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Sample records for rocket engine nozzles

  1. Rocket engine nozzle attenuator

    NASA Astrophysics Data System (ADS)

    Lewis, David A.

    1993-01-01

    The function of a rocket engine nozzle is to expand the hot engine exhaust gases down to ambient pressure, transforming thermal energy to directed kinetic energy in order to produce thrust. Considering nozzle design, there is an optimum nozzle shape and length, the bell-shaped or contour nozzle. The reason for this specific contour is that the nozzle must be designed in such a manner that the expansion shock waves emanating from the nozzle throat region coincide, and thus diminish the compression effects accompanying the reorientation of flow in the center region of the expansion section. A rocket nozzle must absorb a variety of loads caused by such shocks due to thermal expansion and contraction, as well as shocks from sudden pressurization at startup, and flight accelerations. A rocket engine nozzle is provided which is capable of attenuating nozzle vibrations generated therein during use. The nozzle includes an annular closed chamber surrounding the nozzle adjacent to its gas exhaust end. Within the chamber is a dense but unrestricted particulate mass capable of undergoing frictional movement within the chamber.

  2. Liquid rocket engine nozzles

    NASA Technical Reports Server (NTRS)

    1976-01-01

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

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

    NASA Astrophysics Data System (ADS)

    Forde, Scott; Bulman, Mel; Neill, Todd

    2006-07-01

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

  4. Comparison of two procedures for predicting rocket engine nozzle performance

    NASA Technical Reports Server (NTRS)

    Davidian, Kenneth J.

    1987-01-01

    Two nozzle performance prediction procedures which are based on the standardized JANNAF methodology are presented and compared for four rocket engine nozzles. The first procedure required operator intercedence to transfer data between the individual performance programs. The second procedure is more automated in that all necessary programs are collected into a single computer code, thereby eliminating the need for data reformatting. Results from both procedures show similar trends but quantitative differences. Agreement was best in the predictions of specific impulse and local skin friction coefficient. Other compared quantities include characteristic velocity, thrust coefficient, thrust decrement, boundary layer displacement thickness, momentum thickness, and heat loss rate to the wall. Effects of wall temperature profile used as an input to the programs was investigated by running three wall temperature profiles. It was found that this change greatly affected the boundary layer displacement thickness and heat loss to the wall. The other quantities, however, were not drastically affected by the wall temperature profile change.

  5. Annular nozzle engine technology

    NASA Technical Reports Server (NTRS)

    Martinez, AL

    1992-01-01

    The topics covered include: (1) driver rocket subsystem; (2) annular nozzle engine technology; (3) expansion-deflection nozzle; (4) aerospike-nozzled engine background; (5) aerospike testing; (6) linear aerospike; and (7) the combined cycle engine.

  6. High-speed measurement of nozzle swing angle of rocket engine based on monocular vision

    NASA Astrophysics Data System (ADS)

    Qu, Yufu; Yang, Haijuan

    2015-02-01

    A nozzle angle measurement system based on monocular vision is proposed to achieve high-speed and non-contact angle measurement of rocket engine nozzle. The measurement system consists of two illumination sources, a lens, a target board with spots, a high-speed camera, an image acquisition card and a PC. A target board with spots was fixed on the end of rocket engine nozzle. The image of the target board moved along with the rocket engine nozzle swing was captured by a high-speed camera and transferred to the PC by an image acquisition card. Then a data processing algorithm was utilized to acquire the swing angle of the engine nozzle. Experiment shows that the accuracy of swing angle measurement was 0.2° and the measurement frequency was up to 500Hz.

  7. Rocket nozzle snubber

    NASA Astrophysics Data System (ADS)

    Hoever, Jeffrey P.; Reginato, Richard L.; Bay, Frederick M.; Rolon, Claudia E.; Bazigos, Michael

    1991-10-01

    Stress from the ignition of a rocket is transferred from the rocket nozzle to the actuators and flexible joint seal which move the nozzle causing potential failure of the actuators. This stress can be relieved by using a snubber in the form of a flexible hollow conical toroidal bladder filled with incompressible fluid fitted snugly between the nozzle and the main body of the rocket. Such a snubber transfers the stress to the main body of the rocket through the snubber while deforming to allow the rocket nozzle to slew in response to control commands.

  8. Modification of the PARC Navier-Stokes Code to predict rocket engine nozzle performance

    NASA Technical Reports Server (NTRS)

    Collins, Frank G.; Myruski, Bryan; Orr, Joseph L.

    1990-01-01

    The PARC2D Navier-Stokes Code was modified to compute the performance parameters for rocket engine nozzles. The perfect gas code was applied to the SSME engine nozzle for inviscid, laminar and turbulent flow. Inviscid computations compare well with Rocketdyne computations. Performance degradation due to the boundary layers is very reasonable. Application of the code to nontraditional nozzle geometries and to low Reynolds nozzles is demonstrated. Modification of the code for equilibrium H2/O2 chemistry is described. Thermodynamic and equilibrium constants are determined from statistical mechanics and the transport properties from exact kinetic theory, using collison integrals determined from appropriate intermolecular potentials. The equilibrium code was used to compute the SSME flowfield. Modifications of the flowfield due to the change of composition are described.

  9. Modification of the PARC Navier-Stokes Code to predict rocket engine nozzle performance

    NASA Technical Reports Server (NTRS)

    Collins, Frank G.; Myruski, Bryan; Orr, Joseph L.

    1990-01-01

    The PARC2D Navier-Stokes Code was modified to compute the performance parameters for rocket engine nozzles. The perfect gas code was applied to the SSME engine nozzle for inviscid, laminar and turbulent flow. Inviscid computations compare well with Rocketdyne computations. Performance degradation due to the boundary layers is very reasonable. Application of the code to nontraditional nozzle geometries and to low Reynolds nozzles is demonstrated. Modification of the code for equilibrium H2/O2 chemistry is described. Thermodynamic and equilibrium constants are determined from statistical mechanics and the transport properties from exact kinetic theory, using collison integrals determined from appropriate intermolecular potentials. The equilibrium code was used to compute the SSME flowfield. Modifications of the flowfield due to the change of composition are described.

  10. Heat transfer in rocket engine combustion chambers and nozzles

    NASA Technical Reports Server (NTRS)

    Anderson, P. G.; Chen, Y. S.; Farmer, R. C.

    1992-01-01

    The 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 analysis verified with appropriate test data. Finally, the component analysis 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. Film cooling simulations of film coolant flows typical of the subscale Space Transportation Main Engine (STME) being experimentally studied by Pratt and Whitney have been made, and these results will be presented. Other film coolant experiments have also been simulated to verify the CFD heat transfer model being developed. The status of the study and its relevance as a new design tool are covered. Information is given in viewgraph form.

  11. Heat transfer in rocket engine combustion chambers and nozzles

    NASA Technical Reports Server (NTRS)

    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.

  12. Heat transfer in rocket engine combustion chambers and nozzles

    NASA Astrophysics Data System (ADS)

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

    1993-07-01

    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.

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

    NASA Technical Reports Server (NTRS)

    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.

  14. Experimental Performance of Area Ratio 200, 25 and 8 Nozzles on JP-4 Fuel and Liquid Oxygen Rocket Engine

    NASA Technical Reports Server (NTRS)

    Lovell, J. Calvin; Samanich, Nick E.; Barnett, Donald O.

    1960-01-01

    The performance of an area ratio 200 bell-shaped nozzle, an area ratio 25 bell-shaped nozzle, and an area ratio 8 conic nozzle on a JP-4 fuel and liquid-oxygen rocket engine has been determined. Tests were conducted using a nominal 4000-pound-thrust rocket in the Lewis 10- by 10-foot supersonic tunnel, which provided the altitude environment needed for fully expanded nozzle flow. The area ratio 200 nozzle had a vacuum thrust coefficient of 1.96, compared with 1.82 and 1.70 for the area ratio 25 and 8 nozzles, respectively. These values are approximately equal to those for theoretical frozen expansion. The measured value of vacuum specific impulse for the area ratio 200 nozzle was 317 seconds for a combustion-chamber characteristic velocity of 5200 feet per second. The vacuum-specific-impulse increase for the area-ratio increase from 8 to 200 was 46 seconds.

  15. Rocket nozzle shield

    NASA Astrophysics Data System (ADS)

    Vollersen, Carl A.; Anderson, Thomas W.; Brown, Wayne M.; Pratt, Lozelle L.

    1991-12-01

    A shield for the nozzle of an underwater launched missile can be formed from a rigid plate. The plate is separated from the end of the nozzle by a gap that is at least equal to the throat area of the nozzle. The gap reduces overpressure of the missile during launch while keeping water out of the nozzle by being directed perpendicular to the water flow direction. Important advantages of the invention are its simplicity and low cost. A further advantage is its strength and the ability to absorb a blow from a water jar. Another advantage is the ability to block the flow of water from behind the rocket without blocking the flow of gas.

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

    NASA Technical Reports Server (NTRS)

    Shi, John J.

    2005-01-01

    During the development stage, in order to design/to size the rocket engine components and to reduce the risks, the local dynamic environments as well as dynamic interface loads must be defined. There are two kinds of dynamic environment, i.e. shock transients and steady-state random and sinusoidal vibration environments. Usually, the steady-state random and sinusoidal vibration environments are scalable, but the shock environments are not scalable. In other words, based on similarities only random vibration environments can be defined for a new engine. The methodology covered in this paper provides a way to predict the shock environments and the dynamic loads for new engine systems and new engine components in the early stage of new engine development or engine nozzle modifications.

  17. Experimental investigations on pulse detonation rocket engine with various injectors and nozzles

    NASA Astrophysics Data System (ADS)

    Yan, Yu; Fan, Wei; Wang, Ke; Zhu, Xu-dong; Mu, Yang

    2011-07-01

    Pulse detonation engines (PDEs) may represent a revolutionary approach to propulsion. The engine of simple construction can be easily manufactured. The pulse detonation rocket engine (PDRE) used here are 30 mm in inner diameter and 860 mm in length. Liquid kerosene, gaseous oxygen and nitrogen were used as fuel, oxidizer and purge gas, respectively. Two-phase detonation generating is harder than gaseous detonation due to liquid fuel atomization and mixing of two-phase reactants. It is a difficult task for liquid fuel and gaseous oxidizer to mix and form uniformly distributed mixture in the entire long engine during filling process in a short time. Therefore the velocities of fuel and oxidizer must be well designed to achieve not only the requirement of filling the entire engine but also the requirement of liquid fuel atomization and reactants mixing. Four injectors were tested to improve the atomization of liquid fuel and mixing process of reactants for performance enhancement of PDRE. Injector with small fuel exit area and large gas exit area was found to be effective for liquid fuel atomization and reactants mixing process. The PDRE with injector B performed the best among all the injectors tested. Nozzles are critical components in improving the performance of PDRE. Four kinds of bell-shaped converging-diverging nozzles were also tested here in order to enhance the performance of PDRE. It was found that a nozzle with high contraction ratio and high expansion ratio generated the highest thrust augmentation of 27.3%.

  18. Investigation of Ceramic, Graphite, and Chrome-plated Graphite Nozzles on Rocket Engine

    NASA Technical Reports Server (NTRS)

    Kinney, George R; Lidman, William G

    1949-01-01

    The use of ceramic material for rocket nozzles and the effectiveness of preventing oxidation and erosion of graphite nozzles by chrome-plating the internal surface were investigated. A supported ceramic nozzle, cracked by initial operation, was operated a second time without further cracking or damage. Chrome-plating the internal surface of graphite nozzles effectively prevented oxidation and erosion that occurred during operation with unprotected graphite.

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

    NASA Technical Reports Server (NTRS)

    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.

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

    NASA Technical Reports Server (NTRS)

    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.

  1. Subscale Carbon-Carbon Nozzle Extension Development and Hot Fire Testing in Support of Upper Stage Liquid Rocket Engines

    NASA Technical Reports Server (NTRS)

    Gradl, Paul; Valentine, Peter; Crisanti, Matthew; Greene, Sandy Elam

    2016-01-01

    Upper stage and in-space liquid rocket engines are optimized for performance through the use of high area ratio nozzles to fully expand combustion gases to low exit pressures increasing exhaust velocities. Due to the large size of such nozzles and the related engine performance requirements, carbon-carbon (C/C) composite nozzle extensions are being considered for use in order to reduce weight impacts. NASA and industry partner Carbon-Carbon Advanced Technologies (C-CAT) are working towards advancing the technology readiness level of large-scale, domestically-fabricated, C/C nozzle extensions. These C/C extensions have the ability to reduce the overall costs of extensions relative to heritage metallic and composite extensions and to decrease weight by 50%. Material process and coating developments have advanced over the last several years, but hot fire testing to fully evaluate C/C nozzle extensions in relevant environments has been very limited. NASA and C-CAT have designed, fabricated and hot fire tested multiple subscale nozzle extension test articles of various C/C material systems, with the goal of assessing and advancing the manufacturability of these domestically producible materials as well as characterizing their performance when subjected to the typical environments found in a variety of liquid rocket and scramjet engines. Testing at the MSFC Test Stand 115 evaluated heritage and state-of-the-art C/C materials and coatings, demonstrating the capabilities of the high temperature materials and their fabrication methods. This paper discusses the design and fabrication of the 1.2k-lbf sized carbon-carbon nozzle extensions, provides an overview of the test campaign, presents results of the hot fire testing, and discusses potential follow-on development work.

  2. Carbon-Carbon Nozzle Extension Development in Support of In-Space and Upper-Stage Liquid Rocket Engines

    NASA Technical Reports Server (NTRS)

    Gradl, Paul R.; Valentine, Peter G.

    2017-01-01

    Upper stage and in-space liquid rocket engines are optimized for performance through the use of high area ratio nozzles to fully expand combustion gases to low exit pressures, increasing exhaust velocities. Due to the large size of such nozzles, and the related engine performance requirements, carbon-carbon (C-C) composite nozzle extensions are being considered to reduce weight impacts. Currently, the state-of-the-art is represented by the metallic and foreign composite nozzle extensions limited to approximately 2000 degrees F. used on the Atlas V, Delta IV, Falcon 9, and Ariane 5 launch vehicles. NASA and industry partners are working towards advancing the domestic supply chain for C-C composite nozzle extensions. These development efforts are primarily being conducted through the NASA Small Business Innovation Research (SBIR) program in addition to other low level internal research efforts. This has allowed for the initial material development and characterization, subscale hardware fabrication, and completion of hot-fire testing in relevant environments. NASA and industry partners have designed, fabricated and hot-fire tested several subscale domestically produced C-C extensions to advance the material and coatings fabrication technology for use with a variety of liquid rocket and scramjet engines. Testing at NASA's Marshall Space Flight Center (MSFC) evaluated heritage and state-of-the-art C-C materials and coatings, demonstrating the initial capabilities of the high temperature materials and their fabrication methods. This paper discusses the initial material development, design and fabrication of the subscale carbon-carbon nozzle extensions, provides an overview of the test campaign, presents results of the hot fire testing, and discusses potential follow-on development work. The follow on work includes the fabrication of ultra-high temperature materials, larger C-C nozzle extensions, material characterization, sub-element testing and hot-fire testing at

  3. Development and Testing of Carbon-Carbon Nozzle Extensions for Upper Stage Liquid Rocket Engines

    NASA Technical Reports Server (NTRS)

    Valentine, Peter G.; Gradl, Paul R.; Greene, Sandra E.

    2017-01-01

    Carbon-carbon (C-C) composite nozzle extensions are of interest for use on a variety of launch vehicle upper stage engines and in-space propulsion systems. The C-C nozzle extension technology and test capabilities being developed are intended to support National Aeronautics and Space Administration (NASA) and Department of Defense (DOD) requirements, as well as those of the broader Commercial Space industry. For NASA, C-C nozzle extension technology development primarily supports the NASA Space Launch System (SLS) and NASA's Commercial Space partners. Marshall Space Flight Center (MSFC) efforts are aimed at both (a) further developing the technology and databases needed to enable the use of composite nozzle extensions on cryogenic upper stage engines, and (b) developing and demonstrating low-cost capabilities for testing and qualifying composite nozzle extensions. Recent, on-going, and potential future work supporting NASA, DOD, and Commercial Space needs will be discussed. Information to be presented will include (a) recent and on-going mechanical, thermal, and hot-fire testing, as well as (b) potential future efforts to further develop and qualify domestic C-C nozzle extension solutions for the various upper stage engines under development.

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

    NASA Technical Reports Server (NTRS)

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

    2010-01-01

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

  5. Nitrous oxide cooling in hybrid rocket nozzles

    NASA Astrophysics Data System (ADS)

    Lemieux, Patrick

    2010-02-01

    The Department of Mechanical Engineering at the California Polytechnic State University, San Luis Obispo, has developed an innovative program of experimental research and development on hybrid rocket motors (where the fuel and the oxidizer are in different phases prior to combustion). One project currently underway involves the development of aerospike nozzles for such motors. These nozzles, however, are even more susceptible to throat ablation than regular converging-diverging nozzles, due the nature of their flow expansion mechanism. This paper presents the result of a recent development project focused on reducing throat ablation in hybrid rocket motor nozzles. Although the method is specifically targeted at increasing the life and operating range of aerospike nozzles, this paper describes its proof-of-concept implementation on conventional nozzles. The method is based on a regenerative cooling mechanism that differs in practice from that used in liquid propellant motors. A series of experimental tests demonstrate that this new method is not only effective at reducing damage in the most ablative region of the nozzle, but that the nozzle can survive multiple test runs.

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

    NASA Technical Reports Server (NTRS)

    Kacynski, Kenneth J.; Hoffman, Joe D.

    1993-01-01

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

  7. Dual nozzle design update. [on liquid rocket engines for advanced earth-to-orbit transportation systems

    NASA Technical Reports Server (NTRS)

    Obrien, C. J.

    1982-01-01

    Dual-nozzle engines, such as the dual-throat and dual-expander engines, are being evaluated for advanced earth-to-orbit transportation systems. Potential derivatives of the Space Shuttle and completely new vehicles might benefit from these advanced engines. In this paper, progress in the design of single-fuel and dual-fuel dual-nozzle engines is summarized. Dual-nozzle engines include those burning propellants such as LOX/RP-1/LH2, LOX/LC3H8/LH2, LOX/LCH4/LH2, LOX/LH2/LH2, LOX/LCH4/LCH4, LOX/LC3H8/C3H8 and N2O4/MMH/LH2. Engine data are applicable for thrust levels from 200,000 through 670,000 lbF. The results indicate that several versions of these engines utilize state-of-the-art technology and that even advanced versions of these engines do not require a major breakthrough in technology.

  8. Transient Two-Dimensional Analysis of Side Load in Liquid Rocket Engine Nozzles

    NASA Technical Reports Server (NTRS)

    Wang, Ten-See

    2004-01-01

    Two-dimensional planar and axisymmetric numerical investigations on the nozzle start-up side load physics were performed. The objective of this study is to develop a computational methodology to identify nozzle side load physics using simplified two-dimensional geometries, in order to come up with a computational strategy to eventually predict the three-dimensional side loads. The computational methodology is based on a multidimensional, finite-volume, viscous, chemically reacting, unstructured-grid, and pressure-based computational fluid dynamics formulation, and a transient inlet condition based on an engine system modeling. The side load physics captured in the low aspect-ratio, two-dimensional planar nozzle include the Coanda effect, afterburning wave, and the associated lip free-shock oscillation. Results of parametric studies indicate that equivalence ratio, combustion and ramp rate affect the side load physics. The side load physics inferred in the high aspect-ratio, axisymmetric nozzle study include the afterburning wave; transition from free-shock to restricted-shock separation, reverting back to free-shock separation, and transforming to restricted-shock separation again; and lip restricted-shock oscillation. The Mach disk loci and wall pressure history studies reconfirm that combustion and the associated thermodynamic properties affect the formation and duration of the asymmetric flow.

  9. Flow separation in rocket nozzles under high altitude condition

    NASA Astrophysics Data System (ADS)

    Stark, R.; Génin, C.

    2017-01-01

    The knowledge of flow separation in rocket nozzles is crucial for rocket engine design and optimum performance. Typically, flow separation is studied under sea-level conditions. However, this disregards the change of the ambient density during ascent of a launcher. The ambient flow properties are an important factor concerning the design of altitude-adaptive rocket nozzles like the dual bell nozzle. For this reason an experimental study was carried out to study the influence of the ambient density on flow separation within conventional nozzles.

