of American Robert Hutchings Goddard (1882-1945) and his development of liquid-fueled rockets that first produced a controlled rocket flight. Fascinated by rocketry since boyhood, Goddard designed, built, and launched the world's first liquid-fueled rocket in 1926. Ridiculed by the press for suggesting that rockets could be flown to the moon, he continued his experiments, supported partly by the Smithsonian Institution and defended by Charles Lindbergh. T

  5. SEISMIC MODELING ENGINES PHASE 1 FINAL REPORT

    Energy Technology Data Exchange (ETDEWEB)

    BRUCE P. MARION

    2006-02-09

    Seismic modeling is a core component of petroleum exploration and production today. Potential applications include modeling the influence of dip on anisotropic migration; source/receiver placement in deviated-well three-dimensional surveys for vertical seismic profiling (VSP); and the generation of realistic data sets for testing contractor-supplied migration algorithms or for interpreting AVO (amplitude variation with offset) responses. This project was designed to extend the use of a finite-difference modeling package, developed at Lawrence Berkeley Laboratories, to the advanced applications needed by industry. The approach included a realistic, easy-to-use 2-D modeling package for the desktop of the practicing geophysicist. The feasibility of providing a wide-ranging set of seismic modeling engines was fully demonstrated in Phase I. The technical focus was on adding variable gridding in both the horizontal and vertical directions, incorporating attenuation, improving absorbing boundary conditions and adding the optional coefficient finite difference methods.

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

    Science.gov (United States)

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

    2003-01-01

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

  7. Generomak: Fusion physics, engineering and costing model

    Energy Technology Data Exchange (ETDEWEB)

    Delene, J.G.; Krakowski, R.A.; Sheffield, J.; Dory, R.A.

    1988-06-01

    A generic fusion physics, engineering and economics model (Generomak) was developed as a means of performing consistent analysis of the economic viability of alternative magnetic fusion reactors. The original Generomak model developed at Oak Ridge by Sheffield was expanded for the analyses of the Senior Committee on Environmental Safety and Economics of Magnetic Fusion Energy (ESECOM). This report describes the Generomak code as used by ESECOM. The input data used for each of the ten ESECOM fusion plants and the Generomak code output for each case is given. 14 refs., 3 figs., 17 tabs.

  8. BIOMASS REBURNING - MODELING/ENGINEERING STUDIES

    Energy Technology Data Exchange (ETDEWEB)

    Vladimir Zamansky; Chris Lindsey; Vitali Lissianski

    2000-01-28

    This project is designed to develop engineering and modeling tools for a family of NO{sub x} control technologies utilizing biomass as a reburning fuel. During the ninth reporting period (September 27--December 31, 1999), EER prepared a paper Kinetic Model of Biomass Reburning and submitted it for publication and presentation at the 28th Symposium (International) on Combustion, University of Edinburgh, Scotland, July 30--August 4, 2000. Antares Group Inc, under contract to Niagara Mohawk Power Corporation, evaluated the economic feasibility of biomass reburning options for Dunkirk Station. A preliminary report is included in this quarterly report.

  9. Heat transfer in rocket combustion chambers

    Science.gov (United States)

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

    1993-01-01

    Complexities of liquid rocket engine heat transfer which involve the injector faceplate and film cooled walls are being investigated by computational analysis. A conjugate heat transfer analysis was used to describe localized heating phenomena associated with particular injector configurations and film coolant flows. These components were analyzed, and the analyses verified when appropriate test data were available. The component analyses are being synthesized into an overall flowfield/heat transfer model. A Navier-Stokes flow solver, the FDNS code, was used to make the analyses. Particular attention was given to the representation of the thermodynamic properties of the fluid streams. Unit flow models of specific coaxial injector elements have been developed and are being used to describe the flame structure near the injector faceplate.

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

    Science.gov (United States)

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

    2000-01-01

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

  11. Transient Burning Rate Model for Solid Rocket Motor Internal Ballistic Simulations

    Directory of Open Access Journals (Sweden)

    David R. Greatrix

    2008-01-01

    Full Text Available A general numerical model based on the Zeldovich-Novozhilov solid-phase energy conservation result for unsteady solid-propellant burning is presented in this paper. Unlike past models, the integrated temperature distribution in the solid phase is utilized directly for estimating instantaneous burning rate (rather than the thermal gradient at the burning surface. The burning model is general in the sense that the model may be incorporated for various propellant burning-rate mechanisms. Given the availability of pressure-related experimental data in the open literature, varying static pressure is the principal mechanism of interest in this study. The example predicted results presented in this paper are to a substantial extent consistent with the corresponding experimental firing response data.

  12. Mean Value Modelling of a Turbocharged SI Engine

    DEFF Research Database (Denmark)

    Müller, Martin; Hendricks, Elbert; Sorenson, Spencer C.

    1998-01-01

    An important paradigm for the modelling of naturallly aspirated (NA) spark ignition (SI) engines for control purposes is the Mean Value Engine Model (MVEM). Such models have a time resolution which is just sufficient to capture the main details of the dynamic performance of NA SI engines but not ...

  13. Engineering model development and test results

    Science.gov (United States)

    Wellman, John A.

    1993-08-01

    The correctability of the primary mirror spherical error in the Wide Field/Planetary Camera (WF/PC) is sensitive to the precise alignment of the incoming aberrated beam onto the corrective elements. Articulating fold mirrors that provide +/- 1 milliradian of tilt in 2 axes are required to allow for alignment corrections in orbit as part of the fix for the Hubble space telescope. An engineering study was made by Itek Optical Systems and the Jet Propulsion Laboratory (JPL) to investigate replacement of fixed fold mirrors within the existing WF/PC optical bench with articulating mirrors. The study contract developed the base line requirements, established the suitability of lead magnesium niobate (PMN) actuators and evaluated several tilt mechanism concepts. Two engineering model articulating mirrors were produced to demonstrate the function of the tilt mechanism to provide +/- 1 milliradian of tilt, packaging within the space constraints and manufacturing techniques including the machining of the invar tilt mechanism and lightweight glass mirrors. The success of the engineering models led to the follow on design and fabrication of 3 flight mirrors that have been incorporated into the WF/PC to be placed into the Hubble Space Telescope as part of the servicing mission scheduled for late 1993.

  14. Particle Size Distributions Measured in the Stratospheric Plumes of Three Rockets During the ACCENT Missions

    Science.gov (United States)

    Wiedinmyer, C.; Brock, C. A.; Reeves, J. M.; Ross, M. N.; Schmid, O.; Toohey, D.; Wilson, J. C.

    2001-12-01

    The global impact of particles emitted by rocket engines on stratospheric ozone is not well understood, mainly due to the lack of comprehensive in situ measurements of the size distributions of these emitted particles. During the Atmospheric Chemistry of Combustion Emissions Near the Tropopause (ACCENT) missions in 1999, the NASA WB-57F aircraft carried the University of Denver N-MASS and FCAS instruments into the stratospheric plumes from three rockets. Size distributions of particles with diameters from 4 to approximately 2000 nm were calculated from the instrument measurements using numerical inversion techniques. The data have been averaged over 30-second intervals. The particle size distributions observed in all of the rocket plumes included a dominant mode near 60 nm diameter, probably composed of alumina particles. A smaller mode at approximately 25 nm, possibly composed of soot particles, was seen in only the plumes of rockets that used liquid oxygen and kerosene as a propellant. Aircraft exhaust emitted by the WB-57F was also sampled; the size distributions within these plumes are consistent with prior measurements in aircraft plumes. The size distributions for all rocket intercepts have been fitted to bimodal, lognormal distributions to provide input for global models of the stratosphere. Our data suggest that previous estimates of the solid rocket motor alumina size distributions may underestimate the alumina surface area emission index, and so underestimate the particle surface area available for heterogeneous chlorine activation reactions in the global stratosphere.

  15. Engineering model for ultrafast laser microprocessing

    Science.gov (United States)

    Audouard, E.; Mottay, E.