  10. Evaluation of coated columbium test panels having application to a secondary nozzle extension for the RL10 rocket engine system, parts 1 and 2

    NASA Technical Reports Server (NTRS)

    Murphy, Kenneth S.; Castro, Joaquin H.

    1988-01-01

    The activity performed on the screening and evaluation of various coatings for application on columbium alloy test panels representative of a radiation-cooled nozzle extension for the RL10 rocket engine is summarized. Vendors and processes of candidate coatings were evaluated. Post engine test evaluations of the two selected coatings are discussed.

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

    NASA Technical Reports Server (NTRS)

    Kacynski, Kenneth J.; Hoffman, Joe D.

    1994-01-01

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

  12. Flow Separation Side Loads Excitation of Rocket Nozzle FEM

    NASA Technical Reports Server (NTRS)

    Smalley, Kurt B.; Brown, Andrew; Ruf, Joseph; Gilbert, John

    2007-01-01

    Modern rocket nozzles are designed to operate over a wide range of altitudes, and are also built with large aspect ratios to enable high efficiencies. Nozzles designed to operate over specific regions of a trajectory are being replaced in modern launch vehicles by those that are designed to operate from earth to orbit. This is happening in parallel with modern manufacturing and wall cooling techniques allowing for larger aspect ratio nozzles to be produced. Such nozzles, though operating over a large range of altitudes and ambient pressures, are typically designed for one specific altitude. Above that altitude the nozzle flow is 'underexpanded' and below that altitude, the nozzle flow is 'overexpanded'. In both conditions the nozzle produces less than the maximum possible thrust at that altitude. Usually the nozzle design altitude is well above sea level, leaving the nozzle flow in an overexpanded state for its start up as well as for its ground testing where, if it is a reusable nozzle such as the Space Shuttle Main Engine (SSME), the nozzle will operate for the majority of its life. Overexpansion in a rocket nozzle presents the critical, and sometimes design driving, problem of flow separation induced side loads. To increase their understanding of nozzle side loads, engineers at MSFC began an investigation in 2000 into the phenomenon through a task entitled "Characterization and Accurate Modeling of Rocket Engine Nozzle Side Loads", led by A. Brown. The stated objective of this study was to develop a methodology to accurately predict the character and magnitude of nozzle side loads. The study included further hot-fire testing of the MC-l engine, cold flow testing of subscale nozzles, CFD analyses of both hot-fire and cold flow nozzle testing, and finite element (fe.) analysis of the MC-1 engine and cold flow tested nozzles. A follow on task included an effort to formulate a simplified methodology for modeling a side load during a two nodal diameter fluid

  13. Transient Three-Dimensional Analysis of Side Load in Liquid Rocket Engine Nozzles

    NASA Technical Reports Server (NTRS)

    Wang, Ten-See

    2004-01-01

    Three-dimensional numerical investigations on the nozzle start-up side load physics were performed. The objective of this study is to identify the three-dimensional side load physics and to compute the associated aerodynamic side load using an anchored computational methodology. The computational methodology is based on an unstructured-grid, and pressure-based computational fluid dynamics formulation, and a simulated inlet condition based on a system calculation. Finite-rate chemistry was used throughout the study so that combustion effect is always included, and the effect of wall cooling on side load physics is studied. The side load physics captured include the afterburning wave, transition from free- shock to restricted-shock separation, and lip Lambda shock oscillation. With the adiabatic nozzle, free-shock separation reappears after the transition from free-shock separation to restricted-shock separation, and the subsequent flow pattern of the simultaneous free-shock and restricted-shock separations creates a very asymmetric Mach disk flow. With the cooled nozzle, the more symmetric restricted-shock separation persisted throughout the start-up transient after the transition, leading to an overall lower side load than that of the adiabatic nozzle. The tepee structures corresponding to the maximum side load were addressed.

  14. Apollo Contour Rocket Nozzle in the Propulsion Systems Laboratory

    NASA Image and Video Library

    1964-07-21

    Bill Harrison and Bud Meilander check the setup of an Apollo Contour rocket nozzle in the Propulsion Systems Laboratory at the National Aeronautics and Space Administration (NASA) Lewis Research Center. The Propulsion Systems Laboratory contained two 14-foot diameter test chambers that could simulate conditions found at very high altitudes. The facility was used in the 1960s to study complex rocket engines such as the Pratt and Whitney RL-10 and rocket components such as the Apollo Contour nozzle, seen here. Meilander oversaw the facility’s mechanics and the installation of test articles into the chambers. Harrison was head of the Supersonic Tunnels Branch in the Test Installations Division. Researchers sought to determine the impulse value of the storable propellant mix, classify and improve the internal engine performance, and compare the results with analytical tools. A special setup was installed in the chamber that included a device to measure the thrust load and a calibration stand. Both cylindrical and conical combustion chambers were examined with the conical large area ratio nozzles. In addition, two contour nozzles were tested, one based on the Apollo Service Propulsion System and the other on the Air Force’s Titan transtage engine. Three types of injectors were investigated, including a Lewis-designed model that produced 98-percent efficiency. It was determined that combustion instability did not affect the nozzle performance. Although much valuable information was obtained during the tests, attempts to improve the engine performance were not successful.

  15. Analytical study of nozzle performance for nuclear thermal rockets

    NASA Technical Reports Server (NTRS)

    Davidian, Kenneth O.; Kacynski, Kenneth J.

    1991-01-01

    Nuclear propulsion has been identified as one of the key technologies needed for human exploration of the Moon and Mars. The Nuclear Thermal Rocket (NTR) uses a nuclear reactor to heat hydrogen to a high temperature followed by expansion through a conventional convergent-divergent nozzle. A parametric study of NTR nozzles was performed using the Rocket Engine Design Expert System (REDES) at the NASA Lewis Research Center. The REDES used the JANNAF standard rigorous methodology to determine nozzle performance over a range of chamber temperatures, chamber pressures, thrust levels, and different nozzle configurations. A design condition was set by fixing the propulsion system exit radius at five meters and throat radius was varied to achieve a target thrust level. An adiabatic wall was assumed for the nozzle, and its length was assumed to be 80 percent of a 15 degree cone. The results conclude that although the performance of the NTR, based on infinite reaction rates, looks promising at low chamber pressures, finite rate chemical reactions will cause the actual performance to be considerably lower. Parameters which have a major influence on the delivered specific impulse value include the chamber temperature and the chamber pressures in the high thrust domain. Other parameters, such as 2-D and boundary layer effects, kinetic rates, and number of nozzles, affect the deliverable performance of an NTR nozzle to a lesser degree. For a single nozzle, maximum performance of 930 seconds and 1030 seconds occur at chamber temperatures of 2700 and 3100 K, respectively.

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

    NASA Astrophysics Data System (ADS)

    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.

  17. Advanced Solid Rocket Motor nozzle development status

    NASA Astrophysics Data System (ADS)

    Kearney, W. J.; Moss, J. D.

    1993-06-01

    This paper presents a status update of the design and development of an improved nozzle for the Advanced Solid Rocket Motor (ASRM). The ASRM nozzle incorporates advanced state-of-the-art design features and materials which contribute to enhanced safety, reliability, performance, and producibility for the space shuttle boosters. During 1992 the nozzle design progressed through a successful Preliminary Design Review (PDR). An improved ablative material development program also culminated in the selection of new standard and low density carbon cloth phenolic prepreg offering reduced variability and improved process attributes. A subscale motor test series to evaluate new materials and design features was also completed. An overview update of the matured design characteristics, supporting analysis, key development-program results and program status and plans is reported.

  18. Advanced Solid Rocket Motor nozzle development status

    NASA Technical Reports Server (NTRS)

    Kearney, W. J.; Moss, J. D.

    1993-01-01

    This paper presents a status update of the design and development of an improved nozzle for the Advanced Solid Rocket Motor (ASRM). The ASRM nozzle incorporates advanced state-of-the-art design features and materials which contribute to enhanced safety, reliability, performance, and producibility for the space shuttle boosters. During 1992 the nozzle design progressed through a successful Preliminary Design Review (PDR). An improved ablative material development program also culminated in the selection of new standard and low density carbon cloth phenolic prepreg offering reduced variability and improved process attributes. A subscale motor test series to evaluate new materials and design features was also completed. An overview update of the matured design characteristics, supporting analysis, key development-program results and program status and plans is reported.

  19. Advanced Solid Rocket Motor nozzle development status

    NASA Technical Reports Server (NTRS)

    Kearney, W. J.; Moss, J. D.

    1993-01-01

    This paper presents a status update of the design and development of an improved nozzle for the Advanced Solid Rocket Motor (ASRM). The ASRM nozzle incorporates advanced state-of-the-art design features and materials which contribute to enhanced safety, reliability, performance, and producibility for the space shuttle boosters. During 1992 the nozzle design progressed through a successful Preliminary Design Review (PDR). An improved ablative material development program also culminated in the selection of new standard and low density carbon cloth phenolic prepreg offering reduced variability and improved process attributes. A subscale motor test series to evaluate new materials and design features was also completed. An overview update of the matured design characteristics, supporting analysis, key development-program results and program status and plans is reported.

  20. A performance comparison of two small rocket nozzles

    NASA Technical Reports Server (NTRS)

    Arrington, Lynn A.; Reed, Brian D.; Rivera, Angel, Jr.

    1996-01-01

    An experimental study was conducted on two small rockets (110 N thrust class) to directly compare a standard conical nozzle with a bell nozzle optimized for maximum thrust using the Rao method. In large rockets, with throat Reynolds numbers of greater than 1 x 10(exp 5), bell nozzles outperform conical nozzles. In rockets with throat Reynolds numbers below 1 x 10(exp 5), however, test results have been ambiguous. An experimental program was conducted to test two small nozzles at two different fuel film cooling percentages and three different chamber pressures. Test results showed that for the throat Reynolds number range from 2 x 10(exp 4) to 4 x 10(exp 4), the bell nozzle outperformed the conical nozzle. Thrust coefficients for the bell nozzle were approximately 4 to 12 percent higher than those obtained with the conical nozzle. As expected, testing showed that lowering the fuel film cooling increased performance for both nozzle types.

  1. Nuclear thermal rocket nozzle testing and evaluation program

    NASA Technical Reports Server (NTRS)

    Davidian, Kenneth O.; Kacynski, Kenneth J.

    1993-01-01

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

  2. A hydrogen-oxygen rocket engine coolant passage design program (RECOP) for fluid-cooled thrust chambers and nozzles

    NASA Technical Reports Server (NTRS)

    Tomsik, Thomas M.

    1994-01-01

    The design of coolant passages in regeneratively cooled thrust chambers is critical to the operation and safety of a rocket engine system. Designing a coolant passage is a complex thermal and hydraulic problem requiring an accurate understanding of the heat transfer between the combustion gas and the coolant. Every major rocket engine company has invested in the development of thrust chamber computer design and analysis tools; two examples are Rocketdyne's REGEN code and Aerojet's ELES program. In an effort to augment current design capabilities for government and industry, the NASA Lewis Research Center is developing a computer model to design coolant passages for advanced regeneratively cooled thrust chambers. The RECOP code incorporates state-of-the-art correlations, numerical techniques and design methods, certainly minimum requirements for generating optimum designs of future space chemical engines. A preliminary version of the RECOP model was recently completed and code validation work is in progress. This paper introduces major features of RECOP and compares the analysis to design points for the first test case engine; the Pratt & Whitney RL10A-3-3A thrust chamber.

  3. Status of flow separation prediction in liquid propellant rocket nozzles

    NASA Technical Reports Server (NTRS)

    Schmucker, R. H.

    1974-01-01

    Flow separation which plays an important role in the design of a rocket engine nozzle is discussed. For a given ambient pressure, the condition of no flow separation limits the area ratio and, therefore, the vacuum performance. Avoidance of performance loss due to area ratio limitation requires a correct prediction of the flow separation conditions. To provide a better understanding of the flow separation process, the principal behavior of flow separation in a supersonic overexpanded rocket nozzle is described. The hot firing separation tests from various sources are summarized, and the applicability and accuracy of the measurements are described. A comparison of the different data points allows an evaluation of the parameters that affect flow separation. The pertinent flow separation predicting methods, which are divided into theoretical and empirical correlations, are summarized and the numerical results are compared with the experimental points.

  4. Reusable Solid Rocket Motor Nozzle Joint 5 Redesign

    NASA Technical Reports Server (NTRS)

    Lui, R. C.; Stratton, T. C.; LaMont, D. T.

    2003-01-01

    Torque tension testing of a newly designed Reusable Solid Rocket Motor nozzle bolted assembly was successfully completed. Test results showed that the 3-sigma preload variation was as expected at the required input torque level and the preload relaxation were within the engineering limits. A shim installation technique was demonstrated as a simple process to fill a shear lip gap between nozzle housings in the joint region. A new automated torque system was successfully demonstrated in this test. This torque control tool was found to be very precise and accurate. The bolted assembly performance was further evaluated using the Nozzle Structural Test Bed. Both current socket head cap screw and proposed multiphase alloy bolt configurations were tested. Results indicated that joint skip and bolt bending were significantly reduced with the new multiphase alloy bolt design. This paper summarizes all the test results completed to date.

  5. Reusable Solid Rocket Motor Nozzle Joint 5 Redesign

    NASA Technical Reports Server (NTRS)

    Lui, R. C.; Stratton, T. C.; LaMont, D. T.

    2003-01-01

    Torque tension testing of a newly designed Reusable Solid Rocket Motor nozzle bolted assembly was successfully completed. Test results showed that the 3-sigma preload variation was as expected at the required input torque level and the preload relaxation were within the engineering limits. A shim installation technique was demonstrated as a simple process to fill a shear lip gap between nozzle housings in the joint region. A new automated torque system was successfully demonstrated in this test. This torque control tool was found to be very precise and accurate. The bolted assembly performance was further evaluated using the Nozzle Structural Test Bed. Both current socket head cap screw and proposed multiphase alloy bolt configurations were tested. Results indicated that joint skip and bolt bending were significantly reduced with the new multiphase alloy bolt design. This paper summarizes all the test results completed to date.

  6. Altitude compensating ablative stiffening band for rocket motor nozzles

    NASA Astrophysics Data System (ADS)

    Brown, J. Lynn; McIntire, Vaughn W., Jr.; Clontz, Leslie A.; Peckham, Richard J.; Dixon, Alan B. C.; West, James C.

    1993-03-01

    A rocket motor nozzle with an ablative internal structural member which provides rigidity to the rocket motor both prior to and during motor operation is described. The rocket motor nozzle of the present invention includes an outer shell which converges to a throat and then diverges to form an entrance. An ablative stiffening band, secured to the inner surface of the nozzle and extending from the aft portion of the nozzle to the throat, is contoured to allow the exhaust gases to flow smoothly through a central longitudinal opening in the band.

  7. Parametric study of solar thermal rocket nozzle performance

    NASA Technical Reports Server (NTRS)

    Pearson, J. Boise; Landrum, D. Brian; Hawk, Clark W.

    1995-01-01

    This paper details a numerical investigation of performance losses in low-thrust solar thermal rocket nozzles. The effects of nozzle geometry on three types of losses were studied; finite rate dissociation-recombination kinetic losses, two dimensional axisymmetric divergence losses, and compressible viscous boundary layer losses. Short nozzle lengths and supersonic flow produce short residence times in the nozzle and a nearly frozen flow, resulting in large kinetic losses. Variations in geometry have a minimal effect on kinetic losses. Divergence losses are relatively small, and careful shaping of the nozzle can nearly eliminate them. The boundary layer in these small nozzles can grow to a major fraction of nozzle radius, and cause large losses. These losses are attributed to viscous drag on the nozzle walls and flow blockage by the boundary layer, especially in the throat region. Careful shaping of the nozzle can produce a significant reduction in viscous losses.

  8. General view of a Solid Rocket Motor Nozzle in the ...

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

    General view of a Solid Rocket Motor Nozzle in the Solid Rocket Booster (SRB) Assembly and Refurbishment Facility at Kennedy Space Center, being prepared to be mated with the Aft Skirt. In this view you can see the attach brackets where the Thrust Vector Control System actuators connect to the nozzle which can swivel the nozzle up to 3.5 degrees to redirect the thrust to steer and maintain the Shuttle's programmed trajectory. - Space Transportation System, Solid Rocket Boosters, Lyndon B. Johnson Space Center, 2101 NASA Parkway, Houston, Harris County, TX

  9. Russian Rocket Engine Test

    NASA Technical Reports Server (NTRS)

    1998-01-01

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

  10. Rocket nozzle coolant channel thermal analysis program (E25107)

    NASA Technical Reports Server (NTRS)

    Thompson, W. R.

    1972-01-01

    A complete description of the liquid cooled rocket nozzle analysis program (E25107) is presented, including a users manual, program listing, and a sample problem. The program is recommended for use in designing liquid cooled rocket nozzles. In addition, it is adaptable to any system in which a liquid-cooled tubular structure is used to contain and direct the flow of a hot gas.

  11. Star 48 solid rocket motor nozzle analyses and instrumented firings

    NASA Technical Reports Server (NTRS)

    Porter, R. L.

    1986-01-01

    The analyses and testing performed by NASA in support of an expanded and improved nozzle design data base for use by the U.S. solid rocket motor industry is presented. A production nozzle with a history of one ground failure and two flight failures was selected for analyses and testing. The stress analysis was performed with the Champion computer code developed by the U.S. Navy. Several improvements were made to the code. Strain predictions were made and compared to test data. Two short duration motor firings were conducted with highly instrumented nozzles. The first nozzle had 58 thermocouples, 66 strain gages, and 8 bondline pressure measurements. The second nozzle had 59 thermocouples, 68 strain measurements, and 8 bondline pressure measurements. Most of this instrumentation was on the nonmetallic parts, and provided significantly more thermal and strain data on the nonmetallic components of a nozzle than has been accumulated in a solid rocket motor test to date.

  12. Rocket Engine Oscillation Diagnostics

    NASA Technical Reports Server (NTRS)

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

    2002-01-01

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

  13. Robust Rocket Engine Concept

    NASA Technical Reports Server (NTRS)

    Lorenzo, Carl F.

    1995-01-01

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

  14. Rocket engine numerical simulator

    NASA Technical Reports Server (NTRS)

    Davidian, Ken

    1993-01-01

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

  15. Rocket engine numerical simulation

    NASA Technical Reports Server (NTRS)

    Davidian, Ken

    1993-01-01

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

  16. Rocket nozzle thermal shock tests in an arc heater facility

    NASA Technical Reports Server (NTRS)

    Painter, James H.; Williamson, Ronald A.

    1986-01-01

    A rocket motor nozzle thermal structural test technique that utilizes arc heated nitrogen to simulate a motor burn was developed. The technique was used to test four heavily instrumented full-scale Star 48 rocket motor 2D carbon/carbon segments at conditions simulating the predicted thermal-structural environment. All four nozzles survived the tests without catastrophic or other structural failures. The test technique demonstrated promise as a low cost, controllable alternative to rocket motor firing. The technique includes the capability of rapid termination in the event of failure, allowing post-test analysis.

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

    NASA Astrophysics Data System (ADS)

    Honoré, M.; Brox, L.; Hallberg, M.

    2012-03-01

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

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

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

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

  19. Russian Rocket Engine Test

    NASA Technical Reports Server (NTRS)

    1998-01-01

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

  20. Rainbows and Rocket Engine

    NASA Image and Video Library

    2017-02-22

    Rainbows and rocket engines – doesn’t get much better than that! Check out these gorgeous aerial views from today’s Space Launch System RS-25 engine test @NASA’s Stennis Space Center. PAO Name:Kim Henry Phone Number:256-544-1899 Email Address: kimberly.m.henry@nasa.gov

  1. Ground-based firing of a linear aerospike rocket engine

    NASA Technical Reports Server (NTRS)

    1990-01-01

    In this 23-second clip the Rocketdyne linear aerospike rocket engine is fired in a test stand at the Santa Susana facility of Rocketdyne in Chatsworth, California. Unlike conventional rocket engines the linear aerospike has no bell nozzle to contain and direct the expanding gases. Instead, the gases are aimed down a V-shaped ramp, which makes up one side of the 'nozzle,' with the side open to the atmosphere allowing optimum expansion at all altitudes while providing thrust.

  2. Blast Effects of Twin Variable-Cant Rocket Nozzles on Visibility During Landing

    NASA Technical Reports Server (NTRS)

    1964-01-01

    Blast Effects of Twin Variable-Cant Rocket Nozzles on Visibility During Landing on a Particle-Covered Surface. A limited investigation has been conducted to determine the jet-blast effect of twin variable-cant supersonic nozzles. These tests were made to examine the result of using canted main rocket engines to sweep the blast debris outward from the proposed landing area of a rocket-powered vehicle making a vertical approach to a touchdown. Cant angles from 0 degrees to 75 degrees, at intervals of 15 degrees, were tested at low ambient pressure and at atmospheric ambient pressure. Nozzle chamber pressures used were 400 psi and 2000 psi. [Entire movie available on DVD from CASI as Doc ID 20070030964. Contact help@sti.nasa.gov

  3. Modified RS2101 rocket engine study program

    NASA Technical Reports Server (NTRS)

    1971-01-01

    The purpose of the program is to perform design studies and analyses to determine the effects of incorporating a 60:1 expansion area ratio nozzle extension, extended firing time, and modified operating conditions and environments on the MM'71 rocket engine assembly. An injector-to-thrust chamber seal study was conducted to define potential solutions for leakage past this joint. The results and recommendations evolving from the engine thermal analyses, the injector-to-thrust chamber seal studies, and the nozzle extension joint stress analyses are presented.

  4. Numerical simulation of film-cooled ablative rocket nozzles

    NASA Technical Reports Server (NTRS)

    Landrum, D. B.; Beard, R. M.

    1996-01-01

    The objective of this research effort was to evaluate the impact of incorporating an additional cooling port downstream between the injector and nozzle throat in the NASA Fast Track chamber. A numerical model of the chamber was developed for the analysis. The analysis did not model ablation but instead correlated the initial ablation rate with the initial nozzle wall temperature distribution. The results of this study provide guidance in the development of a potentially lighter, second generation ablative rocket nozzle which maintains desired performance levels.

  5. Convoluted nozzle design for the RL10 derivative 2B engine

    NASA Technical Reports Server (NTRS)

    1985-01-01

    The convoluted nozzle is a conventional refractory metal nozzle extension that is formed with a portion of the nozzle convoluted to show the extendible nozzle within the length of the rocket engine. The convoluted nozzle (CN) was deployed by a system of four gas driven actuators. For spacecraft applications the optimum CN may be self-deployed by internal pressure retained, during deployment, by a jettisonable exit closure. The convoluted nozzle is included in a study of extendible nozzles for the RL10 Engine Derivative 2B for use in an early orbit transfer vehicle (OTV). Four extendible nozzle configurations for the RL10-2B engine were evaluated. Three configurations of the two position nozzle were studied including a hydrogen dump cooled metal nozzle and radiation cooled nozzles of refractory metal and carbon/carbon composite construction respectively.

  6. Overview of rocket engine control

    NASA Technical Reports Server (NTRS)

    Lorenzo, Carl F.; Musgrave, Jeffrey L.

    1991-01-01

    The issues of Chemical Rocket Engine Control are broadly covered. The basic feedback information and control variables used in expendable and reusable rocket engines, such as Space Shuttle Main Engine, are discussed. The deficiencies of current approaches are considered and a brief introduction to Intelligent Control Systems for rocket engines (and vehicles) is presented.

  7. Liquid Rocket Engine Testing

    NASA Technical Reports Server (NTRS)

    Rahman, Shamim

    2005-01-01

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

  8. Liquid rocket engine turbines

    NASA Technical Reports Server (NTRS)

    1974-01-01

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

  9. Engine for Redstone Rocket

    NASA Technical Reports Server (NTRS)

    2004-01-01

    This photograph is of the engine for the Redstone rocket. The Redstone ballistic missile was a high-accuracy, liquid-propelled, surface-to-surface missile developed by the Army Ballistic Missile Agency, Redstone Arsenal, in Huntsville, Alabama, under the direction of Dr. von Braun. The Redstone engine was a modified and improved version of the Air Force's Navaho cruise missile engine of the late forties. The A-series, as this would be known, utilized a cylindrical combustion chamber as compared with the bulky, spherical V-2 chamber. By 1951, the Army was moving rapidly toward the design of the Redstone missile, and the production was begun in 1952. Redstone rockets became the 'reliable workhorse' for America's early space program. As an example of its versatility, the Redstone was utilized in the booster for Explorer 1, the first American satellite, with no major changes to the engine or missile.

  10. Rocket nozzle lip flow by direct simulation Monte Carlo method

    NASA Technical Reports Server (NTRS)

    Hueser, J. E.; Brock, F. J.; Melfi, L. T., Jr.; Bird, G. A.

    1985-01-01

    The flow in the immediate vicinity of a rocket nozzle lip has been analyzed, and the results are presented. A rapid change in gas composition is observed in the flow around the lip depending principally on species molecular mass and local flow angle. The divergence of axial and radial temperatures indicating breakdown of translational equilibrium is observed in the results.

  11. Stress-testing of the throat of a rocket's nozzle

    NASA Technical Reports Server (NTRS)

    Estes, E. G.

    1969-01-01

    Test motor in which high initial pressure can be reduced suddenly provides a method of testing stress effects in the throat of a rockets nozzle. Motors operating pressure is increased to aggravate tensile stresses in a submerged throat. Opposing compression stresses are limited by control of the operating pressure.

  12. Thermographic Leak Detection of the Space Shuttle Main Engine Nozzle

    NASA Technical Reports Server (NTRS)

    Walker, James L.; Russell, Samuel S.

    1999-01-01

    The Space Shuttle Main Engines Nozzles consist of over one thousand tapered Inconel coolant tubes brazed to a stainless steel structural jacket. Liquid Hydrogen flows through the tubing, from the aft to forward end of the nozzle, under high pressure to maintain a thermal balance between the rocket exhaust and the nozzle wall. Three potential problems occur within the SSME nozzle coolant tubes as a result of manufacturing anomalies and the highly volatile service environment including poor or incomplete bonding of the tubes to the structural jacket, cold wall leaks and hot wall leaks. Of these conditions the identification of cold wall leaks has been the most problematic. The methods and results presented in this summary addresses the thermographic identification of cold wall "interstitial" leaks between the structural jacket and coolant tubes of the Space Shuttle Main Engines Nozzles.

  13. Jupiter Rocket Engine

    NASA Technical Reports Server (NTRS)

    2004-01-01

    Engine for the Jupiter rocket. The Jupiter vehicle was a direct derivative of the Redstone. The Army Ballistic Missile Agency (ABMA) at Redstone Arsenal, Alabama, continued Jupiter development into a successful intermediate ballistic missile, even though the Department of Defense directed its operational development to the Air Force. ABMA maintained a role in Jupiter RD, including high-altitude launches that added to ABMA's understanding of rocket vehicle operations in the near-Earth space environment. It was knowledge that paid handsome dividends later.

  14. Numerical study of nozzle wall cooling for nuclear thermal rockets

    NASA Technical Reports Server (NTRS)

    Kim, Suk C.; Stubbs, Robert M.

    1993-01-01

    The flowfields and performance of nuclear thermal rockets, which utilize radiation and film-cooling to cool the nozzle extension, are studied by solving the Navier-Stokes equations and species equations. The thrust level of the rocket for the present study is about 75,000 lb(f) for a chamber pressure of 68 atm(l,000 psi) and a chamber temperature of 2700 K. The throat radius of the nozzle is 0.0936 m and the area ratios of the nozzles are 300 and 500. It is assumed that the flow is chemically frozen and the turbulence is simulated by the modified Baldwin-Lomax turbulence model. The calculated results for various area ratios and film mass-flow rates are presented as Mach number contours, variations of nozzle wall temperature, exit profiles, and vacuum specific impulses. The present study shows that by selecting the flow rate of the film-cooling hydrogen and area ratio of the nozzle correctly, high area ratio nozzle extensions can be cooled effectively with radiation and film-cooling without significant penalty in performance.

  15. Fabrication of Composite Combustion Chamber/Nozzle for Fastrac Engine

    NASA Technical Reports Server (NTRS)

    Lawrence, T.; Beshears, R.; Burlingame, S.; Peters, W.; Prince, M.; Suits, M.; Tillery, S.; Burns, L.; Kovach, M.; Roberts, K.

    2001-01-01

    The Fastrac Engine developed by the Marshall Space Flight Center for the X-34 vehicle began as a low cost engine development program for a small booster system. One of the key components to reducing the engine cost was the development of an inexpensive combustion chamber/nozzle. Fabrication of a regeneratively cooled thrust chamber and nozzle was considered too expensive and time consuming. In looking for an alternate design concept, the Space Shuttle's Reusable Solid Rocket Motor Project provided an extensive background with ablative composite materials in a combustion environment. An integral combustion chamber/nozzle was designed and fabricated with a silica/phenolic ablative liner and a carbon/epoxy structural overwrap. This paper describes the fabrication process and developmental hurdles overcome for the Fastrac engine one-piece composite combustion chamber/nozzle.

  16. Fabrication of Composite Combustion Chamber/Nozzle for Fastrac Engine

    NASA Technical Reports Server (NTRS)

    Lawerence, T.; Beshears, R.; Burlingame, S.; Peters, W.; Prince, M.; Suits, M.; Tillery, S.; Burns, L.; Kovach, M.; Roberts, K.; hide

    2000-01-01

    The Fastrac Engine developed by the Marshall Space Flight Center for the X-34 vehicle began as a low cost engine development program for a small booster system. One of the key components to reducing the engine cost was the development of an inexpensive combustion chamber/nozzle. Fabrication of a regeneratively cooled thrust chamber and nozzle was considered too expensive and time consuming. In looking for an alternate design concept, the Space Shuttle's Reusable Solid Rocket Motor Project provided an extensive background with ablative composite materials in a combustion environment. An integral combustion chamber/nozzle was designed and fabricated with a silica/phenolic ablative liner and a carbon/epoxy structural overwrap. This paper describes the fabrication process and developmental hurdles overcome for the Fastrac engine one-piece composite combustion chamber/nozzle.

  17. Analytical investigation of solid rocket nozzle failure

    NASA Technical Reports Server (NTRS)

    Mccoy, K. E.; Hester, J.

    1985-01-01

    On April 5, 1983, an Inertial Upper Stage (IUS) spacecraft experienced loss of control during the burn of the second of two solid rocket motors. The anomaly investigation showed the cause to be a malfunction of the solid rocket motor. This paper presents a description of the IUS system, a failure analysis summary, an account of the thermal testing and computer modeling done at Marshall Space Flight Center, a comparison of analysis results with thermal data obtained from motor static tests, and describes some of the design enhancement incorporated to prevent recurrence of the anomaly.

  18. Transpiration And Regenerative Cooling Of Rocket Engine

    NASA Technical Reports Server (NTRS)

    Obrien, Charles J.

    1989-01-01

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

  19. Experimental Studies of the Heat Transfer to RBCC Rocket Nozzles for CFD Application to Design Methodologies

    NASA Technical Reports Server (NTRS)

    Santoro, Robert J.; Pal, Sibtosh

    1999-01-01

    Rocket thrusters for Rocket Based Combined Cycle (RBCC) engines typically operate with hydrogen/oxygen propellants in a very compact space. Packaging considerations lead to designs with either axisymmetric or two-dimensional throat sections. Nozzles tend to be either two- or three-dimensional. Heat transfer characteristics, particularly in the throat, where the peak heat flux occurs, are not well understood. Heat transfer predictions for these small thrusters have been made with one-dimensional analysis such as the Bartz equation or scaling of test data from much larger thrusters. The current work addresses this issue with an experimental program that examines the heat transfer characteristics of a gaseous oxygen (GO2)/gaseous hydrogen (GH2) two-dimensional compact rocket thruster. The experiments involved measuring the axial wall temperature profile in the nozzle region of a water-cooled gaseous oxygen/gaseous hydrogen rocket thruster at a pressure of 3.45 MPa. The wall temperature measurements in the thruster nozzle in concert with Bartz's correlation are utilized in a one-dimensional model to obtain axial profiles of nozzle wall heat flux.

  20. Development of moldable carbonaceous materials for ablative rocket nozzles.

    NASA Technical Reports Server (NTRS)

    Lockhart, R. J.; Bortz, S. A.; Schwartz, M. A.

    1972-01-01

    Description of a materials system developed for use as low-cost ablative nozzles for NASA's 260-in. solid rocket motor. Petroleum coke and carbon black fillers were employed; high density was achieved by controlling particle size distribution. An alumina catalyzed furfuryl ester resin which produced high carbon residues after pyrolysis was employed as the binder. Staple carbon fibers improved the strength and crack resistance of molded bodies. In static firing tests of two subscale nozzles, this material compared favorably in erosion rate with several other ablative systems.

  1. Development of moldable carbonaceous materials for ablative rocket nozzles.

    NASA Technical Reports Server (NTRS)

    Lockhart, R. J.; Bortz, S. A.; Schwartz, M. A.

    1972-01-01

    Description of a materials system developed for use as low-cost ablative nozzles for NASA's 260-in. solid rocket motor. Petroleum coke and carbon black fillers were employed; high density was achieved by controlling particle size distribution. An alumina catalyzed furfuryl ester resin which produced high carbon residues after pyrolysis was employed as the binder. Staple carbon fibers improved the strength and crack resistance of molded bodies. In static firing tests of two subscale nozzles, this material compared favorably in erosion rate with several other ablative systems.

  2. Analysis of film cooling in rocket nozzles

    NASA Technical Reports Server (NTRS)

    Woodbury, Keith A.

    1993-01-01

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

  3. Liquid Rocket Engine Testing Overview

    NASA Technical Reports Server (NTRS)

    Rahman, Shamim

    2005-01-01

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

  4. Space Shuttle Solid Rocket Nozzle Joint 2 Gas Path Study

    NASA Technical Reports Server (NTRS)

    Lui, R. C.; Hyer, R. L.

    1999-01-01

    Each Space Shuttle solid rocket nozzle contains five joints. By design, each nozzle joint utilizes room-temperature vulcanized (RTV) rubber as a thermal barrier to protect the nozzle metal parts and O-rings from heat degradation due to motor combustion gas. Typically, combustion gas does not penetrate through the RTV in nozzle Joints. However, Joint currently experiences gas penetration through the RTV over 90 percent of the time. The occurrence of Joint gas paths is a very complex phenomenon with several known contributors such as the unique geometry of the joint, motor pressure, nozzle vectoring, manufacturing induced RTV voids. RTV material capability, and joint skip (relative radial displacement of joint components). The occurrence of Joint gas paths appears to be a combination of these causes. To better understand the causes of gas paths in Joint, the circumferential locations of gas paths in the RTV were identified and correlated with nozzle vectoring, RTV closeour locations (where the RTV backfill is completed), and X-ray low-density. anomaly (LDA) indications. Results that no specific cause was determined.

  5. Space Shuttle Solid Rocket Nozzle Joint 2 Gas Path Study

    NASA Technical Reports Server (NTRS)

    Lui, R. C.; Hyer, R. L.

    1999-01-01

    Each Space Shuttle solid rocket nozzle contains five joints. By design, each nozzle joint utilizes room-temperature vulcanized (RTV) rubber as a thermal barrier to protect the nozzle metal parts and O-rings from heat degradation due to motor combustion gas. Typically, combustion gas does not penetrate through the RTV in nozzle Joints. However, Joint currently experiences gas penetration through the RTV over 90 percent of the time. The occurrence of Joint gas paths is a very complex phenomenon with several known contributors such as the unique geometry of the joint, motor pressure, nozzle vectoring, manufacturing induced RTV voids. RTV material capability, and joint skip (relative radial displacement of joint components). The occurrence of Joint gas paths appears to be a combination of these causes. To better understand the causes of gas paths in Joint, the circumferential locations of gas paths in the RTV were identified and correlated with nozzle vectoring, RTV closeour locations (where the RTV backfill is completed), and X-ray low-density. anomaly (LDA) indications. Results that no specific cause was determined.

  6. AXISYMMETRIC, THROTTLEABLE NON-GIMBALLED ROCKET ENGINE

    NASA Technical Reports Server (NTRS)

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

    2005-01-01

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

  7. Transpiration cooled throat for hydrocarbon rocket engines

    NASA Technical Reports Server (NTRS)

    May, Lee R.; Burkhardt, Wendel M.

    1991-01-01

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

  8. Development of Thermal Barriers For Solid Rocket Motor Nozzle Joints

    NASA Technical Reports Server (NTRS)

    Steinetz, Bruce M.; Dunlap, Patrick H., Jr.

    2000-01-01

    Joints in the Space Shuttle solid rocket motors are sealed by O-rings to contain combustion gases inside the rocket that reach pressures of up to 900 psi and temperatures of up to 5500 F. To provide protection for the O-rings, the motors are insulated with either phenolic or rubber insulation. Gaps in the joints leading up to the O-rings are filled with polysulfide joint-fill compounds as an additional level of protection. The current RSRM nozzle-to-case joint design incorporating primary, secondary, and wiper O-rings experiences gas paths through the joint-fill compound to the innermost wiper O-ring in about one out of every seven motors. Although this does not pose a safety hazard to the motor, it is an undesirable condition that NASA and rocket manufacturer Thiokol want to eliminate. Each nozzle-to-case joint gas path results in extensive reviews and evaluation before flights can be resumed. Thiokol and NASA Marshall are currently working to improve the nozzle-to-case joint design by implementing a more reliable J-leg design that has been used successfully in the field and igniter joint. They are also planning to incorporate the NASA Glenn braided carbon fiber thermal barrier into the joint. The thermal barrier would act as an additional level of protection for the O-rings and allow the elimination of the joint-fill compound from the joint.

  9. Thermal Barriers Developed for Solid Rocket Motor Nozzle Joints

    NASA Technical Reports Server (NTRS)

    Steinetz, Bruce M.; Dunlap, Patrick H., Jr.

    2000-01-01

    Space shuttle solid rocket motor case assembly joints are sealed with conventional O-ring seals that are shielded from 5500 F combustion gases by thick layers of insulation and by special joint-fill compounds that fill assembly splitlines in the insulation. On a number of occasions, NASA has observed hot gas penetration through defects in the joint-fill compound of several of the rocket nozzle assembly joints. In the current nozzle-to-case joint, NASA has observed penetration of hot combustion gases through the joint-fill compound to the inboard wiper O-ring in one out of seven motors. Although this condition does not threaten motor safety, evidence of hot gas penetration to the wiper O-ring results in extensive reviews before resuming flight. The solid rocket motor manufacturer (Thiokol) approached the NASA Glenn Research Center at Lewis Field about the possibility of applying Glenn's braided fiber preform seal as a thermal barrier to protect the O-ring seals. Glenn and Thiokol are working to improve the nozzle-to-case joint design by implementing a more reliable J-leg design and by using a braided carbon fiber thermal barrier that would resist any hot gases that the J-leg does not block.

  10. ION ROCKET ENGINE

    DOEpatents

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

    1961-12-19

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

  11. Vacuum plasma spray applications on liquid fuel rocket engines

    NASA Technical Reports Server (NTRS)

    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.

  12. Vacuum plasma spray applications on liquid fuel rocket engines

    NASA Astrophysics Data System (ADS)

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

    1992-07-01

    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.

  13. JANNAF boundary layer integral matrix procedure. [turbulent boundary layer/rocket nozzles - prediction analysis techniques

    NASA Technical Reports Server (NTRS)

    Evans, R. M.

    1975-01-01

    Accurate predictions of the thrust loss due to boundary layer effects and of the wall heat flux are very important to the design and performance evaluation of rocket nozzles. A method used in analytical procedures for liquid fuel rocket engine performance prediction and evaluation is presented. A computer program is examined that is a fast and accurate procedure for solving the set of boundary layer equation (momentum, energy, and species) for laminar or turbulent, chemically reacting flows with a wide variety of boundary conditions. Results of comparison of the various turbulent models are presented. A summary of the modifications and additions to the program is examined.

  14. Manufacturing Process Developments for Regeneratively-Cooled Channel Wall Rocket Nozzles

    NASA Technical Reports Server (NTRS)

    Gradl, Paul; Brandsmeier, Will

    2016-01-01

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

  15. Digital program analyzes supersonic flow field within bell-shaped rocket nozzles

    NASA Technical Reports Server (NTRS)

    Elliott, J. J.; Stromsta, R. R.

    1970-01-01

    Digital computer program computes and analyzes supersonic flow field in axisymmetric rocket nozzle for specified gas properties, nozzle geometry, and input or starting line. Method of characteristics is used for solution of set of hyperbolic partial differential equations.

  16. Development of Thermal Barriers for Solid Rocket Motor Nozzle Joints

    NASA Technical Reports Server (NTRS)

    Steinetz, Bruce M.; Dunlap, Patrick H., Jr.