    2016-03-01

    Ultrafast laser micro-machining relies on complex laser-matter interaction processes, leading to a virtually athermal laser ablation. The development of industrial ultrafast laser applications benefits from a better understanding of these processes. To this end, a number of sophisticated scientific models have been developed, providing valuable insights in the physics of the interaction. Yet, from an engineering point of view, they are often difficult to use, and require a number of adjustable parameters. We present a simple engineering model for ultrafast laser processing, applied in various real life applications: percussion drilling, line engraving, and non normal incidence trepanning. The model requires only two global parameters. Analytical results are derived for single pulse percussion drilling or simple pass engraving. Simple assumptions allow to predict the effect of non normal incident beams to obtain key parameters for trepanning drilling. The model is compared to experimental data on stainless steel with a wide range of laser characteristics (time duration, repetition rate, pulse energy) and machining conditions (sample or beam speed). Ablation depth and volume ablation rate are modeled for pulse durations from 100 fs to 1 ps. Trepanning time of 5.4 s with a conicity of 0.15° is obtained for a hole of 900 μm depth and 100 μm diameter.

  16. A Flight Dynamics Model for a Multi-Actuated Flexible Rocket Vehicle

    Science.gov (United States)

    Orr, Jeb S.

    2011-01-01

    A comprehensive set of motion equations for a multi-actuated flight vehicle is presented. The dynamics are derived from a vector approach that generalizes the classical linear perturbation equations for flexible launch vehicles into a coupled three-dimensional model. The effects of nozzle and aerosurface inertial coupling, sloshing propellant, and elasticity are incorporated without restrictions on the position, orientation, or number of model elements. The present formulation is well suited to matrix implementation for large-scale linear stability and sensitivity analysis and is also shown to be extensible to nonlinear time-domain simulation through the application of a special form of Lagrange s equations in quasi-coordinates. The model is validated through frequency-domain response comparison with a high-fidelity planar implementation.

  17. Enhanced Core Noise Modeling for Turbofan Engines

    Science.gov (United States)

    Stone, James R.; Krejsa, Eugene A.; Clark, Bruce J.

    2011-01-01

    This report describes work performed by MTC Technologies (MTCT) for NASA Glenn Research Center (GRC) under Contract NAS3-00178, Task Order No. 15. MTCT previously developed a first-generation empirical model that correlates the core/combustion noise of four GE engines, the CF6, CF34, CFM56, and GE90 for General Electric (GE) under Contract No. 200-1X-14W53048, in support of GRC Contract NAS3-01135. MTCT has demonstrated in earlier noise modeling efforts that the improvement of predictive modeling is greatly enhanced by an iterative approach, so in support of NASA's Quiet Aircraft Technology Project, GRC sponsored this effort to improve the model. Since the noise data available for correlation are total engine noise spectra, it is total engine noise that must be predicted. Since the scope of this effort was not sufficient to explore fan and turbine noise, the most meaningful comparisons must be restricted to frequencies below the blade passage frequency. Below the blade passage frequency and at relatively high power settings jet noise is expected to be the dominant source, and comparisons are shown that demonstrate the accuracy of the jet noise model recently developed by MTCT for NASA under Contract NAS3-00178, Task Order No. 10. At lower power settings the core noise became most apparent, and these data corrected for the contribution of jet noise were then used to establish the characteristics of core noise. There is clearly more than one spectral range where core noise is evident, so the spectral approach developed by von Glahn and Krejsa in 1982 wherein four spectral regions overlap, was used in the GE effort. Further analysis indicates that the two higher frequency components, which are often somewhat masked by turbomachinery noise, can be treated as one component, and it is on that basis that the current model is formulated. The frequency scaling relationships are improved and are now based on combustor and core nozzle geometries. In conjunction with the Task

  18. Developing Project Duration Models in Software Engineering

    Institute of Scientific and Technical Information of China (English)

    Pierre Bourque; Serge Oligny; Alain Abran; Bertrand Fournier

    2007-01-01

    Based on the empirical analysis of data contained in the International Software Benchmarking Standards Group(ISBSG) repository, this paper presents software engineering project duration models based on project effort. Duration models are built for the entire dataset and for subsets of projects developed for personal computer, mid-range and mainframeplatforms. Duration models are also constructed for projects requiring fewer than 400 person-hours of effort and for projectsre quiring more than 400 person-hours of effort. The usefulness of adding the maximum number of assigned resources as asecond independent variable to explain duration is also analyzed. The opportunity to build duration models directly fromproject functional size in function points is investigated as well.