    1999-01-01

    The Space Shuttle solid rocket motor case assembly joints are sealed using conventional 0-ring seals. The 5500+F combustion gases are kept a safe distance away from the seals by thick layers of insulation. Special joint-fill compounds are used to fill the joints in the insulation to prevent a direct flowpath to the seals. On a number of occasions. NASA has observed in several of the rocket nozzle assembly joints hot gas penetration through defects in the joint- fill compound. The current nozzle-to-case joint design incorporates primary, secondary and wiper (inner-most) 0-rings and polysulfide joint-fill compound. In the current design, 1 out of 7 motors experience hot gas to the wiper 0-ring. Though the condition does not threaten motor safety, evidence of hot gas to the wiper 0-ring results in extensive reviews before resuming flight. NASA and solid rocket motor manufacturer Thiokol are working to improve the nozzle-to-case joint design by implementing a more reliable J-leg design and a thermal barrier, This paper presents burn-resistance, temperature drop, flow and resiliency test results for several types of NASA braided carbon-fiber thermal barriers. Burn tests were performed to determine the time to burn through each of the thermal barriers when exposed to the flame of an oxy-acetylene torch (5500 F), representative of the 5500 F solid rocket motor combustion temperatures. Thermal barriers braided out of carbon fibers endured the flame for over 6 minutes, three times longer than solid rocket motor burn time. Tests were performed on two thermal barrier braid architectures, denoted Carbon-3 and Carbon-6, to measure the temperature drop across and along the barrier in a compressed state when subjected to the flame of an oxyacetylene torch. Carbon-3 and Carbon-6 thermal barriers were excellent insulators causing temperature drops through their diameter of up to a 2800 and 2560 F. respectively. Gas temperature 1/4" downstream of the thermal barrier were within the

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

    NASA Technical Reports Server (NTRS)

    Davidian, Kenneth O.

    1990-01-01

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

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

    NASA Technical Reports Server (NTRS)

    Davidian, Kenneth O.

    1990-01-01

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

  19. Acoustic structures in the near-field from clustered rocket nozzles

    NASA Astrophysics Data System (ADS)

    Canchero, Andres; Tinney, Charles E.; Murray, Nathan E.; Ruf, Joseph H.

    2014-11-01

    The plume and acoustic field produced by a cluster of two and four rocket nozzles is visualized by way of retroreflective shadowgraphy. Steady state and transient operations (startup/shutdown) were conducted in the fully-anechoic chamber and open jet facility of The University of Texas at Austin. The laboratory scale rocket nozzles comprise thrust-optimized parabolic contours, which during start-up, experience free shock separated flow, restricted shock separated flow and an end-effects regime prior to flowing full. Shadowgraphy images with synchronized surveys of the acoustic loads produced in close vicinity to the rocket clusters and wall static pressure profiles are first compared with several RANS simulations during steady operations. A Proper Orthogonal Decomposition of various regions in the shadowgraphy images is then performed to elucidate the prominent features residing in the supersonic annular flow region, the acoustic near field and the interaction zone that resides between the nozzle plumes. POD modes are used to detect propagation paths of the acoustic waves and shock cell structures in the supersonic shear layer. Spectral peak frequencies on the propagation paths are associated with the shock cell length, which are responsible for generating broadband shock noise. Aerospace Engineering & Engineering Mechanics.

  20. Coal-Fired Rocket Engine

    NASA Technical Reports Server (NTRS)

    Anderson, Floyd A.

    1987-01-01

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

  1. Wall Pressure Unsteadiness and Side Loads in Overexpanded Rocket Nozzles

    NASA Technical Reports Server (NTRS)

    Baars, Woutijn J.; Tinney, Charles E.; Ruf, Joseph H.; Brown, Andrew M.; McDaniels, David M.

    2012-01-01

    Surveys of both the static and dynamic wall pressure signatures on the interior surface of a sub-scale, cold-flow and thrust optimized parabolic nozzle are conducted during fixed nozzle pressure ratios corresponding to FSS and RSS states. The motive is to develop a better understanding for the sources of off-axis loads during the transient start-up of overexpanded rocket nozzles. During FSS state, pressure spectra reveal frequency content resembling SWTBLI. Presumably, when the internal flow is in RSS state, separation bubbles are trapped by shocks and expansion waves; interactions between the separated flow regions and the waves produce asymmetric pressure distributions. An analysis of the azimuthal modes reveals how the breathing mode encompasses most of the resolved energy and that the side load inducing mode is coherent with the response moment measured by strain gauges mounted upstream of the nozzle on a flexible tube. Finally, the unsteady pressure is locally more energetic during RSS, albeit direct measurements of the response moments indicate higher side load activity when in FSS state. It is postulated that these discrepancies are attributed to cancellation effects between annular separation bubbles.

  2. Atomic hydrogen rocket engine

    NASA Technical Reports Server (NTRS)

    Etters, R. D.; Flurchick, K.

    1981-01-01

    A rocket using atomic hydrogen propellant is discussed. An essential feature of the proposed engine is that the atomic hydrogen fuel is used as it is produced, thus eliminating the necessity of storage. The atomic hydrogen flows into a combustion chamber and recombines, producing high velocity molecular hydrogen which flows out an exhaust port. Standard thermodynamics, kinetic theory and wall recombination cross-sections are used to predict a thrust of approximately 1.4 N for a RF hydrogen flow rate of 4 x 10 to the 22nd/sec. Specific impulses are nominally from 1000 to 2000 sec. It is predicted that thrusts on the order of one Newton and specific impulses of up to 2200 sec are attainable with nominal RF discharge fluxes on the order of 10 to the 22nd atoms/sec; further refinements will probably not alter these predictions by more than a factor of two.

  3. Shuttle solid rocket motor nozzle alternate ablative evaluation

    NASA Technical Reports Server (NTRS)

    Powers, L. B.; Bailey, R. L.; Morrison, B. H.

    1981-01-01

    A series of subscale tests are shown to suggest that a lower-cost ablative material than the rayon-based carbon ablative currently used in the Space Shuttle Solid Rocket Motor (SRM) may be used as a substitute. Six such ablatives with outstanding performance characteristics, using spun PAN and continuous pitch and PAN fibers instead of the present, continuous rayon, were identified in the course of tests with HTPB/AL/AP solid propellant grains with a burn time of 12 sec. The test nozzle features an initial throat diameter of 2.2 in. and a 6.1 expansion ratio. In addition to nozzle structural feature drawings, extensive test data tables and propellant formulation and properties tables are provided.

  4. Rocket Engine Thrust Chamber Assembly

    NASA Technical Reports Server (NTRS)

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

    2001-01-01

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

  5. Application of Optimization Techniques to Design of Unconventional Rocket Nozzle Configurations

    NASA Technical Reports Server (NTRS)

    Follett, W.; Ketchum, A.; Darian, A.; Hsu, Y.

    1996-01-01

    Several current rocket engine concepts such as the bell-annular tri-propellant engine, and the linear aerospike being proposed for the X-33 require unconventional three dimensional rocket nozzles which must conform to rectangular or sector shaped envelopes to meet integration constraints. These types of nozzles exist outside the current experience database, therefore, the application of efficient design methods for these propulsion concepts is critical to the success of launch vehicle programs. The objective of this work is to optimize several different nozzle configurations, including two- and three-dimensional geometries. Methodology includes coupling computational fluid dynamic (CFD) analysis to genetic algorithms and Taguchi methods as well as implementation of a streamline tracing technique. Results of applications are shown for several geometeries including: three dimensional thruster nozzles with round or super elliptic throats and rectangualar exits, two- and three-dimensional thrusters installed within a bell nozzle, and three dimensional thrusters with round throats and sector shaped exits. Due to the novel designs considered for this study, there is little experience which can be used to guide the effort and limit the design space. With a nearly infinite parameter space to explore, simple parametric design studies cannot possibly search the entire design space within the time frame required to impact the design cycle. For this reason, robust and efficient optimization methods are required to explore and exploit the design space to achieve high performance engine designs. Five case studies which examine the application of various techniques in the engineering environment are presented in this paper.

  6. Wall temperature distribution calculation for a rocket nozzle contour

    NASA Technical Reports Server (NTRS)

    Omori, S.; Gross, K. W.; Krebsbach, A.

    1972-01-01

    The JANNAF turbulent boundary layer (TBL) computer program, applicable to rocket nozzles, requires a wall temperature distribution among other input parameters to determine boundary layer behavior, heat transfer, and performance degradation. The inclusion of a complete regenerative cooling cycle model with associate geometry, material and fluid property data provides a capability to internally calculate wall temperature profiles on the hot gas and coolant flow-side, as well as the coolant flow bulk temperature variation. Besides the regular heat transfer and performance degradation calculations, the new concept can be used to optimize the cooling cycle, flow requirements, and cooling jacket geometry.

  7. Air-Breathing Rocket Engines

    NASA Technical Reports Server (NTRS)

    1998-01-01

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

  8. Engine throat/nozzle optics for plume spectroscopy

    NASA Technical Reports Server (NTRS)

    Bickford, R. L.; Duncan, D. B.

    1991-01-01

    The Task 2.0 Engine Throat/Nozzle Optics for Plume Spectroscopy, effort was performed under the NASA LeRC Development of Life Prediction Capabilities for Liquid Propellant Rocket Engines program. This Task produced the engineering design of an optical probe to enable spectroscopic measurements within the SSME main chamber. The probe mounts on the SSME nozzle aft manifold and collects light emitted from the throat plane and chamber. Light collected by the probe is transferred to a spectrometer through a fiber optic cable. The design analyses indicate that the probe will function throughout the engine operating cycle and is suitable for both test stand and flight operations. By detecting metallic emissions that are indicative of component degradation or incipient failure, engine shutdown can be initiated before catastrophic failure. This capability will protect valuable test stand hardware and provide enhanced mission safety.

  9. Design and Fabrication Development of J-2X Engine Metallic Nozzle Extension

    NASA Technical Reports Server (NTRS)

    Kopicz, C.; Gradl, P.

    2015-01-01

    Maximized rocket engine performance is in part derived from expanding combustion gasses through the rocket nozzle. For upper stage engines the nozzles can be quite large. On the J-2X engine, an uncooled extension of a regeneratively cooled nozzle is used to expand the combustion gasses to a targeted exit pressure which is defined by an altitude for the desired maximum performance. Creating a J-2X nozzle extension capable of surviving the loads of test and flight environments while meeting engine system performance requirements required development of new processes and facilities. Meeting the challenges of the development resulted in concurrent J-2X nozzle extension design and fabrication. This paper describes how some of the design and fabrication challenges were resolved.

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

    NASA Technical Reports Server (NTRS)

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

    1972-01-01

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

  11. Fundamental Understanding of Propellant/Nozzle Interaction for Rocket Nozzle Erosion Minimization Under Very High Pressure Conditions

    DTIC Science & Technology

    2005-08-31

    physical mechanisms of erosion. It is known that rocket motors with tungsten nozzles and non-aluminized AP- based propellants result in a significant...mentioned mechanical properties of the material change with temperature and the incorporation of this effect would also include the effect of heat transfer...evaluate their effects on nozzle erosion process. A nozzle erosion code has been updated to include comprehensive heterogeneous surface reaction mechanism

  12. Reusable Solid Rocket Motor Nozzle Joint-4 Thermal Analysis

    NASA Technical Reports Server (NTRS)

    Clayton, J. Louie

    2001-01-01

    This study provides for development and test verification of a thermal model used for prediction of joint heating environments, structural temperatures and seal erosions in the Space Shuttle Reusable Solid Rocket Motor (RSRM) Nozzle Joint-4. The heating environments are a result of rapid pressurization of the joint free volume assuming a leak path has occurred in the filler material used for assembly gap close out. Combustion gases flow along the leak path from nozzle environment to joint O-ring gland resulting in local heating to the metal housing and erosion of seal materials. Analysis of this condition was based on usage of the NASA Joint Pressurization Routine (JPR) for environment determination and the Systems Improved Numerical Differencing Analyzer (SINDA) for structural temperature prediction. Model generated temperatures, pressures and seal erosions are compared to hot fire test data for several different leak path situations. Investigated in the hot fire test program were nozzle joint-4 O-ring erosion sensitivities to leak path width in both open and confined joint geometries. Model predictions were in generally good agreement with the test data for the confined leak path cases. Worst case flight predictions are provided using the test-calibrated model. Analysis issues are discussed based on model calibration procedures.

  13. Pulse Detonation Rocket Engine Research at NASA Marshall

    NASA Technical Reports Server (NTRS)

    Morris, Christopher I.

    2003-01-01

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

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

    NASA Technical Reports Server (NTRS)

    Morris, Christopher I.

    2003-01-01

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

  15. RS-25 Rocket Engine Test

    NASA Image and Video Library

    2017-08-09

    The 8.5-minute test conducted at NASA’s Stennis Space Center is part of a series of tests designed to put the upgraded former space shuttle engines through the rigorous temperature and pressure conditions they will experience during a launch. The tests also support the development of a new controller, or “brain,” for the engine, which monitors engine status and communicates between the rocket and the engine, relaying commands to the engine and transmitting data back to the rocket.

  16. Recovering Aerodynamic Side Loads on Rocket Nozzles using Quasi-Static Strain-Gage Measurements

    NASA Technical Reports Server (NTRS)

    Brown, Andrew; Ruf, Joseph H.; McDaniels, David M.

    2009-01-01

    During over-expanded operation of rocket nozzles, which is defined to be when the exit pressure is greater than internal pressure over some part of the nozzle, the nozzle will experience a transverse forcing function due to the pressure differential across the nozzle wall. Over-expansion occurs during the nozzle start-up and shutdown transient, even in high-altitude engines, because most test facilities cannot completely reproduce the near-vacuum pressures at those altitudes. During this transient, the pressure differential moves axially down the nozzle as it becomes pressurized, but this differential is never perfectly symmetric circumferentially. The character of the forcing function is highly complex and defined by a series of restricted and free shock separations. The subject of this paper is the determination of the magnitude of this loading during sub-scale testing via measurement of the structural dynamic response of the nozzle and its support structure. An initial attempt at back-calculating this load using the inverse of the transfer function was performed, but this attempt was shown to be highly susceptible to numerical error. The final method chosen was to use statically calibrated strain data and to filter out the system fundamental frequency such that the measured response yields close to the correct dynamic loading function. This method was shown to capture 93% of the pressure spectral energy using controlled load shaker testing. This method is one of the only practical ways for the inverse determination of the forcing function for non-stationary excitations, and, to the authors' knowledge, has not been described in the literature to date.

  17. Applying Taguchi Methods To Brazing Of Rocket-Nozzle Tubes

    NASA Technical Reports Server (NTRS)

    Gilbert, Jeffrey L.; Bellows, William J.; Deily, David C.; Brennan, Alex; Somerville, John G.

    1995-01-01

    Report describes experimental study in which Taguchi Methods applied with view toward improving brazing of coolant tubes in nozzle of main engine of space shuttle. Dr. Taguchi's parameter design technique used to define proposed modifications of brazing process reducing manufacturing time and cost by reducing number of furnace brazing cycles and number of tube-gap inspections needed to achieve desired small gaps between tubes.

  18. Applying Taguchi Methods To Brazing Of Rocket-Nozzle Tubes

    NASA Technical Reports Server (NTRS)

    Gilbert, Jeffrey L.; Bellows, William J.; Deily, David C.; Brennan, Alex; Somerville, John G.

    1995-01-01

    Report describes experimental study in which Taguchi Methods applied with view toward improving brazing of coolant tubes in nozzle of main engine of space shuttle. Dr. Taguchi's parameter design technique used to define proposed modifications of brazing process reducing manufacturing time and cost by reducing number of furnace brazing cycles and number of tube-gap inspections needed to achieve desired small gaps between tubes.

  19. An Analysis of Nuclear-Rocket Nozzle Cooling

    NASA Technical Reports Server (NTRS)

    Robbins, William H.; Bachkin, Daniel; Medeiros, Arthur A.

    1960-01-01

    A nuclear-rocket regenerative-cooling analysis was conducted over a range of reactor power of 46 to 1600 megawatts and is summarized herein. Although the propellant (hydrogen) is characterized by a large heat-sink capacity, an analysis of the local heat-flux capability of the coolant at the nozzle throat indicated that, for conventional values of system pressure drop, the cooling capability was inadequate to maintain a selected wall temperature of 1440 R. Several techniques for improving the cooling capability were discussed, for example, high pressure drop, high wall temperature, refractory wall coatings, thin highly conductive walls, and film cooling. In any specific design a combination of methods will probably be utilized to achieve successful cooling.

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

    NASA Technical Reports Server (NTRS)

    Morris, Christopher

    2002-01-01

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

  1. Reentry aerodynamics forces and moments on the engine nozzle of the 146-inch solid rocket booster model 473 tested in MSFC 14 by 14 inch trisonic wind tunnel (SA30F)

    NASA Technical Reports Server (NTRS)

    Johnson, J. D.; Braddock, W. F.

    1975-01-01

    A test of a model of the Space Shuttle Solid Rocket Boosters (SRB's) was performed in a 14 x 14 inch Trisonic Wind Tunnel to determine the aerodynamic forces and moments imposed on the nozzle of the SRB during reentry. The model, with scale dimensions equal to 0.5479 of the actual SRB dimensions, was instrumented with a six-component force balance attached to the model nozzle so that only forces and moments acting on the nozzle were measured. A total of 137 runs (20 deg pitch polars) were performed during this test. The angle of attack ranged from 60 to 185 deg, the Reynolds number from 5.2 million to 7.6 million. The Mach numbers investigated were 1.96, 2.74, and 3.48. Five external protuberances were simulated. The effective roll angle simulated was 180 deg. The effects of three different heat shield configurations were investigated.

  2. Performance of a RBCC Engine in Rocket-Operation

    NASA Astrophysics Data System (ADS)

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

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

  3. Turbo Pump Fed Micro-Rocket Engine

    NASA Astrophysics Data System (ADS)

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

    2004-10-01

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

  4. Aircraft Engine Exhaust Nozzle System for Jet Noise Reduction

    NASA Technical Reports Server (NTRS)

    Thomas, Russell H. (Inventor); Czech, Michael J. (Inventor); Elkoby, Ronen (Inventor)

    2014-01-01

    The aircraft exhaust engine nozzle system includes a fan nozzle to receive a fan flow from a fan disposed adjacent to an engine disposed above an airframe surface of the aircraft, a core nozzle disposed within the fan nozzle and receiving an engine core flow, and a pylon structure connected to the core nozzle and structurally attached with the airframe surface to secure the engine to the aircraft.

  5. Rocket engine condition monitoring system

    SciTech Connect

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

    1989-01-01

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

  6. Formation of Vortex Structures in the Prenozzle Space of an Engine with a Vectorable Thrust Nozzle

    NASA Astrophysics Data System (ADS)

    Volkov, K. N.; Emel'yanov, V. N.; Denisikhin, S. V.

    2016-05-01

    A numerical simulation of the hydrodynamic effects arising in the process of work of the vectorable thrust nozzle of a solid-propellant rocket engine has been performed. The fields of the flows of combustion products in the channel of a charge, the prenozzle space, and the nozzle unit were calculated for different angles of vectoring of the nozzle. The distributions of the gasdynamic parameters of the flow of combustion products in the prenozzle space, corresponding to their efflux from the cylindrical and star-shaped channels of charges, were compared. The formation of a vortex flow in the neighborhood of the back cover of the nozzle was considered.

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

    NASA Technical Reports Server (NTRS)

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

    1993-01-01

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

  8. Simulation of the radiation-convective heat transfer in multinozzle assemblies of rocket engines

    NASA Astrophysics Data System (ADS)

    Volkov, N. N.; Volkova, L. I.; Tsatsuev, S. M.

    2012-12-01

    The method and results of numerical modeling of the radiation-convective heat transfer and thermal state in the systems of multinozzle rocket-engine (RE) assemblies are presented. The method is implemented in a form of a software module entered as the component into the program of calculation of the nonsteady thermal state of the RE nozzles. The results of calculation by the consolidated program are given, and the two-dimensional thermal fields on the external and internal surfaces of mouthpieces of the four-nozzle liquid rocket engine allow us to refine the thermal state of the nozzles themselves and evaluate the radiation heat flows in the engine module.

  9. Rocket-Based Combined Cycle Engine Concept Development

    NASA Technical Reports Server (NTRS)

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

    2001-01-01

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

  10. C/C and C/SiC Nozzle Extensions - A Breakthrough to Improve Upper and Lower Stage Engines Performance

    NASA Astrophysics Data System (ADS)

    Pichon, T.; /Lacombe, A.; Mercier, A.; Ferrey, A.