  19. 基于动态云BP网络的液体火箭发动机故障诊断方法%Fault diagnosis method for liquid-propellant rocket engines based on the dynamic cloud-BP neural network

    Institute of Scientific and Technical Information of China (English)

    刘垠杰; 黄强; 程玉强; 吴建军

    2012-01-01

    将云模型与BP(backpropagation)神经网络以串联方式有机结合,首先利用云变换方法进行网络的结构辨识和云模型的特征提取,同时通过在输入层引入单位延时环节描述发动机工作过程动态特性,研究提出了基于动态云BP网络的液体火箭发动机故障诊断方法.结合实际试车数据的验证结果表明,该方法能够准确识别发动机已有的3种故障模式,通过在试车数据中添加0期望、0.2标准差的随机噪声的方法来模拟环境噪声和测试过程中产生的随机噪声,根据持续性原则,方法仍能够正确进行故障检测与分类.方法单步运行时长为1.124x10-4,完全能够满足实时性要求.%A fault diagnosis method for liquid-propellant rocket engines was proposed based on the dynamic cloud-BP(back propagation) neural network in the way of the integration of cloud model and BP neural network.The Cloud transform method was used to identify the network configuration and to extract the cloud features.And a unit time-delay was also introduced into the input layer to describe the dynamic characteristics of the engine.Results with test data show that the method can isolate the existed 3 fault modes precisely.A 0 expectation,0.2 standard deviation noise was used to simulate the entironmental noise and stochastic noise,and the method can still detect and classify the fault accurately acount to lasting-rule.The method can run in real-time with the single processing time being 1.124×10-4 s.

  20. Aerodynamics and flow characterisation of multistage rockets

    Science.gov (United States)

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

    2017-05-01

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

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

    Institute of Scientific and Technical Information of China (English)

    严宇; 范玮; 王可; 穆杨

    2011-01-01

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

  2. Modeling of Heat Transfer and Ablation of Refractory Material Due to Rocket Plume Impingement

    Science.gov (United States)

    Harris, Michael F.; Vu, Bruce T.

    2012-01-01

    CR Tech's Thermal Desktop-SINDA/FLUINT software was used in the thermal analysis of a flame deflector design for Launch Complex 39B at Kennedy Space Center, Florida. The analysis of the flame deflector takes into account heat transfer due to plume impingement from expected vehicles to be launched at KSC. The heat flux from the plume was computed using computational fluid dynamics provided by Ames Research Center in Moffet Field, California. The results from the CFD solutions were mapped onto a 3-D Thermal Desktop model of the flame deflector using the boundary condition mapping capabilities in Thermal Desktop. The ablation subroutine in SINDA/FLUINT was then used to model the ablation of the refractory material.

  3. Developing engineering processes through integrated modelling of product and process

    DEFF Research Database (Denmark)

    Nielsen, Jeppe Bjerrum; Hvam, Lars

    2012-01-01

    This article aims at developing an operational tool for integrated modelling of product assortments and engineering processes in companies making customer specific products. Integrating a product model in the design of engineering processes will provide a deeper understanding of the engineering...... activities as well as insight into how product features affect the engineering processes. The article suggests possible ways of integrating models of products with models of engineering processes. The models have been tested and further developed in an action research study carried out in collaboration...