    2002-01-01

    The need to increase the payload capacity of the current launchers drives rocket engine manufacturers to seek higher thrust level, specific impulse and thrust to weight ratio. A particularly efficient way to do this is the use of increased expansion ratio nozzle extensions for upper stage engines, and using thermostructural composite materials in order to allow higher temperature material limitations and to decrease mass. The latter is applicable to both upper and lower stage engines. Up to the mid 90s, the use of composite nozzles has been limited to solid rocket nozzles, but recent developments led to flight qualification on liquid rocket engines, on the RL10-B2 engine of the DELTA III launcher. This engine is equipped with a large extendible Carbon/Carbon Novoltex Sepcarb nozzle developed by Snecma Propulsion Solide under a contract from Pratt &Whitney San Jose. This paper describes the technological background of Snecma Propulsion Solide concerning the design and manufacturing of large size composite nozzles for liquid rocket engines. It provides an up-to-date status of the demonstrations already performed on different engines (HM7, RL10-B2 in particular) and details all the recent progress on technical and manufacturing performance. The manufacturing process has also been improved and simplified in order to allow the manufacturing of larger scale nozzles, at lower cost. Finally, this paper evidences that this technology is today mature and is ready to be implemented on existing or future liquid rocket engines being developed.

  11. Verification of thermal analysis codes for modeling solid rocket nozzles

    NASA Technical Reports Server (NTRS)

    Keyhani, M.

    1993-01-01

    One of the objectives of the Solid Propulsion Integrity Program (SPIP) at Marshall Space Flight Center (MSFC) is development of thermal analysis codes capable of accurately predicting the temperature field, pore pressure field and the surface recession experienced by decomposing polymers which are used as thermal barriers in solid rocket nozzles. The objective of this study is to provide means for verifications of thermal analysis codes developed for modeling of flow and heat transfer in solid rocket nozzles. In order to meet the stated objective, a test facility was designed and constructed for measurement of the transient temperature field in a sample composite subjected to a constant heat flux boundary condition. The heating was provided via a steel thin-foil with a thickness of 0.025 mm. The designed electrical circuit can provide a heating rate of 1800 W. The heater was sandwiched between two identical samples, and thus ensure equal power distribution between them. The samples were fitted with Type K thermocouples, and the exact location of the thermocouples were determined via X-rays. The experiments were modeled via a one-dimensional code (UT1D) as a conduction and phase change heat transfer process. Since the pyrolysis gas flow was in the direction normal to the heat flow, the numerical model could not account for the convection cooling effect of the pyrolysis gas flow. Therefore, the predicted values in the decomposition zone are considered to be an upper estimate of the temperature. From the analysis of the experimental and the numerical results the following are concluded: (1) The virgin and char specific heat data for FM 5055 as reported by SoRI can not be used to obtain any reasonable agreement between the measured temperatures and the predictions. However, use of virgin and char specific heat data given in Acurex report produced good agreement for most of the measured temperatures. (2) Constant heat flux heating process can produce a much higher

  12. Rocket engine development

    NASA Astrophysics Data System (ADS)

    Butler, Kathleen N.

    1993-04-01

    The paper discusses the National Launch System Space Transportation Main Engine (STME) development program, which is a joint USAF/NASA project to produce a heavy-lift launch vehicle at a cost significantly below that of current systems and which incorporates efforts by three member companies. The STME project is organized into seven product development teams (systems engineering and integration, engine system, thrust chamber assembly, fuel turbopump assembly, oxidizer turbopump assembly, control system, and engine hardware), each led by one of the three member companies along with a NASA deputy. Most teams are assigned several components, each having a product-oriented concurrent engineering team involving all relevant disciplines.

  13. Fluid thrust control system. [for liquid propellant rocket engines

    NASA Technical Reports Server (NTRS)

    Howell, W. L.; Jansen, H. B.; Lehmann, E. N. (Inventor)

    1968-01-01

    A pure fluid thrust control system is described for a pump-fed, regeneratively cooled liquid propellant rocket engine. A proportional fluid amplifier and a bistable fluid amplifier control overshoot in the starting of the engine and take it to a predetermined thrust. An ejector type pump is provided in the line between the liquid hydrogen rocket nozzle heat exchanger and the turbine driving the fuel pump to aid in bringing the fluid at this point back into the regular system when it is not bypassed. The thrust control system is intended to function in environments too severe for mechanical controls.

  14. Centrifugal pumps for rocket engines

    NASA Technical Reports Server (NTRS)

    Campbell, W. E.; Farquhar, J.

    1974-01-01

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

  15. Effect of injector configuration in rocket nozzle film cooling

    NASA Astrophysics Data System (ADS)

    Kumar, A. Lakshya; Pisharady, J. C.; Shine, S. R.

    2016-04-01

    Experimental and numerical investigations are carried out to analyze the effect of coolant injector configuration on overall film cooling performance in a divergent section of a rocket nozzle. Two different injector orientations are investigated: (1) shaped slots with a divergence angle of 15° (semi-divergent injector) (2) fully divergent slot (fully divergent injector). A 2-dimensional, axis-symmetric, multispecies computational model using finite volume formulation has been developed and validated against the experimental data. The experiments provided a consistent set of measurements for cooling effectiveness for different blowing ratios ranging from 3.7 to 6. Results show that the semi divergent configuration leads to higher effectiveness compared to fully divergent slot at all blowing ratios. The spatially averaged effectiveness results show that the difference between the two configurations is significant at higher blowing ratios. The increase in effectiveness was around 2 % at BR = 3.7 whereas it was around 12 % in the case of BR = 6. Numerical results show the presence of secondary flow recirculation zones near the jet exit for both the injectors. An additional recirculation zone present in the case of fully divergent injector caused an increase in mixing of the coolant and mainstream, and a reduction in film cooling performance.

  16. Rocket Engine Altitude Simulation Technologies

    NASA Technical Reports Server (NTRS)

    Woods, Jody L.; Lansaw, John

    2010-01-01

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

  17. Proposed Flight Research of a Dual-Bell Rocket Nozzle Using the NASA F-15 Airplane

    NASA Technical Reports Server (NTRS)

    Jones, Daniel S.; Bui, Trong T.; Ruf, Joseph H.

    2013-01-01

    For more than a half-century, several types of altitude-compensating rocket nozzles have been proposed and analyzed, but very few have been adequately tested in a relevant flight environment. One type of altitude-compensating nozzle is the dual-bell rocket nozzle, which was first introduced into literature in 1949. Despite the performance advantages that have been predicted, both analytically and through static test data, the dual-bell nozzle has still not been adequately tested in a relevant flight environment. This paper proposes a method for conducting testing and research with a dual-bell rocket nozzle in a flight environment. We propose to leverage the existing NASA F-15 airplane and Propulsion Flight Test Fixture as the flight testbed, with the dual-bell nozzle operating during captive-carried flights, and with the nozzle subjected to a local flow field similar to that of a launch vehicle. The primary objective of this effort is not only to advance the technology readiness level of the dual-bell nozzle, but also to gain a greater understanding of the nozzle mode transitional sensitivity to local flow-field effects, and to quantify the performance benefits with this technology. The predicted performance benefits are significant, and may result in reducing the cost of delivering payloads to low-Earth orbit.

  18. Proposed Flight Research of a Dual-Bell Rocket Nozzle Using the NASA F-15 Airplane

    NASA Technical Reports Server (NTRS)

    Jones, Daniel S.; Bui, Trong T.; Ruf, Joseph H.

    2013-01-01

    For more than a half-century, several types of altitude-compensating rocket nozzles have been proposed and analyzed, but very few have been adequately tested in a relevant flight environment. One type of altitude-compensating nozzle is the dual-bell rocket nozzle, which was first introduced into literature in 1949. Despite the performance advantages that have been predicted, both analytically and through static test data, the dual-bell nozzle has still not been adequately tested in a relevant flight environment. This presentation proposes a method for conducting testing and research with a dual-bell rocket nozzle in a flight environment. We propose to leverage the existing NASA F-15 airplane and Propulsion Flight Test Fixture as the flight testbed, with the dual-bell nozzle operating during captive-carried flights, and with the nozzle subjected to a local flow field similar to that of a launch vehicle. The primary objective of this effort is not only to advance the technology readiness level of the dual-bell nozzle, but also to gain a greater understanding of the nozzle mode transitional sensitivity to local flow-field effects, and to quantify the performance benefits with this technology. The predicted performance benefits are significant, and may result in reducing the cost of delivering payloads to low-Earth orbit.

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

    NASA Technical Reports Server (NTRS)

    Morris, Christopher I.

    2002-01-01

    Pulsed detonation rocket engines (PDREs) have generated considerable research interest in recent years as a chemical propulsion system potentially offering improved performance and reduced complexity compared to conventional rocket engines. The detonative mode of combustion employed by these devices offers a thermodynamic advantage over the constant-pressure deflagrative combustion mode used in conventional rocket engines and gas turbines. However, while this theoretical advantage has spurred a great deal of interest in building PDRE devices, the unsteady blowdown process intrinsic to the PDRE has made realistic estimates of the actual propulsive performance problematic. The recent review article by Kailasanath highlights some of the difficulties in comparing the available experimental measurements with numerical models. In a previous paper by the author, parametric studies of the performance of a single, straight-tube PDRE were reported. A 1-D, unsteady method of characteristics code, employing a constant-gamma assumption behind the detonation front, was developed for that study. Models of this type are computationally inexpensive, and are particularly useful for parametric performance comparisons. For example, a plot showing the specific impulse of various PDRE and steady-state rocket engine (SSRE) configurations as a function of blowdown pressure ratio. The performance curves clearly indicate that a straight-tube PDRE is superior in specific impulse to a SSRE with a sonic nozzle over the entire range of pressure ratios. Note, however, that a straight-tube PDRE in general does not compare favorably to a SSRE fitted with an optimized de Laval supersonic nozzle, particularly at the high pressure ratios typical for boost or in-space rocket applications. However, the calculations also show that if a dynamically optimized, supersonic de Laval nozzle could be could be fitted to a PDRE, then the specific impulse of the device would exceed that of a comparable SSRE

  20. Reusable Rocket Engine Maintenance Study

    NASA Technical Reports Server (NTRS)

    Macgregor, C. A.

    1982-01-01

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

  1. AJ26 rocket engine test

    NASA Image and Video Library

    2010-11-10

    Fire and steam signal a successful test firing of Orbital Sciences Corporation's Aerojet AJ26 rocket engine at John C. Stennis Space Center. AJ26 engines will be used to power Orbital's Taurus II space vehicle on commercial cargo flights to the International Space Station. On Nov. 10, operators at Stennis' E-1 Test Stand conducted a 10-second test fire of the engine, the first of a series of three verification tests. Orbital has partnered with NASA to provide eight missions to the ISS by 2015.

  2. Upper Stage Engine Composite Nozzle Extensions

    NASA Technical Reports Server (NTRS)

    Valentine, Peter G.; Allen, Lee R.; Gradl, Paul R.; Greene, Sandra E.; Sullivan, Brian J.; Weller, Leslie J.; Koenig, John R.; Cuneo, Jacques C.; Thompson, James; Brown, Aaron; Shigley, John K.; Dovey, Henry N.; Roberts, Robert K.

    2015-01-01

    Carbon-carbon (C-C) composite nozzle extensions are of interest for use on a variety of launch vehicle upper stage engines and in-space propulsion systems. The C-C nozzle extension technology and test capabilities being developed are intended to support National Aeronautics and Space Administration (NASA) and United States Air Force (USAF) requirements, as well as broader industry needs. Recent and on-going efforts at the Marshall Space Flight Center (MSFC) are aimed at both (a) further developing the technology and databases for nozzle extensions fabricated from specific CC materials, and (b) developing and demonstrating low-cost capabilities for testing composite nozzle extensions. At present, materials development work is concentrating on developing a database for lyocell-based C-C that can be used for upper stage engine nozzle extension design, modeling, and analysis efforts. Lyocell-based C-C behaves in a manner similar to rayon-based CC, but does not have the environmental issues associated with the use of rayon. Future work will also further investigate technology and database gaps and needs for more-established polyacrylonitrile- (PAN-) based C-C's. As a low-cost means of being able to rapidly test and screen nozzle extension materials and structures, MSFC has recently established and demonstrated a test rig at MSFC's Test Stand (TS) 115 for testing subscale nozzle extensions with 3.5-inch inside diameters at the attachment plane. Test durations of up to 120 seconds have been demonstrated using oxygen/hydrogen propellants. Other propellant combinations, including the use of hydrocarbon fuels, can be used if desired. Another test capability being developed will allow the testing of larger nozzle extensions (13.5- inch inside diameters at the attachment plane) in environments more similar to those of actual oxygen/hydrogen upper stage engines. Two C-C nozzle extensions (one lyocell-based, one PAN-based) have been fabricated for testing with the larger

  3. Nozzle Numerical Analysis Of The Scimitar Engine

    NASA Astrophysics Data System (ADS)

    Battista, F.; Marini, M.; Cutrone, L.

    2011-05-01

    This work describes part of the activities on the LAPCAT-II A2 vehicle, in which starting from the available conceptual vehicle design and the related pre- cooled turbo-ramjet engine called SCIMITAR, well- thought assumptions made for performance figures of different components during the iteration process within LAPCAT-I will be assessed in more detail. In this paper it is presented a numerical analysis aimed at the design optimization of the nozzle contour of the LAPCAT A2 SCIMITAR engine designed by Reaction Engines Ltd. (REL) (see Figure 1). In particular, nozzle shape optimization process is presented for cruise conditions. All the computations have been carried out by using the CIRA C3NS code in non equilibrium conditions. The effect of considering detailed or reduced chemical kinetic schemes has been analyzed with a particular focus on the production of pollutants. An analysis of engine performance parameters, such as thrust and combustion efficiency has been carried out.

  4. Combustor nozzles in gas turbine engines

    DOEpatents

    Johnson, Thomas Edward; Keener, Christopher Paul; Stewart, Jason Thurman; Ostebee, Heath Michael

    2017-09-12

    A micro-mixer nozzle for use in a combustor of a combustion turbine engine, the micro-mixer nozzle including: a fuel plenum defined by a shroud wall connecting a periphery of a forward tube sheet to a periphery of an aft tubesheet; a plurality of mixing tubes extending across the fuel plenum for mixing a supply of compressed air and fuel, each of the mixing tubes forming a passageway between an inlet formed through the forward tubesheet and an outlet formed through the aft tubesheet; and a wall mixing tube formed in the shroud wall.

  5. Device for installing rocket engines

    NASA Technical Reports Server (NTRS)

    George, T. R., Jr. (Inventor)

    1976-01-01

    A device for installing rocket engines is reported that is supported at a cant relative to vertical by an axially extensible, tiltable pedestal. A lifting platform supports the rocket engine at its thrust chamber exit, including a mount having a concentric base characterized by a concave bearing surface, a plurality of uniformly spaced legs extended radially from the base, and an annular receiver coaxially aligned with the base and affixed to the distal ends of said legs for receiving the thrust chamber exit. The lifting platform rests on a seat concentrically related to the pedestal and affixed to an extended end portion thereof having a convex bearing surface mated in sliding engagement with the concave bearing surface of the annular base for accommodating a rocking motion of the platform.

  6. Experimental performance of a high-area-ratio rocket nozzle at high combustion chamber pressure

    NASA Technical Reports Server (NTRS)

    Jankovsky, Robert S.; Kazaroff, John M.; Pavli, Albert J.

    1996-01-01

    An experimental investigation was conducted to determine the thrust coefficient of a high-area-ratio rocket nozzle at combustion chamber pressures of 12.4 to 16.5 MPa (1800 to 2400 psia). A nozzle with a modified Rao contour and an expansion area ratio of 1025:1 was tested with hydrogen and oxygen at altitude conditions. The same nozzle, truncated to an area ratio of 440:1, was also tested. Values of thrust coefficient are presented along with characteristic exhaust velocity efficiencies, nozzle wall temperatures, and overall thruster specific impulse.

  7. Rocket Engine Numerical Simulator (RENS)

    NASA Technical Reports Server (NTRS)

    Davidian, Kenneth O.

    1997-01-01

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

  8. Rocket Engine Numerical Simulator (RENS)

    NASA Technical Reports Server (NTRS)

    Davidian, Kenneth O.

    1997-01-01

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

  9. Liquid rocket engine turbopump gears

    NASA Technical Reports Server (NTRS)

    1974-01-01

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

  10. Measuring Model Rocket Engine Thrust Curves

    ERIC Educational Resources Information Center

    Penn, Kim; Slaton, William V.

    2010-01-01

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

  11. Measuring Model Rocket Engine Thrust Curves

    ERIC Educational Resources Information Center

    Penn, Kim; Slaton, William V.

    2010-01-01

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

  12. Liquid Rocket Engine Testing

    DTIC Science & Technology

    2016-10-21

    technologies and advancing TRL’s • Our product is data • 100’s of sensor channels • Extreme care goes into signal conditioning and processing • Real time...and storable propellants • Liquid Oxygen (LOX) • RP-1 (Kerosene, very similar to JP-8) • Liquid Hydrogen • Liquid methane • Pressure = Performance in...demonstrator of a Kerosene-LOX, 250,000 lbf, 3000 psi oxygen -rich staged combustion engine (ORSC) • AFRL’s Test Stand 2A recently completed a two-year

  13. Numerical and experimental study of liquid breakup process in solid rocket motor nozzle

    NASA Astrophysics Data System (ADS)

    Yen, Yi-Hsin

    Rocket propulsion is an important travel method for space exploration and national defense, rockets needs to be able to withstand wide range of operation environment and also stable and precise enough to carry sophisticated payload into orbit, those engineering requirement makes rocket becomes one of the state of the art industry. The rocket family have been classified into two major group of liquid and solid rocket based on the fuel phase of liquid or solid state. The solid rocket has the advantages of simple working mechanism, less maintenance and preparing procedure and higher storage safety, those characters of solid rocket make it becomes popular in aerospace industry. Aluminum based propellant is widely used in solid rocket motor (SRM) industry due to its avalibility, combusion performance and economical fuel option, however after aluminum react with oxidant of amonimum perchrate (AP), it will generate liquid phase alumina (Al2O3) as product in high temperature (2,700˜3,000 K) combustion chamber enviornment. The liquid phase alumina particles aggromorate inside combustion chamber into larger particle which becomes major erosion calprit on inner nozzle wall while alumina aggromorates impinge on the nozzle wall surface. The erosion mechanism result nozzle throat material removal, increase the performance optimized throat diameter and reduce nozzle exit to throat area ratio which leads to the reduction of exhaust gas velocity, Mach number and lower the propulsion thrust force. The approach to avoid particle erosion phenomenon taking place in SRM's nozzle is to reduce the alumina particle size inside combustion chamber which could be done by further breakup of the alumina droplet size in SRM's combustion chamber. The study of liquid breakup mechanism is an important means to smaller combustion chamber alumina droplet size and mitigate the erosion tack place on rocket nozzle region. In this study, a straight two phase air-water flow channel experiment is set up

  14. A detailed description of the uncertainty analysis for High Area Ratio Rocket Nozzle tests at the NASA Lewis Research Center

    NASA Technical Reports Server (NTRS)

    Davidian, Kenneth J.; Dieck, Ronald H.; Chuang, Isaac

    1987-01-01

    A preliminary uncertainty analysis has been performed for the High Area Ratio Rocket Nozzle test program which took place at the altitude test capsule of the Rocket Engine Test Facility at the NASA Lewis Research Center. Results from the study establish the uncertainty of measured and calculated parameters required for the calculation of rocket engine specific impulse. A generalized description of the uncertainty methodology used is provided. Specific equations and a detailed description of the analysis are presented. Verification of the uncertainty analysis model was performed by comparison with results from the experimental program's data reduction code. Final results include an uncertainty for specific impulse of 1.30 percent. The largest contributors to this uncertainty were calibration errors from the test capsule pressure and thrust measurement devices.

  15. NASA Orbit Transfer Rocket Engine Technology Program

    NASA Technical Reports Server (NTRS)

    1984-01-01

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

  16. Rocketdyne's XLR-132 rocket engine

    NASA Astrophysics Data System (ADS)

    Pauckert, Ronald P.; Billings, Ralph Bancroft

    In the mid-1980s Rocketdyne entered into a contract with the U.S. Air Force, Phillips Laboratory, to develop and demonstrate the advanced technology required for a high-performance, storable-propellant liquid rocket engine. The engine is designed to provide propulsion for a variety of needs, including spacecraft deployment, sortie operations, orbit transfer vehicle propulsion, and satellite maneuvering. Significant potential payload benefits of propulsion systems using high-performance pump-fed engines (relative to pressure-fed systems) are shown. Additional design goals were high reliability, adaptability to the various uses, durability, and light weight. The propellants selected were monomethylhydrazine and nitrogen tetroxide. The engine, designated XLR-132 by the Air Force, is Rocketdyne's model RS47.

  17. Orbit transfer rocket engine technology program

    NASA Astrophysics Data System (ADS)

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

    1993-10-01

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

  18. Orbit transfer rocket engine technology program

    NASA Technical Reports Server (NTRS)

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

    1993-01-01

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

  19. Liquid rocket engine test facility engineering challenges

    NASA Astrophysics Data System (ADS)

    Ellerbrock, Hartwig; Ziegenhagen, Stefan

    2006-12-01

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

  20. Rapid Fabrication Techniques for Liquid Rocket Channel Wall Nozzles

    NASA Technical Reports Server (NTRS)

    Gradl, Paul R.

    2016-01-01

    The functions of a regeneratively-cooled nozzle are to (1) expand combustion gases to increase exhaust gas velocity while, (2) maintaining adequate wall temperatures to prevent structural failure, and (3) transfer heat from the hot gases to the coolant fluid to promote injector performance and stability. Regeneratively-cooled nozzles are grouped into two categories: tube-wall nozzles and channel wall nozzles. A channel wall nozzle is designed with an internal liner containing a series of integral coolant channels that are closed out with an external jacket. Manifolds are attached at each end of the nozzle to distribute coolant to and away from the channels. A variety of manufacturing techniques have been explored for channel wall nozzles, including state of the art laser-welded closeouts and pressure-assisted braze closeouts. This paper discusses techniques that NASA MSFC is evaluating for rapid fabrication of channel wall nozzles that address liner fabrication, slotting techniques and liner closeout techniques. Techniques being evaluated for liner fabrication include large-scale additive manufacturing of freeform-deposition structures to create the liner blanks. Abrasive water jet milling is being evaluated for cutting the complex coolant channel geometries. Techniques being considered for rapid closeout of the slotted liners include freeform deposition, explosive bonding and Cold Spray. Each of these techniques, development work and results are discussed in further detail in this paper.

  1. Pulse Detonation Rocket Engine Research at NASA Marshall

    NASA Technical Reports Server (NTRS)

    Morris, Christopher I.

    2003-01-01

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

  2. Instrumentation for In-Flight SSME Rocket Engine Plume Spectroscopy

    NASA Technical Reports Server (NTRS)

    Madzsar, George C.; Bickford, Randall L.; Duncan, David B.

    1994-01-01

    This paper describes instrumentation that is under development for an in-flight demonstration of a plume spectroscopy system on the space shuttle main engine. The instrumentation consists of a nozzle mounted optical probe for observation of the plume, and a spectrometer for identification and quantification of plume content. This instrumentation, which is intended for use as a diagnostic tool to detect wear and incipient failure in rocket engines, will be validated by a hardware demonstration on the Technology Test Bed engine at the Marshall Space Flight Center.

  3. Methane Dual Expander Aerospike Nozzle Rocket Engine

    DTIC Science & Technology

    2012-03-22

    v Acknowledgments I would like to thank my thesis advisor, Lt. Col Hartsfield, who has put great effort in discussing, teaching , and...42 NPSS Thermochemistry ................................................................................................44 Methane FPT Generation...Next is a more focused description of how NPSS uses thermochemistry and an explanation on the development of the different fluid property tables

  4. Engineers demonstrate the pocket rocket

    NASA Technical Reports Server (NTRS)

    1995-01-01

    Part of Stennis Space Center's mission with its traveling exhibits is to educate the younger generation on how propulsion systems work. A popular tool is the 'pocket rocket,' which demonstrates how a hybrid rocket works. A hybrid rocket is a cross breed between a solid fuel rocket and a liquid fuel rocket.

  5. Engineers demonstrate the pocket rocket

    NASA Technical Reports Server (NTRS)

    1996-01-01

    Part of Stennis Space Center's mission with its traveling exhibits is to educate the younger generation on how propulsion systems work. A popular tool is the 'pocket rocket,' which demonstrates how a hybrid rocket works. A hybrid rocket is a cross breed between a solid fuel rocket and a liquid fuel rocket.

  6. Measuring Model Rocket Engine Thrust Curves

    NASA Astrophysics Data System (ADS)

    Penn, Kim; Slaton, William V.

    2010-12-01

    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 LabPro2 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 simple engine holder can be constructed and used with Vernier's LabPro and force probe to record data that students can use to compare to sample data from the rocket manufacturer or the National Association of Rocketry's3 engine certification sheets, calculate total impulse, and make predictions for model rocket launches. PASCO markets a rocket engine test bracket4 that mounts to its PASPORT force sensor for similar measurements. The engine holder described here is very economical, and all the parts can be obtained from a local hardware store or home center.

  7. Jet Engine Exhaust Nozzle Flow Effector

    NASA Technical Reports Server (NTRS)

    Turner, Travis L. (Inventor); Cano, Roberto J. (Inventor); Silcox, Richard J. (Inventor); Buehrle, Ralph D. (Inventor); Cagle, Christopher M. (Inventor); Cabell, Randolph H. (Inventor); Hilton, George C. (Inventor)

    2011-01-01

    A jet engine exhaust nozzle flow effector is a chevron formed with a radius of curvature with surfaces of the flow effector being defined and opposing one another. At least one shape memory alloy (SMA) member is embedded in the chevron closer to one of the chevron's opposing surfaces and substantially spanning from at least a portion of the chevron's root to the chevron's tip.

  8. Jet Engine Exhaust Nozzle Flow Effector

    NASA Technical Reports Server (NTRS)

    Turner, Travis L. (Inventor); Cano, Roberto J. (Inventor); Silox, Richard J. (Inventor); Buehrle, Ralph D. (Inventor); Cagle, Christopher M. (Inventor); Cabell, Randolph H. (Inventor); Hilton, George C. (Inventor)

    2014-01-01

    A jet engine exhaust nozzle flow effector is a chevron formed with a radius of curvature with surfaces of the flow effector being defined and opposing one another. At least one shape memory alloy (SMA) member is embedded in the chevron closer to one of the chevron's opposing surfaces and substantially spanning from at least a portion of the chevron's root to the chevron's tip.

  9. Altitude-Compensating Nozzle (ACN) Project: Planning for Dual-Bell Rocket Nozzle Flight Testing on the NASA F-15B

    NASA Technical Reports Server (NTRS)

    Jones, Daniel S.; Bui, Trong T.; Ruf, Joseph H.

    2013-01-01

    For more than a half-century, several types of altitude-compensating nozzles have been proposed and analyzed, but very few have been adequately tested in a relevant flight environment. One type of altitude-compensating nozzle is the dual-bell rocket nozzle, which was first introduced into literature in 1949. Although the dual-bell rocket nozzle has been thoroughly studied, this nozzle has still not been tested in a relevant flight environment. This poster presents the top-level rationale and preliminary plans for conducting flight research with the dual-bell rocket nozzle, while exhausting the plume into the freestream flow field at various altitudes. The primary objective is to gain a greater understanding of the nozzle plume sensitivity to freestream flight effects, which will also include detailed measurements of the plume mode transition within the nozzle. To accomplish this goal, the NASA F-15B is proposed as the testbed for advancing the technology readiness level of this greatly-needed capability. All proposed tests include the quantitative performance analysis of the dual-bell rocket nozzle as compared with the conventional-bell nozzle.

  10. Reusable rocket engine optical condition monitoring

    NASA Technical Reports Server (NTRS)

    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.

  11. Casting propellant in rocket engine

    NASA Technical Reports Server (NTRS)

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

    1976-01-01

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

  12. Development Status of Reusable Rocket Engine

    NASA Astrophysics Data System (ADS)

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

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

  13. Conjugate conductive, convective, and radiative heat transfer in rocket engines

    SciTech Connect

    Naraghi, M.H.N.; DeLise, J.C.

    1995-12-31

    A comprehensive conductive, convective and radiative model for thermal analysis of rocket thrust chambers and nozzles is presented. In this model, the rocket thrust chamber and nozzle are subdivided into a number of stations along the longitudinal direction. At each station a finite element scheme is used to evaluate wall temperature distribution. The hot-gas-side convective heat transport is evaluated by numerically solving the compressible boundary layer equations and the radiative fluxes are evaluated by implementing an exchange factor scheme. The convective heat flux in the cooling channel is modeled based on the existing closed form correlations for rocket cooling channels. The conductive, convective and radiative processes are conjugated through an iterative procedure. The hot-gas-side heat transfer coefficients evaluated based on this model are compared to the experimental results reported in the literature. The computed convective heat transfer coefficients agree very well with experimental data for most of the engine except the throat where a discrepancy of approximately 20% exists. The model is applied to a typical regeneratively cooled rocket engine and the resulting wall temperature and heat flux distribution are presented.

  14. Dual-fuel, dual-mode rocket engine

    NASA Technical Reports Server (NTRS)

    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.

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

    NASA Astrophysics Data System (ADS)

    Quentmeyer, Richard J.; Roncace, Elizabeth A.

    1993-07-01

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

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

    NASA Technical Reports Server (NTRS)

    Quentmeyer, Richard J.; Roncace, Elizabeth A.

    1993-01-01

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

  17. Axisymmetric Numerical Modeling of Pulse Detonation Rocket Engines

    NASA Technical Reports Server (NTRS)

    Morris, Christopher I.

    2005-01-01

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

  18. Experimental thrust performance of a high-area-ratio rocket nozzle

    NASA Technical Reports Server (NTRS)

    Pavli, Albert J.; Kacynski, Kenneth J.; Smith, Tamara A.

    1987-01-01

    An experimental investigation was conducted to determine the thrust performance attainable from high-area-ratio rocket nozzles. A modified Rao-contoured nozzle with an expansion area of 1030 was test fired with hydrogen-oxygen propellants at altitude conditions. The nozzle was also tested as a truncated nozzle, at an expansion area ratio of 428. Thrust coefficient and thrust coefficient efficiency values are presented for each configuration at various propellant mixture ratios (oxygen/fuel). Several procedural techniques were developed permitting improved measurement of nozzle performance. The more significant of these were correcting the thrust for the aneroid effects, determining the effective chamber pressure, and referencing differential pressure transducers to a vacuum reference tank.

  19. Computer codes for thermal analysis of a solid rocket motor nozzle

    NASA Technical Reports Server (NTRS)

    Chauhan, Rajinder Singh

    1988-01-01

    A number of computer codes are available for performing thermal analysis of solid rocket motor nozzles. Aerotherm Chemical Equilibrium (ACE) computer program can be used to perform one-dimensional gas expansion to determine the state of the gas at each location of a nozzle. The ACE outputs can be used as input to a computer program called Momentum/Energy Integral Technique (MEIT) for predicting boundary layer development development, shear, and heating on the surface of the nozzle. The output from MEIT can be used as input to another computer program called Aerotherm Charring Material Thermal Response and Ablation Program (CMA). This program is used to calculate oblation or decomposition response of the nozzle material. A code called Failure Analysis Nonlinear Thermal and Structural Integrated Code (FANTASTIC) is also likely to be used for performing thermal analysis of solid rocket motor nozzles after the program is duly verified. A part of the verification work on FANTASTIC was done by using one and two dimension heat transfer examples with known answers. An attempt was made to prepare input for performing thermal analysis of the CCT nozzle using the FANTASTIC computer code. The CCT nozzle problem will first be solved by using ACE, MEIT, and CMA. The same problem will then be solved using FANTASTIC. These results will then be compared for verification of FANTASTIC.

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

    NASA Technical Reports Server (NTRS)

    Morris, Christopher I.

    2004-01-01

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

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

    SciTech Connect

    Youngblood, Stewart

    2015-08-01

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

  2. Aeroelastic stability analysis of flexible overexpanded rocket nozzle

    NASA Astrophysics Data System (ADS)

    Bekka, N.; Sellam, M.; Chpoun, A.

    2016-07-01

    The aim of this paper is to present a new aeroelastic stability model taking into account the viscous effects for a supersonic nozzle flow in overexpanded regimes. This model is inspired by the Pekkari model which was developed initially for perfect fluid flow. The new model called the "Modified Pekkari Model" (MPM) considers a more realistic wall pressure profile for the case of a free shock separation inside the supersonic nozzle using the free interaction theory of Chapman. To reach this objective, a code for structure computation coupled with aerodynamic excitation effects is developed that allows the analysis of aeroelastic stability for the overexpanded nozzles. The main results are presented in a comparative manner using existing models (Pekkari model and its extended version) and the modified Pekkari model developed in this work.

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

    NASA Technical Reports Server (NTRS)

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

    2010-01-01

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

  4. Mixing Nozzles for Liquid-Fuel Rocket Motors

    DTIC Science & Technology

    1949-08-01

    Introduct ion 3 2 Performance of a simple swirl nozzle used as an emulsifier 3 2.1 Apparatus 3 2.2 Experimental results 4 2.3 Characteristics of the emulsion...only one liquid passes through the nozzle , 2. 2 Purformance of a simple swirl no:,zle used an an emulsifier 2.1 ,pparatus A noszle Tro , i h rirogun...B respectivel- thfraono the avrera,e percentage of waterfor each series of VA . , thc . urc in llcolui,n s or 6 for onenozzle are C.parated by a tli

  5. Characterization of Rotating Detonation Engine Exhaust Through Nozzle Guide Vanes

    DTIC Science & Technology

    2013-03-21

    THROUGH NOZZLE GUIDE VANES THESIS Presented to the Faculty Department of Aeronautics and Astronautics Graduate School of Engineering and Management Air...the first Nozzle Guide Vane (NGV) section from a T63 gas turbine engine to a 6 inch diameter RDE was designed and built for this study. Pressure...CHARACTERIZATION OF ROTATING DETONATION ENGINE EXHAUST THROUGH NOZZLE GUIDE VANES THESIS Nick D. DeBarmore, Second Lieutenant, USAF AFIT/GAE/ENY/13

  6. Commercial Rocket Engine Readied for Test

    NASA Image and Video Library

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

  7. NASA Engineer Examines the Design of a Regeneratively-Cooled Rocket Engine

    NASA Image and Video Library

    1958-12-21

    An engineer at the National Aeronautics and Space Administration (NASA) Lewis Research Center examines a drawing showing the assembly and details of a 20,000-pound thrust regeneratively cooled rocket engine. The engine was being designed for testing in Lewis’ new Rocket Engine Test Facility, which began operating in the fall of 1957. The facility was the largest high-energy test facility in the country that was capable of handling liquid hydrogen and other liquid chemical fuels. The facility’s use of subscale engines up to 20,000 pounds of thrust permitted a cost-effective method of testing engines under various conditions. The Rocket Engine Test Facility was critical to the development of the technology that led to the use of hydrogen as a rocket fuel and the development of lightweight, regeneratively-cooled, hydrogen-fueled rocket engines. Regeneratively-cooled engines use the cryogenic liquid hydrogen as both the propellant and the coolant to prevent the engine from burning up. The fuel was fed through rows of narrow tubes that surrounded the combustion chamber and nozzle before being ignited inside the combustion chamber. The tubes are visible in the liner sitting on the desk. At the time, Pratt and Whitney was designing a 20,000-pound thrust liquid-hydrogen rocket engine, the RL-10. Two RL-10s would be used to power the Centaur second-stage rocket in the 1960s. The successful development of the Centaur rocket and the upper stages of the Saturn V were largely credited to the work carried out Lewis.

  8. Development of high temperature materials for solid propellant rocket nozzle applications

    NASA Technical Reports Server (NTRS)

    Manning, C. R., Jr.; Lineback, L. D.

    1974-01-01

    Aspects of the development and characteristics of thermal shock resistant hafnia ceramic material for use in solid propellant rocket nozzles are presented. The investigation of thermal shock resistance factors for hafnia based composites, and the preparation and analysis of a model of elastic materials containing more than one crack are reported.

  9. Rocket exhaust plume computer program improvement. Volume 1: Summary: Method of characteristics nozzle and plume programs

    NASA Technical Reports Server (NTRS)

    Ratliff, A. W.; Smith, S. D.; Penny, N. M.

    1972-01-01

    A summary is presented of the various documents that discuss and describe the computer programs and analysis techniques which are available for rocket nozzle and exhaust plume calculations. The basic method of characteristics program is discussed, along with such auxiliary programs as the plume impingement program, the plot program and the thermochemical properties program.

  10. Particle Impact Erosion. Volume 4. User’s Manual Erosion Prediction Procedure for Rocket Nozzle Expansion Region

    DTIC Science & Technology

    1983-05-01

    line across the nozzle for the method-of-characteristics ( MOC ) solution used for the supersonic flow in the nozzle exit cone. After the transonic...Erosion Prediction Procedure for Rockat Nozzle Expansion Region I May 1983 Authors: Aerojet Strategic Propulsion Company J. D. Mockenhaupt P. 0. Box...Rocket Nozzle 6. PERFORMING OG. REPORT NUMBER Expansion Region LAX-FR-01 7. AUTHOR(e) It. CONTRACT OR GRANT NUMBER(#) - J. D. Mockenhaupt and J. H

  11. MHD thrust vectoring of a rocket engine

    NASA Astrophysics Data System (ADS)

    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.

  12. Program For Optimization Of Nuclear Rocket Engines

    NASA Technical Reports Server (NTRS)

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

    1994-01-01

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

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

    NASA Astrophysics Data System (ADS)

    Jones, Daniel Ray

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

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

    NASA Technical Reports Server (NTRS)

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

    1991-01-01

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

  15. Investigation of Low-Reynolds-Number Rocket Nozzle Design Using PNS-Based Optimization Procedure

    NASA Technical Reports Server (NTRS)

    Hussaini, M. Moin; Korte, John J.

    1996-01-01

    An optimization approach to rocket nozzle design, based on computational fluid dynamics (CFD) methodology, is investigated for low-Reynolds-number cases. This study is undertaken to determine the benefits of this approach over those of classical design processes such as Rao's method. A CFD-based optimization procedure, using the parabolized Navier-Stokes (PNS) equations, is used to design conical and contoured axisymmetric nozzles. The advantage of this procedure is that it accounts for viscosity during the design process; other processes make an approximated boundary-layer correction after an inviscid design is created. Results showed significant improvement in the nozzle thrust coefficient over that of the baseline case; however, the unusual nozzle design necessitates further investigation of the accuracy of the PNS equations for modeling expanding flows with thick laminar boundary layers.

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

    NASA Technical Reports Server (NTRS)

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

    1991-01-01

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

  17. Rocket Engines Displayed for 1966 Inspection at Lewis Research Center

    NASA Image and Video Library

    1966-10-21

    An array of rocket engines displayed in the Propulsion Systems Laboratory for the 1966 Inspection held at the National Aeronautics and Space Administration (NASA) Lewis Research Center. Lewis engineers had been working on chemical, nuclear, and solid rocket engines throughout the 1960s. The engines on display are from left to right: two scale models of the Aerojet M-1, a Rocketdyne J-2, a Pratt and Whitney RL-10, and a Rocketdyne throttleable engine. Also on display are several ejector plates and nozzles. The Chemical Rocket Division resolved issues such as combustion instability and screech, and improved operation of cooling systems and turbopumps. The 1.5-million pound thrust M-1 engine was the largest hydrogen-fueled rocket engine ever created. It was a joint project between NASA Lewis and Aerojet-General. Although much larger in size, the M-1 used technology developed for the RL-10 and J-2. The M-1 program was cancelled in late 1965 due to budget cuts and the lack of a post-Apollo mission. The October 1966 Inspection was the culmination of almost a year of events held to mark the centers’ 25th anniversary. The three‐day Inspection, Lewis’ first since 1957, drew 2000 business, industry, and government executives and included an employee open house. The visitors witnessed presentations at the major facilities and viewed the Gemini VII spacecraft, a Centaur rocket, and other displays in the hangar. In addition, Lewis’ newest facility, the Zero Gravity Facility, was shown off for the first time.

  18. Combustion Processes in Hybrid Rocket Engines

    NASA Technical Reports Server (NTRS)

    Venkateswaran,S.; Merkle, C. L.