  4. Development and application of theoretical models for Rotating Detonation Engine flowfields

    Science.gov (United States)

    Fievisohn, Robert

    As turbine and rocket engine technology matures, performance increases between successive generations of engine development are becoming smaller. One means of accomplishing significant gains in thermodynamic performance and power density is to use detonation-based heat release instead of deflagration. This work is focused on developing and applying theoretical models to aid in the design and understanding of Rotating Detonation Engines (RDEs). In an RDE, a detonation wave travels circumferentially along the bottom of an annular chamber where continuous injection of fresh reactants sustains the detonation wave. RDEs are currently being designed, tested, and studied as a viable option for developing a new generation of turbine and rocket engines that make use of detonation heat release. One of the main challenges in the development of RDEs is to understand the complex flowfield inside the annular chamber. While simplified models are desirable for obtaining timely performance estimates for design analysis, one-dimensional models may not be adequate as they do not provide flow structure information. In this work, a two-dimensional physics-based model is developed, which is capable of modeling the curved oblique shock wave, exit swirl, counter-flow, detonation inclination, and varying pressure along the inflow boundary. This is accomplished by using a combination of shock-expansion theory, Chapman-Jouguet detonation theory, the Method of Characteristics (MOC), and other compressible flow equations to create a shock-fitted numerical algorithm and generate an RDE flowfield. This novel approach provides a numerically efficient model that can provide performance estimates as well as details of the large-scale flow structures in seconds on a personal computer. Results from this model are validated against high-fidelity numerical simulations that may require a high-performance computing framework to provide similar performance estimates. This work provides a designer a new

  5. Influence of gunpowder start system on starting performance of liquid rocket engines%火药起动系统对发动机起动性能的影响分析

    Institute of Scientific and Technical Information of China (English)

    孙海雨; 刘志让

    2012-01-01

    Aiming at the pumping pressure open cycle liquid rocket engine which is started by solid start cartridge (SSC), the performance of the start system is studied in this paper. The simulated model of SSC in the start system and the numerical model of the powder gas pipeline were established to simulate the process of gunpowder start. The influence of SSC and gas pipeline parameters on start performance of engine is analyzed to ensure the main influence parameters and regular patterns. It is found that the powder quantity of SSC and the diameter of the first throat in the powder gas pipeline are the most effective factors to the engine's start characteristic. The diameter of the powder gas pipeline's second throat and the diameter of the powder gas pipeline's outlet are the least ones in the case of that the powder gas pipeline's flow field keeps rated condition. The simulated result of the start system was proven in engine hot tests.%针对采用火药起动器起动的泵压开式循环液体火箭发动机,对其起动系统进行了分析和研究。建立了液体火箭发动机火药起动器计算模型和起动系统燃气管路流场计算模型。将所建立的起动系统模型应用于发动机系统仿真,对发动机火药起动过程进行仿真,分析了起动系统中火药起动器参数和燃气管路参数对发动机起动性能的影响,确定了主要影响参数和影响规律。火药起动器火药药柱内径、火药药柱长度以及燃气管路火药起动器喷管喉部直径为强影响因素;燃气管路涡轮喷嘴喉部直径和管路出口直径在确保发动机火药起动主要工况段燃气管路流场流态为额定工况流态的前提下,为弱影响因素。试验数据验证表明,发动机起动系统的仿真结果正确、可信。

  6. Investigation of different modeling approaches for computational fluid dynamics simulation of high-pressure rocket combustors

    Science.gov (United States)

    Ivancic, B.; Riedmann, H.; Frey, M.; Knab, O.; Karl, S.; Hannemann, K.

    2016-07-01

    The paper summarizes technical results and first highlights of the cooperation between DLR and Airbus Defence and Space (DS) within the work package "CFD Modeling of Combustion Chamber Processes" conducted in the frame of the Propulsion 2020 Project. Within the addressed work package, DLR Göttingen and Airbus DS Ottobrunn have identified several test cases where adequate test data are available and which can be used for proper validation of the computational fluid dynamics (CFD) tools. In this paper, the first test case, the Penn State chamber (RCM1), is discussed. Presenting the simulation results from three different tools, it is shown that the test case can be computed properly with steady-state Reynolds-averaged Navier-Stokes (RANS) approaches. The achieved simulation results reproduce the measured wall heat flux as an important validation parameter very well but also reveal some inconsistencies in the test data which are addressed in this paper.