    1996-01-01

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

  19. Effect of Nozzle Nonlinearities upon Nonlinear Stability of Liquid Propellant Rocket Motors

    NASA Technical Reports Server (NTRS)

    Padmanabhan, M. S.; Powell, E. A.; Zinn, B. T.

    1975-01-01

    A three dimensional, nonlinear nozzle admittance relation is developed by solving the wave equation describing finite amplitude oscillatory flow inside the subsonic portion of a choked, slowly convergent axisymmetric nozzle. This nonlinear nozzle admittance relation is then used as a boundary condition in the analysis of nonlinear combustion instability in a cylindrical liquid rocket combustor. In both nozzle and chamber analyses solutions are obtained using the Galerkin method with a series expansion consisting of the first tangential, second tangential, and first radial modes. Using Crocco's time lag model to describe the distributed unsteady combustion process, combustion instability calculations are presented for different values of the following parameters: (1) time lag, (2) interaction index, (3) steady-state Mach number at the nozzle entrance, and (4) chamber length-to-diameter ratio. In each case, limit cycle pressure amplitudes and waveforms are shown for both linear and nonlinear nozzle admittance conditions. These results show that when the amplitudes of the second tangential and first radial modes are considerably smaller than the amplitude of the first tangential mode the inclusion of nozzle nonlinearities has no significant effect on the limiting amplitude and pressure waveforms.

  20. Space Shuttle Main Engine nozzle thermal protection system

    NASA Technical Reports Server (NTRS)

    Nordlund, R. M.

    1985-01-01

    Two of the three Space Shuttle Main Engine (SSME) nozzles are exposed to significant reentry aeroheating loads. To ensure reusability of the Nozzle Assembly, the nozzle primary structure must not exceed specific temperature limits. Due to the thermal, pressure, and dynamic flexing of the nozzle during a mission cycle, an appropriate insulating system must have significant flexibility. Recent missions have demonstrated nozzle reentry aeroheating rates and heat loads much higher than predictions, higher than the capability of the original insulating system. A new insulating system has been developed using similar materials in an aerodynamically 'smooth' shape to both reduce the incoming heating and increase radiation cooling.

  1. Space Shuttle Main Engine nozzle thermal protection system

    NASA Technical Reports Server (NTRS)

    Nordlund, R. M.

    1985-01-01

    Two of the three Space Shuttle Main Engine (SSME) nozzles are exposed to significant reentry aeroheating loads. To ensure reusability of the Nozzle Assembly, the nozzle primary structure must not exceed specific temperature limits. Due to the thermal, pressure, and dynamic flexing of the nozzle during a mission cycle, an appropriate insulating system must have significant flexibility. Recent missions have demonstrated nozzle reentry aeroheating rates and heat loads much higher than predictions, higher than the capability of the original insulating system. A new insulating system has been developed using similar materials in an aerodynamically 'smooth' shape to both reduce the incoming heating and increase radiation cooling.

  2. Conceptual Design for a Dual-Bell Rocket Nozzle System Using a NASA F-15 Airplane as the Flight Testbed

    NASA Technical Reports Server (NTRS)

    Jones, Daniel S.; Ruf, Joseph H.; Bui, Trong T.; Martinez, Martel; St. John, Clinton W.

    2014-01-01

    The dual-bell rocket nozzle was first proposed in 1949, offering a potential improvement in rocket nozzle performance over the conventional-bell nozzle. Despite the performance advantages that have been predicted, both analytically and through static test data, the dual-bell nozzle has still not been adequately tested in a relevant flight environment. In 2013 a proposal was constructed that offered a NASA F-15 airplane as the flight testbed, with the plan to operate a dual-bell rocket nozzle during captive-carried flight. If implemented, this capability will permit nozzle operation into an external flow field similar to that of a launch vehicle, and facilitate an improved understanding of dual-bell nozzle plume sensitivity to external flow-field effects. More importantly, this flight testbed can be utilized to help quantify the performance benefit with the dual-bell nozzle, as well as to advance its technology readiness level. Toward this ultimate goal, this paper provides plans for future flights to quantify the external flow field of the airplane near the nozzle experiment, as well as details on the conceptual design for the dual-bell nozzle cold-flow propellant feed system integration within the NASA F-15 Propulsion Flight Test Fixture. The current study shows that this concept of flight research is feasible, and could result in valuable flight data for the dual-bell nozzle.

  3. Conceptual Design for a Dual-Bell Rocket Nozzle System Using a NASA F-15 Airplane as the Flight Testbed

    NASA Technical Reports Server (NTRS)

    Jones, Daniel S.; Ruf, Joseph H.; Bui, Trong T.; Martinez, Martel; St. John, Clinton W.

    2014-01-01

    The dual-bell rocket nozzle was first proposed in 1949, offering a potential improvement in rocket nozzle performance over the conventional-bell nozzle. Despite the performance advantages that have been predicted, both analytically and through static test data, the dual-bell nozzle has still not been adequately tested in a relevant flight environment. In 2013 a proposal was constructed that offered a NASA F-15 airplane as the flight testbed, with the plan to operate a dual-bell rocket nozzle during captive-carried flight. If implemented, this capability will permit nozzle operation into an external flow field similar to that of a launch vehicle, and facilitate an improved understanding of dual-bell nozzle plume sensitivity to external flow-field effects. More importantly, this flight testbed can be utilized to help quantify the performance benefit with the dual-bell nozzle, as well as to advance its technology readiness level. This presentation provides highlights of a technical paper that outlines this ultimate goal, including plans for future flights to quantify the external flow field of the airplane near the nozzle experiment, as well as details on the conceptual design for the dual-bell nozzle cold-flow propellant feed system integration within the NASA F-15 Propulsion Flight Test Fixture. The current study shows that this concept of flight research is feasible, and could result in valuable flight data for the dual-bell nozzle.

  4. Conceptual Design for a Dual-Bell Rocket Nozzle System Using a NASA F-15 Airplane as the Flight Testbed

    NASA Technical Reports Server (NTRS)

    Jones, Daniel S.; Ruf, Joseph H.; Bui, Trong T.; Martinez, Martel; St. John, Clinton W.

    2014-01-01

    The dual-bell rocket nozzle was first proposed in 1949, offering a potential improvement in rocket nozzle performance over the conventional-bell nozzle. Despite the performance advantages that have been predicted, both analytically and through static test data, the dual-bell nozzle has still not been adequately tested in a relevant flight environment. In 2013 a proposal was constructed that offered a National Aeronautics and Space Administration (NASA) F-15 airplane as the flight testbed, with the plan to operate a dual-bell rocket nozzle during captive-carried flight. If implemented, this capability will permit nozzle operation into an external flow field similar to that of a launch vehicle, and facilitate an improved understanding of dual-bell nozzle plume sensitivity to external flow-field effects. More importantly, this flight testbed can be utilized to help quantify the performance benefit with the dual-bell nozzle, as well as to advance its technology readiness level. Toward this ultimate goal, this report provides plans for future flights to quantify the external flow field of the airplane near the nozzle experiment, as well as details on the conceptual design for the dual-bell nozzle cold-flow propellant feed system integration within the NASA F-15 Propulsion Flight Test Fixture. The current study shows that this concept of flight research is feasible, and could result in valuable flight data for the dual-bell nozzle.

  5. Air-breathing Rocket Engine Test

    NASA Technical Reports Server (NTRS)

    1999-01-01

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

  6. Nitrous Oxide/Paraffin Hybrid Rocket Engines

    NASA Technical Reports Server (NTRS)

    Zubrin, Robert; Snyder, Gary

    2010-01-01

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

  7. Lift-Off Acoustics Prediction of Clustered Rocket Engines in the Near Field

    NASA Technical Reports Server (NTRS)

    Vu, Bruce; Plotkin, Ken

    2010-01-01

    This slide presentation presents a method of predicting acoustics during lift-off of the clustered rocket engines in the near field. Included is a definition of the near field, and the use of deflectors and shielding. There is discussion about the use of PAD, a software system designed to calculate the acoustic levels from the lift of of clustered rocket enginee, including updates to extend the calculation to directivity, water suppression, and clustered nozzles.

  8. Devising rocket power for smaller engines

    SciTech Connect

    Burruss, R.

    1996-04-01

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

  9. Nozzle erosion characterization and minimization for high-pressure rocket motor applications

    NASA Astrophysics Data System (ADS)

    Evans, Brian

    Understanding of the processes that cause nozzle throat erosion and developing methods for mitigation of erosion rate can allow higher operating pressures for advanced rocket motors. However, erosion of the nozzle throat region, which is a strong function of operating pressure, must be controlled to realize the performance gains of higher operating pressures. The objective of this work was the study the nozzle erosion rates at a broad range of pressures from 7 to 34.5 MPa (1,000 to 5,000 psia) using two different rocket motors. The first is an instrumented solidpropellant motor (ISPM), which uses two baseline solid propellants; one is a non-metallized propellant called Propellant S and the other is a metallized propellant called Propellant M. The second test rig is a non-metallized solid-propellant rocket motor simulator (RMS). The RMS is a gas rocket with the ability to vary the combustion-product species composition by systematically varying the flow rates of gaseous reactants. Several reactant mixtures were utilized in the study to determine the relative importance of different oxidizing species (such as H2O, OH, and CO2). Both test rigs are equipped with a windowed nozzle section for real-time X-ray radiography diagnostics of the instantaneous throat variations for deducing the instantaneous erosion rates. The nozzle test section for both motors can also incorporate a nozzle boundary-layer control system (NBLCS) as a means of nozzle erosion mitigation. The effectiveness of the NBLCS at preventing nozzle throat erosion was demonstrated for both the RMS and the ISPM motors at chamber pressures up to 34 MPa (4930 psia). All tests conducted with the NBLCS showed signs of coning of the propellant surface, leading to increased mass burning rate and resultant chamber pressure. Two correlations were developed for the nozzle erosion rates from solid propellant testing, one for metallized propellant and one for non-metallized propellants. The non-metallized propellant

  10. Electrodynamic actuators for rocket engine valves

    NASA Technical Reports Server (NTRS)

    Fiet, O.; Doshi, D.

    1972-01-01

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

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

    NASA Technical Reports Server (NTRS)

    Morris, Christopher I.

    2002-01-01

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

  12. Plasma Diagnostics Development for Advanced Rocket Engines

    NASA Astrophysics Data System (ADS)

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

    2000-10-01

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

  13. Temperature Dependent Modal Test/Analysis Correlation of X-34 Fastrac Composite Rocket Nozzle

    NASA Technical Reports Server (NTRS)

    Brown, Andrew M.; Brunty, Joseph A. (Technical Monitor)

    2001-01-01

    A unique high temperature modal test and model correlation/update program has been performed on the composite nozzle of the FASTRAC engine for the NASA X-34 Reusable Launch Vehicle. The program was required to provide an accurate high temperature model of the nozzle for incorporation into the engine system structural dynamics model for loads calculation; this model is significantly different from the ambient case due to the large decrease in composite stiffness properties due to heating. The high-temperature modal test was performed during a hot-fire test of the nozzle. Previously, a series of high fidelity modal tests and finite element model correlation of the nozzle in a free-free configuration had been performed. This model was then attached to a modal-test verified model of the engine hot-fire test stand and the ambient system mode shapes were identified. A reduced set of accelerometers was then attached to the nozzle, the engine fired full-duration, and the frequency peaks corresponding to the ambient nozzle modes individually isolated and tracked as they decreased during the test. To update the finite-element model of the nozzle to these frequency curves, the percentage differences of the anisotropic composite moduli due to temperature variation from ambient, which had been used in the initial modeling and which were obtained by small sample coupon testing, were multiplied by an iteratively determined constant factor. These new properties were used to create high-temperature nozzle models corresponding to 10 second engine operation increments and tied into the engine system model for loads determination.

  14. Plasma torch testing for thermostructural evaluation of rocket motor nozzle materials

    NASA Technical Reports Server (NTRS)

    Prince, Andrew S.; Bunker, Robert C.; Lawrence, Tim

    1989-01-01

    This paper presents data from the thermostructural testing of tape-wrapped carbon phenolic. This work has been performed with the use of a plasma torch and loading device in an effort to study the anomalous erosion characteristicfs of that seen in the Space Shuttle Solid Rocket Motor Nozzle STS-8A. Testing is conducted in an effort to determine conditions or parameters involved in this mode of failure.

  15. Plasma torch testing for thermostructural evaluation of rocket motor nozzle materials

    NASA Technical Reports Server (NTRS)

    Prince, Andrew S.; Bunker, Robert C.; Lawrence, Tim

    1989-01-01

    This paper presents data from the thermostructural testing of tape-wrapped carbon phenolic. This work has been performed with the use of a plasma torch and loading device in an effort to study the anomalous erosion characteristicfs of that seen in the Space Shuttle Solid Rocket Motor Nozzle STS-8A. Testing is conducted in an effort to determine conditions or parameters involved in this mode of failure.

  16. Pattern classification approach to rocket engine diagnostics

    SciTech Connect

    Tulpule, S.

    1989-01-01

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

  17. Ceramic composites for rocket engine turbines

    NASA Technical Reports Server (NTRS)

    Herbell, Thomas P.; Eckel, Andrew J.

    1991-01-01

    The use of ceramic materials in the hot section of the fuel turbopump of advanced reusable rocket engines promises increased performance and payload capability, improved component life and economics, and greater design flexibility. Severe thermal transients present during operation of the Space Shuttle Main Engine (SSME), push metallic components to the limit of their capabilities. Future engine requirements might be even more severe. In phase one of this two-phase program, performance benefits were quantified and continuous fiber reinforced ceramic matrix composite components demonstrated a potential to survive the hostile environment of an advaced rocket engine turbopump.

  18. Ceramic composites for rocket engine turbines

    NASA Technical Reports Server (NTRS)

    Herbell, Thomas P.; Eckel, Andrew J.

    1991-01-01

    The use of ceramic materials in the hot section of the fuel turbopump of advanced reusable rocket engines promises increased performance and payload capability, improved component life and economics, and greater design flexibility. Severe thermal transients present during operation of the Space Shuttle Main Engine (SSME), push metallic components to the limit of their capabilities. Future engine requirements might be even more severe. In phase one of this two-phase program, performance benefits were quantified and continuous fiber reinforced ceramic matrix composite components demonstrated a potential to survive the hostile environment of an advanced rocket engine turbopump.

  19. End-effects-regime in full scale and lab scale rocket nozzles

    NASA Astrophysics Data System (ADS)

    Rojo, Raymundo; Tinney, Charles; Baars, Woutijn; Ruf, Joseph

    2014-11-01

    Modern rockets utilize a thrust-optimized parabolic-contour design for their nozzles for its high performance and reliability. However, the evolving internal flow structures within these high area ratio rocket nozzles during start up generate a powerful amount of vibro-acoustic loads that act on the launch vehicle. Modern rockets must be designed to accommodate for these heavy loads or else risk a catastrophic failure. This study quantifies a particular moment referred to as the ``end-effects regime,'' or the largest source of vibro-acoustic loading during start-up [Nave & Coffey, AIAA Paper 1973-1284]. Measurements from full scale ignitions are compared with aerodynamically scaled representations in a fully anechoic chamber. Laboratory scale data is then matched with both static and dynamic wall pressure measurements to capture the associating shock structures within the nozzle. The event generated during the ``end-effects regime'' was successfully reproduced in the both the lab-scale models, and was characterized in terms of its mean, variance and skewness, as well as the spectral properties of the signal obtained by way of time-frequency analyses.

  20. Advanced oxygen-hydrocarbon rocket engine study

    NASA Technical Reports Server (NTRS)

    Obrien, J. O.; Ewen, R. L.; Kent, S.; Meagher, G. M.

    1980-01-01

    The program consists of parametric analysis and design to provide a consistent engine system data base for defining advantages and disadvantages, system performance and operating limits, engine parametric data, and technology requirements for candidate high pressure LO2/Hydrocarbon engine systems. The parametric chamber and nozzle cooling analysis was completed for the four potential coolants: RP-1, LCH4, LO2, and LH2. A summary of the cooling capability of each propellant is presented.

  1. Preliminary Studies of a Pulsed Detonation Rocket Engine

    NASA Technical Reports Server (NTRS)

    Cambier, Jean-Luc; Adelman, H. G.; Menees, G. P.; Edwards, Thomas A. (Technical Monitor)

    1995-01-01

    In the new era of space exploration, there is a strong need for more efficient, cheaper and more reliable propulsion devices. With dramatic increase in specific impulse, the overall mass of fuel to be lifted into orbit is decreased, and this leads, in turn, to much lower mass requirements at lift-off, higher payload ratios and lower launch costs. The Pulsed Detonation engine (PDE) has received much attention lately due to its unique combination of simplicity, light-weight and efficiency. Current investigations focus principally on its use as a low speed, airbreathing engine, although other applications have also been proposed. Its use as a rocket propulsion device was first proposed in 1988 by the present authors. The superior efficiency of the Pulsed Detonation Rocket Engine (PDRE) is due to the near constant volume combustion process of a detonation wave. Our preliminary estimates suggest that the PDRE is theoretically capable of achieving specific impulses as high as 720 sec, a dramatic improvement over the current 480 sec of conventional rocket engines, making it competitive with nuclear thermal rockets. In addition to this remarkable efficiency, the PDRE may eliminate the need for high pressure cryogenic turbopumps, a principal source of failures. The heat transfer rates are also much lower, eliminating the need for nozzle cooling. Overall, the engine is more reliable and has a much lower weight. This paper will describe in detail the operation of the PDRE and calculate its performance, through numerical simulations. Engineering issues will be addressed and discussed, and the impact on mission profiles will also be presented. Finally, the performance of the PDRE using in-situ resources, such as CO and O2 from the martian atmosphere, will also be computed.

  2. Preliminary Studies of a Pulsed Detonation Rocket Engine

    NASA Technical Reports Server (NTRS)

    Cambier, Jean-Luc; Adelman, H. G.; Menees, G. P.; Edwards, Thomas A. (Technical Monitor)

    1995-01-01

    In the new era of space exploration, there is a strong need for more efficient, cheaper and more reliable propulsion devices. With dramatic increase in specific impulse, the overall mass of fuel to be lifted into orbit is decreased, and this leads, in turn, to much lower mass requirements at lift-off, higher payload ratios and lower launch costs. The Pulsed Detonation engine (PDE) has received much attention lately due to its unique combination of simplicity, light-weight and efficiency. Current investigations focus principally on its use as a low speed, airbreathing engine, although other applications have also been proposed. Its use as a rocket propulsion device was first proposed in 1988 by the present authors. The superior efficiency of the Pulsed Detonation Rocket Engine (PDRE) is due to the near constant volume combustion process of a detonation wave. Our preliminary estimates suggest that the PDRE is theoretically capable of achieving specific impulses as high as 720 sec, a dramatic improvement over the current 480 sec of conventional rocket engines, making it competitive with nuclear thermal rockets. In addition to this remarkable efficiency, the PDRE may eliminate the need for high pressure cryogenic turbopumps, a principal source of failures. The heat transfer rates are also much lower, eliminating the need for nozzle cooling. Overall, the engine is more reliable and has a much lower weight. This paper will describe in detail the operation of the PDRE and calculate its performance, through numerical simulations. Engineering issues will be addressed and discussed, and the impact on mission profiles will also be presented. Finally, the performance of the PDRE using in-situ resources, such as CO and O2 from the martian atmosphere, will also be computed.

  3. Air-Breathing Rocket Engine Test

    NASA Technical Reports Server (NTRS)

    2000-01-01

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

  4. Air-Breathing Rocket Engine Test

    NASA Technical Reports Server (NTRS)

    2000-01-01

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

  5. Application of Optical Measurement Techniques During Fabrication and Testing of Liquid Rocket Nozzles

    NASA Technical Reports Server (NTRS)

    Gradl, Paul R.

    2015-01-01

    This paper presents a series of optical measurement techniques that were developed for use during large-scale fabrication and testing of nozzle components. A thorough understanding of hardware throughout the fabrication cycle and hotfire testing is critical to meet component design intent. Regeneratively cooled nozzles and associated tooling require tight control of tolerances during the fabrication process to ensure optimal performance. Additionally, changes in geometry during testing can affect performance of the nozzle and mating components. Structured light scanning and digital image correlation techniques were used to collect data during the fabrication and test of nozzles, in addition to other engine components. This data was used to analyze deformations data during machining, heat treatment, assembly and testing operations. A series of feasibility experiments were conducted for these techniques that led to use on full scale nozzles during the J-2X upper stage engine program in addition to other engine development programs. This paper discusses the methods and results of these measurement techniques throughout the nozzle life cycle and application to other components.