  7. Academic program models for undergraduate biomedical engineering.

    Science.gov (United States)

    Krishnan, Shankar M

    2014-01-01

    There is a proliferation of medical devices across the globe for the diagnosis and therapy of diseases. Biomedical engineering (BME) plays a significant role in healthcare and advancing medical technologies thus creating a substantial demand for biomedical engineers at undergraduate and graduate levels. There has been a surge in undergraduate programs due to increasing demands from the biomedical industries to cover many of their segments from bench to bedside. With the requirement of multidisciplinary training within allottable duration, it is indeed a challenge to design a comprehensive standardized undergraduate BME program to suit the needs of educators across the globe. This paper's objective is to describe three major models of undergraduate BME programs and their curricular requirements, with relevant recommendations to be applicable in institutions of higher education located in varied resource settings. Model 1 is based on programs to be offered in large research-intensive universities with multiple focus areas. The focus areas depend on the institution's research expertise and training mission. Model 2 has basic segments similar to those of Model 1, but the focus areas are limited due to resource constraints. In this model, co-op/internship in hospitals or medical companies is included which prepares the graduates for the work place. In Model 3, students are trained to earn an Associate Degree in the initial two years and they are trained for two more years to be BME's or BME Technologists. This model is well suited for the resource-poor countries. All three models must be designed to meet applicable accreditation requirements. The challenges in designing undergraduate BME programs include manpower, facility and funding resource requirements and time constraints. Each academic institution has to carefully analyze its short term and long term requirements. In conclusion, three models for BME programs are described based on large universities, colleges, and

  8. Integrated approach for hybrid rocket technology development

    Science.gov (United States)

    Barato, Francesco; Bellomo, Nicolas; Pavarin, Daniele

    2016-11-01

    Hybrid rocket motors tend generally to be simple from a mechanical point of view but difficult to optimize because of their complex and still not well understood cross-coupled physics. This paper addresses the previous issue presenting the integrated approach established at University of Padua to develop hybrid rocket based systems. The methodology tightly combines together system analysis and design, numerical modeling from elementary to sophisticated CFD, and experimental testing done with incremental philosophy. As an example of the approach, the paper presents the experience done in the successful development of a hybrid rocket booster designed for rocket assisted take off operations. It is thought that following the proposed approach and selecting carefully the most promising applications it is possible to finally exploit the major advantages of hybrid rocket motors as safety, simplicity, low cost and reliability.

  9. Loss terms in free-piston Stirling engine models

    Science.gov (United States)

    Gordon, Lloyd B.

    1992-01-01

    Various models for free piston Stirling engines are reviewed. Initial models were developed primarily for design purposes and to predict operating parameters, especially efficiency. More recently, however, such models have been used to predict engine stability. Free piston Stirling engines have no kinematic constraints and stability may not only be sensitive to the load, but also to various nonlinear loss and spring constraints. The present understanding is reviewed of various loss mechanisms for free piston Stirling engines and how they have been incorporated into engine models is discussed.

  10. Creating system engineering products with executable models in a model-based engineering environment

    Science.gov (United States)

    Karban, Robert; Dekens, Frank G.; Herzig, Sebastian; Elaasar, Maged; Jankevičius, Nerijus

    2016-08-01

    Applying systems engineering across the life-cycle results in a number of products built from interdependent sources of information using different kinds of system level analysis. This paper focuses on leveraging the Executable System Engineering Method (ESEM) [1] [2], which automates requirements verification (e.g. power and mass budget margins and duration analysis of operational modes) using executable SysML [3] models. The particular value proposition is to integrate requirements, and executable behavior and performance models for certain types of system level analysis. The models are created with modeling patterns that involve structural, behavioral and parametric diagrams, and are managed by an open source Model Based Engineering Environment (named OpenMBEE [4]). This paper demonstrates how the ESEM is applied in conjunction with OpenMBEE to create key engineering products (e.g. operational concept document) for the Alignment and Phasing System (APS) within the Thirty Meter Telescope (TMT) project [5], which is under development by the TMT International Observatory (TIO) [5].

  11. Journal of Modeling, Design and Management of Engineering ...

    African Journals Online (AJOL)

    It has special focus on the application of physical or mathematical modeling, computing, simulation, design and/or ... Power systems Production/Manufacturing systems Process engineering systems ... Department of Mechanical Engineering,

  12. Numerical Modeling of Fluid Transient in Cryogenic Fluid Network of Rocket Propulsion System

    Science.gov (United States)

    Majumdar, Alok; Flachbart, Robin

    2003-01-01

    Fluid transients, also known as water hammer, can have a significant impact on the design and operation of both spacecraft and launch vehicles propulsion systems. These transients often occur at system activation and shut down. For ground safety reasons, many spacecrafts are launched with the propellant lines dry. These lines are often evacuated by the time the spacecraft reaches orbit. When the propellant isolation valve opens during propulsion system activation, propellant rushes into lines creating a pressure surge. During propellant system shutdown, a pressure surge is created due to sudden closure of a valve. During both activation and shutdown, pressure surges must be predicted accurately to ensure structural integrity of the propulsion system fluid network. The method of characteristics is the most widely used method of calculating fluid transients in pipeline [ 1,2]. The method of characteristics, however, has limited applications in calculating flow distribution in complex flow circuits with phase change, heat transfer and rotational effects. A robust cryogenic propulsion system analyzer must have the capability to handle phase change, heat transfer, chemical reaction, rotational effects and fluid transients in conjunction with subsystem flow model for pumps, valves and various pipe fittings. In recent years, such a task has been undertaken at Marshall Space Flight Center with the development of the Generalized Fluid System Simulation Program (GFSSP), which is based on finite volume method in fluid network [3]. GFSSP has been extensively verified and validated by comparing its predictions with test data and other numerical methods for various applications such as internal flow of turbo-pump [4], propellant tank pressurization [5,6], chilldown of cryogenic transfer line [7] and squeeze film damper rotordynamics [8]. The purpose of the present paper is to investigate the applicability of the finite volume method to predict fluid transient in cryogenic flow

  13. Mars 2020 Model Based Systems Engineering Pilot

    Science.gov (United States)

    Dukes, Alexandra Marie

    2017-01-01

    The pilot study is led by the Integration Engineering group in NASA's Launch Services Program (LSP). The Integration Engineering (IE) group is responsible for managing the interfaces between the spacecraft and launch vehicle. This pilot investigates the utility of Model-Based Systems Engineering (MBSE) with respect to managing and verifying interface requirements. The main objectives of the pilot are to model several key aspects of the Mars 2020 integrated operations and interface requirements based on the design and verification artifacts from Mars Science Laboratory (MSL) and to demonstrate how MBSE could be used by LSP to gain further insight on the interface between the spacecraft and launch vehicle as well as to enhance how LSP manages the launch service. The method used to accomplish this pilot started through familiarization of SysML, MagicDraw, and the Mars 2020 and MSL systems through books, tutorials, and NASA documentation. MSL was chosen as the focus of the model since its processes and verifications translate easily to the Mars 2020 mission. The study was further focused by modeling specialized systems and processes within MSL in order to demonstrate the utility of MBSE for the rest of the mission. The systems chosen were the In-Flight Disconnect (IFD) system and the Mass Properties process. The IFD was chosen as a system of focus since it is an interface between the spacecraft and launch vehicle which can demonstrate the usefulness of MBSE from a system perspective. The Mass Properties process was chosen as a process of focus since the verifications for mass properties occur throughout the lifecycle and can demonstrate the usefulness of MBSE from a multi-discipline perspective. Several iterations of both perspectives have been modeled and evaluated. While the pilot study will continue for another 2 weeks, pros and cons of using MBSE for LSP IE have been identified. A pro of using MBSE includes an integrated view of the disciplines, requirements, and

  14. The flight of uncontrolled rockets

    CERN Document Server

    Gantmakher, F R; Dryden, H L

    1964-01-01

    International Series of Monographs on Aeronautics and Astronautics, Division VII, Volume 5: The Flight of Uncontrolled Rockets focuses on external ballistics of uncontrolled rockets. The book first discusses the equations of motion of rockets. The rocket as a system of changing composition; application of solidification principle to rockets; rotational motion of rockets; and equations of motion of the center of mass of rockets are described. The text looks at the calculation of trajectory of rockets and the fundamentals of rocket dispersion. The selection further focuses on the dispersion of f

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

    Institute of Scientific and Technical Information of China (English)

    马兵兵; 翟丽婷; 孙璐

    2012-01-01

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

  16. Model driven product line engineering : core asset and process implications

    OpenAIRE

    Azanza Sesé, Maider

    2011-01-01

    Reuse is at the heart of major improvements in productivity and quality in Software Engineering. Both Model Driven Engineering (MDE) and Software Product Line Engineering (SPLE) are software development paradigms that promote reuse. Specifically, they promote systematic reuse and a departure from craftsmanship towards an industrialization of the software development process. MDE and SPLE have established their benefits separately. Their combination, here called Model Driven Product Line Engin...

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

    Science.gov (United States)

    Zibit, Benjamin Seth

    creation of the jet Propulsion Laboratory, the founding of the Aerojet Corporation, and emphasizes the issue of JPL's close relation to military development of the rocket becomes a core subject of this thesis. Cooperation between engineers in an academic setting and the military was not merely inevitable in the 1940s---it was actively fostered and proved quite profitable to all concerned. The deep relationship between the Guggenheim Aeronautics Laboratory and the Army Air Force was one model of the evolution of a permanent institutional edifice, weaving academic research and military end-use together. The dissertation concludes that what began as a modest effort to understand rocket theory in greater depth led within ten years to both research and development tracks which have profoundly altered the technological and military definition of modern history.

  18. Hybrid Rocket Technology

    National Research Council Canada - National Science Library

    Sankaran Venugopal; K K Rajesh; V Ramanujachari

    2011-01-01

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

  19. Numerical investigations of thermal stratification in cooling channel of liquid rocket engine thrust chamber%液体火箭发动机推力室冷却通道温度分层数值研究

    Institute of Scientific and Technical Information of China (English)

    康玉东; 孙冰; 高翔宇

    2009-01-01

    为了研究冷却剂温度分层的形成机理及其对流动和换热的影响,应用雷诺应力模型(RSM)对液体火箭发动机推力室再生冷却通道的流动与传热进行了三维数值模拟,冷却剂为气氢,考虑其物性随温度和压力的变化.所得结果表明:冷却剂在非流动方向会出现温度分层现象,随着冷却剂的不断受热,温度分层现象越明显,由于喉部二次流加强了冷却剂间的混合,在喉部区域温度分层被减弱,温度分层对冷却剂温升及压降影响较小,严重影响气壁温度的估算.%To study the formation mechanism of thermal stratification in cooling channel and its effects on the flow and heat transfer,three dimensional turbulent fluid flow and heat transfer in a regenerative-cooling channel of liquid rocket engine were numerically investigated with Reynolds stress model(RSM) model,and the coolant was hydrogen,whose thermo-physical properties varied with both temperature and pressure.The results show that thermal stratification occurs at non-flow direction,the extent of thermal stratification becomes increasingly significant as the extent of heating increases,and the thermal stratification of coolant is weakened for the existence of secondary flow in throat region; the thermal stratification has little effect on the bulk temperature increase and hydrodynamic losses of hydrogen,but has significant effect on the calculation of the wall heat fluxes and temperature.

  20. Cognitive engineering models in space systems

    Science.gov (United States)

    Mitchell, Christine M.

    1993-01-01

    , the PDRS was identified as the most accessible system for the demonstration. Pursuant to this a PDRS simulation was obtained from the HCIL and an initial knowledge engineering effort was conducted to understand the operator's tasks in the PDRS application. The preliminary results of the knowledge engineering effort and an initial formulation of an operator function model (OFM) are contained in the appendices.