  6. The Strutjet Rocket Based Combined Cycle Engine

    NASA Technical Reports Server (NTRS)

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

    1998-01-01

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

  7. Reliability of solid rocket motor cases and nozzles

    NASA Technical Reports Server (NTRS)

    Crose, James G.

    1993-01-01

    A recent article in Aerospace America claims that 'the average success ratio of the current U.S. stable of launch vehicles, including upper stages, is about 92 percent (without upper stages it is close to 95 percent). The 8 percent failure probability implies an expected loss of $12M per flight, not including the lost opportunity costs'. Since payload costs are likely to be much greater than launch costs and even more so for the new launch vehicles for the Advanced Launch Development Program (ALDP), the cost of rocket motor unreliability at the current 8 percent rate can run into billions of dollars if expected increases in demand are realized. At an 8 percent failure rate, it is extremely unlikely that failure will occur during the first few ground tests of a new system. At that time, most of the design, analysis and tooling costs of the program have not expended. Since most systems are expected to be used ten to a hundred or more times, the likelihood of one or more failures is very large, and it can be expected that the above losses will be realized in the future. This will occur unless the problems are addressed and remedied. Recent trends suggest the problem is not being addressed adequately.

  8. Reliability of solid rocket motor cases and nozzles

    NASA Astrophysics Data System (ADS)

    Crose, James G.

    1993-02-01

    A recent article in Aerospace America claims that 'the average success ratio of the current U.S. stable of launch vehicles, including upper stages, is about 92 percent (without upper stages it is close to 95 percent). The 8 percent failure probability implies an expected loss of $12M per flight, not including the lost opportunity costs'. Since payload costs are likely to be much greater than launch costs and even more so for the new launch vehicles for the Advanced Launch Development Program (ALDP), the cost of rocket motor unreliability at the current 8 percent rate can run into billions of dollars if expected increases in demand are realized. At an 8 percent failure rate, it is extremely unlikely that failure will occur during the first few ground tests of a new system. At that time, most of the design, analysis and tooling costs of the program have not expended. Since most systems are expected to be used ten to a hundred or more times, the likelihood of one or more failures is very large, and it can be expected that the above losses will be realized in the future. This will occur unless the problems are addressed and remedied. Recent trends suggest the problem is not being addressed adequately.

  9. Numerical simulation of shock-induced separated flows in overexpanded rocket nozzles

    NASA Astrophysics Data System (ADS)

    Shams, A.; Girard, S.; Comte, P.

    2012-01-01

    Flow separation in rocket nozzles is undesirable because of its unsteady and nonsymmetric nature, which leads to dangerous side-loads. At the initial stages of start-up, when a thrust optimized contour (TOC) nozzle operates under overexpanded conditions, free shock separation (FSS) takes place. Under certain conditions, this free separated flow reattaches back to the nozzle wall and forms restricted shock separation (RSS). The appearance of restricted shock separated flow depends upon the nozzle contour in a well-defined range of nozzle pressure ratios (NPR) and is characterized by a cap-shock pattern. The flow transition process from FSS to RSS flow configurations is a complex phenomenon and has been an area of interest for a few decades now. In the present study, an attempt has been made to understand the formation of the cap-shock pattern and the RSS flow configuration in a thrust optimized contour (TOC) nozzle. The presented research work consists of two parts. In the first part of the paper, numerical investigation of flow transition (FSS→RSS) has been performed to understand the formation of the cap-shock pattern, which is believed to be the main cause for this flow transition from FSS to RSS. Axisymmetric numerical calculations on a wide range of NPRs (15 25) are performed to reproduce the forward transition process and are found to be in good agreement with the experiments. In the second part, some light has been shed on various aspects of RSS flow regime. Three-dimensional (3D) numerical simulations have been performed on a wide range of NPRs, i. e., 25.0, 30.0, 38.0, 41.0, and 46.0. Detailed analysis of these numerical results allows examining the evolution of the separation point and the cap-shock pattern with respect to the NPR. Furthermore, some insights based on the axial momentum along the nozzle axis and radial momentum distributions across the quadruple point are given.

  10. AJ26 rocket engine testing news briefing

    NASA Image and Video Library

    2010-02-24

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

  11. Analyzing Thermal Conditions In Rocket Engines

    NASA Technical Reports Server (NTRS)

    Naraghi, M. H. N.

    1994-01-01

    Computer code, RTE, developed to perform three-dimensional thermal analyses of rocket thrust chambers. Calculates rate of heat transfer from combustion gases to coolant, coolant-temperature rise and pressure drop, and temperature profiles within cooling-jacket wall. Also calculates combustion-gas wall static pressure, temperature and enthalpy, as well as coolant pressure, temperature, and mach number for all stations. Program used for any propellant combination and most coolants commonly used in rockets. Code used for both regeneratively and radiatively cooled engines. However, in case of regeneratively cooled engines, applicability limited to engines featuring single-pass cooling and rectangular cooling channels.

  12. Composite Material Application to Liquid Rocket Engines

    NASA Technical Reports Server (NTRS)

    Judd, D. C.

    1982-01-01

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

  13. AJ26 rocket engine testing news briefing

    NASA Technical Reports Server (NTRS)

    2010-01-01

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

  14. Iridium/Rhenium Parts For Rocket Engines

    NASA Technical Reports Server (NTRS)

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

    1991-01-01

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

  15. Iridium/Rhenium Parts For Rocket Engines

    NASA Technical Reports Server (NTRS)

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

    1991-01-01

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

  16. A Versatile Rocket Engine Hot Gas Facility

    NASA Technical Reports Server (NTRS)

    Green, James M.

    1993-01-01

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

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

    NASA Technical Reports Server (NTRS)

    Morris, Christopher I.

    2003-01-01

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

  18. Additive Manufacturing for Affordable Rocket Engines

    NASA Technical Reports Server (NTRS)

    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. Design considerations in clustering nuclear rocket engines

    NASA Technical Reports Server (NTRS)

    Sager, Paul H.

    1992-01-01

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

  20. Design considerations in clustering nuclear rocket engines

    NASA Technical Reports Server (NTRS)

    Sager, Paul H.

    1992-01-01

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

  1. Flow processes in overexpanded chemical rocket nozzles. Part 3: Methods for the aimed flow separation and side load reduction

    NASA Technical Reports Server (NTRS)

    Schmucker, R. H.

    1983-01-01

    Methods aimed at reduction of overexpansion and side load resulting from asymmetric flow separation for rocket nozzles with a high opening ratio are described. The methods employ additional measures for nozzles with a fixed opening ratio. The flow separation can be controlled by several types of nozzle inserts, the properties of which are discussed. Side loads and overexpansion can be reduced by adapting the shape of the nozzle and taking other additional measures for controlled separation of the boundary layer, such as trip wires.

  2. Outbrief - Long Life Rocket Engine Panel

    NASA Technical Reports Server (NTRS)

    Quinn, Jason Eugene

    2004-01-01

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

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

    NASA Technical Reports Server (NTRS)

    Chung, Chan H.; Kim, Suk C.; Stubbs, Robert M.; De Witt, Kenneth J.

    1993-01-01

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

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

    NASA Technical Reports Server (NTRS)

    Chung, Chan H.; Kim, Suk C.; Stubbs, Robert M.; De Witt, Kenneth J.

    1993-01-01

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

  5. Low-thrust chemical rocket engine study

    NASA Technical Reports Server (NTRS)

    Mellish, J. A.

    1981-01-01

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

  6. Boiler and Pressure Balls Monopropellant Thermal Rocket Engine

    NASA Technical Reports Server (NTRS)

    Greene, William D. (Inventor)

    2009-01-01

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

  7. Fiber-Reinforced Superalloys For Rocket Engines

    NASA Technical Reports Server (NTRS)

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

    1990-01-01

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

  8. Vapor Grown Carbon Fiber/Phenolic Matrix Composites for Rocket Nozzles and Heat Shields

    NASA Technical Reports Server (NTRS)

    Patton, R. D.; Pittman, C. U., Jr.; Wang, L.; Day, A.; Hill, J. R.

    2001-01-01

    The ablation and mechanical and thermal properties of vapor grown carbon fiber (VGCF)/phenolic resin composites were evaluated to determine the potential of using this material in solid rocket motor nozzles. Composite specimens with varying VGCF loading (30%-50% wt) including one sample with ex-rayon carbon fiber plies were prepared and exposed to a plasma torch for 20 s with a heat flux of 16.5 MW/sq m at approximately 1650 C. Low erosion rates and little char formation were observed, confirming that these materials were promising for rocket motor nozzle materials. When fiber loadings increased, mechanical properties and ablative properties improved. The VGCF composites had low thermal conductivities (approximately 0.56 W/m-C) indicating they were good insulating materials. If a 65% fiber loading in VGCF composite can be achieved, then ablative properties are projected to be comparable to or better than the composite material currently used on the Space Shuttle Reusable Solid Rocket Motor (RSRM).

  9. Variable Specific Impulse Magnetoplasma Rocket Engine

    NASA Technical Reports Server (NTRS)

    Chang-Diaz, Franklin R. (Inventor)

    2002-01-01

    An engine is disclosed, including a controllable output plasma generator, a controllable heater for selectably raising a temperature of the plasma connected to an outlet of the plasma generator, and a nozzle connected to an outlet of the heater, through which heated plasma is discharged to provide thrust. In one embodiment, the source of plasma is a helicon generator. In one embodiment, the heater is an ion cyclotron resonator. In one embodiment, the nozzle is a radially diverging magnetic field disposed on a discharge side of the heater so that helically travelling particles in the beater exit the heater at high axial velocity. A particular embodiment includes control circuits for selectably directing a portion of radio frequency power from an RF generator to the helicon generator and to the cyclotron resonator so that the thrust output and the specific impulse of the engine can be selectively controlled. A method of propelling a vehicle is also disclosed. The method includes generating a plasma, heating said plasma, and discharging the heated plasma through a nozzle. In one embodiment, the nozzle is a diverging magnetic field. In this embodiment, the heating is performed by applying a radio frequency electro magnetic field to the plasma at the ion cyclotron frequency in an axially polarized DC magnetic field.

  10. Variable Specific Impulse Magnetoplasma Rocket Engine

    NASA Technical Reports Server (NTRS)

    Chang-Diaz, Franklin R. (Inventor)

    2002-01-01

    An engine is disclosed, including a controllable output plasma generator, a controllable heater for selectably raising a temperature of the plasma connected to an outlet of the plasma generator, and a nozzle connected to an outlet of the heater, through which heated plasma is discharged to provide thrust. In one embodiment, the source of plasma is a helicon generator. In one embodiment, the heater is an ion cyclotron resonator. In one embodiment, the nozzle is a radially diverging magnetic field disposed on a discharge side of the heater so that helically travelling particles in the beater exit the heater at high axial velocity. A particular embodiment includes control circuits for selectably directing a portion of radio frequency power from an RF generator to the helicon generator and to the cyclotron resonator so that the thrust output and the specific impulse of the engine can be selectively controlled. A method of propelling a vehicle is also disclosed. The method includes generating a plasma, heating said plasma, and discharging the heated plasma through a nozzle. In one embodiment, the nozzle is a diverging magnetic field. In this embodiment, the heating is performed by applying a radio frequency electro magnetic field to the plasma at the ion cyclotron frequency in an axially polarized DC magnetic field.

  11. Incident shock strength evolution in overexpanded jet flow out of rocket nozzle

    NASA Astrophysics Data System (ADS)

    Silnikov, Mikhail V.; Chernyshov, Mikhail V.

    2017-06-01

    The evolution of the incident shock in the plane overexpanded jet flow or in the axisymmetric one is analyzed theoretically and compared at the whole range of governing flow parameters. Analytical results can be applied to avoid jet flow instability and self-oscillation effects at rocket launch, to improve launch safety and to suppress shock-wave induced noise harmful to environment and personnel. The mathematical model of ;differential conditions of dynamic compatibility; was applied to the curved shock in non-uniform plane or axisymmetrical flow. It allowed us to study such features of the curved incident shock and flow downstream it as shock geometrical curvature, jet boundary curvature, local increase or decrease of the shock strength, flow vorticity rate (local pressure gradient) in the vicinity of the nozzle lip, static pressure gradient in the compressed layer downstream the shock, and many others. All these quantities sufficiently depend on the flow parameters (flow Mach number, jet overexpansion rate, nozzle throat angle, and ration of gas specific heats). These dependencies are sometimes unusual, especially at small Mach numbers. It was also surprising that there is no great difference among all these flowfield features in the plane jet and in the axisymmetrical jet flow out of a nozzle with large throat angle, but all these parameters behave in a quite different way in an axisymmetrical jet at small and moderate nozzle throat angles.

  12. Experimental analysis of SiC-based refractory concrete in hybrid rocket nozzles

    NASA Astrophysics Data System (ADS)

    D'Elia, Raffaele; Bernhart, Gérard; Hijlkema, Jouke; Cutard, Thierry

    2016-09-01

    Hybrid propulsion represents a good alternative to the more widely used liquid and solid systems. This technology combines some important specifications of the latters, as the possibility of re-ignition, thrust modulation, a higher specific impulse than solid systems, a greater simplicity and a lower cost than liquid systems. Nevertheless the highly oxidizing environment represents a major problem as regards the thermo-oxidation and ablative behavior of nozzle materials. The main goal of this research is to characterize a silicon carbide based micro-concrete with a maximum aggregates size of 800 μm, in a hybrid propulsion environment. The nozzle throat has to resist to a highly oxidizing polyethylene/nitrous oxide hybrid environment, under temperatures up to 2900 K. Three tests were performed on concrete-based nozzles in HERA Hybrid Rocket Motor (HRM) test bench at ONERA. Pressure chamber evolution and observations before and after tests are used to investigate the ablated surface at nozzle throat. Ablation behavior and crack generation are discussed and some improvements are proposed.

  13. Shape memory alloy actuated adaptive exhaust nozzle for jet engine

    NASA Technical Reports Server (NTRS)

    Song, Gangbing (Inventor); Ma, Ning (Inventor)

    2009-01-01

    The proposed adaptive exhaust nozzle features an innovative use of the shape memory alloy (SMA) actuators for actively control of the opening area of the exhaust nozzle for jet engines. The SMA actuators remotely control the opening area of the exhaust nozzle through a set of mechanism. An important advantage of using SMA actuators is the reduction of weight of the actuator system for variable area exhaust nozzle. Another advantage is that the SMA actuator can be activated using the heat from the exhaust and eliminate the need of other energy source. A prototype has been designed and fabricated. The functionality of the proposed SMA actuated adaptive exhaust nozzle is verified in the open-loop tests.

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

    NASA Technical Reports Server (NTRS)

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

    2002-01-01

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

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

    NASA Technical Reports Server (NTRS)

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

    2002-01-01

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

  16. AJ26 Rocket Engine Test

    NASA Image and Video Library

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

  17. 2011-2012 Dryden Center Innovation Fund End of the Year Report: Altitude-Compensating Rocket Nozzles

    NASA Technical Reports Server (NTRS)

    Jones, Daniel S.; Bui, Trong T.

    2012-01-01

    This report highlights one of the many successful projects at the NASA Dryden Flight Research Center that was approved for FY12 funding under the Center Innovation Fund. This project was focused on advancing the technology readiness level of one specific type of altitude-compensating nozzle: the dual-bell rocket nozzle. When considering a rocket's performance over its entire integrated trajectory, the dual-bell nozzle has been predicted to achieve a higher total impulse over the conventional bell nozzle, which is expected to result in a greater capability of payload mass to low-Earth orbit. Although the dual-bell rocket nozzle has been thoroughly studied for several decades, this nozzle has still not been adequately tested in a relevant flight-like environment. This report provides highlights and top-level details on the FY12 feasibility effort to advance this promising technology through flight test, a collaborative effort which leverages NASA Marshall's dual-bell nozzle research and development with Dryden's expertise in propulsion-focused flight testing. To accomplish this goal, the NASA F-15B is proposed as the testbed for the initial flight-test campaign to advance this greatly needed capability.

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

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

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

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

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

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

  20. Hybrid rocket engine, theoretical model and experiment

    NASA Astrophysics Data System (ADS)

    Chelaru, Teodor-Viorel; Mingireanu, Florin

    2011-06-01

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

  1. Low-thrust chemical rocket engine study

    NASA Technical Reports Server (NTRS)

    Shoji, J. M.

    1981-01-01

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

  2. Flow processes in overexpanded chemical rocket nozzles. Part 1: Flow separation

    NASA Technical Reports Server (NTRS)

    Schmucker, R. H.

    1973-01-01

    An investigation was made of published nozzle flow separation data in order to determine the parameters which affect the separation condition. A comparison of experimental data with empirical and theoretical separation prediction methods leads to the selection of suitable equations for the separation criterion. The results were used to predict flow separation of the main space shuttle engine.

  3. Flow processes in overexpanded chemical rocket nozzles. Part 1: Flow separation

    NASA Technical Reports Server (NTRS)

    Schmucker, R. H.

    1984-01-01

    An investigation was made of published nozzle flow separation data in order to determine the parameters which affect the separation conditions. A comparison of experimental data with empirical and theoretical separation prediction methods leads to the selection of suitable equations for the separation criterion. The results were used to predict flow separation of the main space shuttle engine.

  4. Technicians Manufacture a Nozzle for the Kiwi B-1-B Engine

    NASA Image and Video Library

    1964-05-21

    Technicians manufacture a nozzle for the Kiwi B-1-B nuclear rocket engine in the Fabrication Shop’s vacuum oven at the National Aeronautics and Space Administration (NASA) Lewis Research Center. The Nuclear Engine for Rocket Vehicle Applications (NERVA) was a joint NASA and Atomic Energy Commission (AEC) endeavor to develop a nuclear-powered rocket for both long-range missions to Mars and as a possible upper-stage for the Apollo Program. The early portion of the program consisted of basic reactor and fuel system research. This was followed by a series of Kiwi reactors built to test basic nuclear rocket principles in a non-flying nuclear engine. The next phase, NERVA, would create an entire flyable engine. The final phase of the program, called Reactor-In-Flight-Test, would be an actual launch test. The AEC was responsible for designing the nuclear reactor and overall engine. NASA Lewis was responsible for developing the liquid-hydrogen fuel system. The turbopump, which pumped the fuels from the storage tanks to the engine, was the primary tool for restarting the engine. The NERVA had to be able to restart in space on its own using a safe preprogrammed startup system. Lewis researchers endeavored to design and test this system. This non-nuclear Kiwi engine, seen here, was being prepared for tests at Lewis’ High Energy Rocket Engine Research Facility (B-1) located at Plum Brook Station. The tests were designed to start an unfueled Kiwi B-1-B reactor and its Aerojet Mark IX turbopump without any external power.

  5. Advanced oxygen-hydrocarbon rocket engine study

    NASA Technical Reports Server (NTRS)

    Obrien, C. J.; Salkeid, R.; Mueggenburg, H.; Ewen, R. L.

    1980-01-01

    Preliminary identification and evaluation of promising LO2/Hydrocarbon rocket engine cycles were used to produce a consistent and reliable data base for vehicle optimization and design studies. cycles G and C were chosen for design analysis. Preliminary design analysis of the heat transfer subsystem was performed to establish major technology requirements.

  6. Advanced oxygen-hydrocarbon rocket engine study

    NASA Technical Reports Server (NTRS)

    Obrien, C. J.

    1980-01-01

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

  7. Turbopump systems for liquid rocket engines

    NASA Technical Reports Server (NTRS)

    1974-01-01

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

  8. Spray Diagnostics in Rocket Engines Using Phase Doppler Analyzer

    NASA Technical Reports Server (NTRS)

    Isaac, Kakkattukuzhy M.

    1996-01-01

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

  9. Thermal stratification potential in rocket engine coolant channels

    NASA Technical Reports Server (NTRS)

    Kacynski, Kenneth J.

    1992-01-01

    The potential for rocket engine coolant channel flow stratification was computationally studied. A conjugate, 3-D, conduction/advection analysis code (SINDA/FLUINT) was used. Core fluid temperatures were predicted to vary by over 360 K across the coolant channel, at the throat section, indicating that the conventional assumption of a fully mixed fluid may be extremely inaccurate. Because of the thermal stratification of the fluid, the walls exposed to the rocket engine exhaust gases will be hotter than an assumption of full mixing would imply. In this analysis, wall temperatures were 160 K hotter in the turbulent mixing case than in the full mixing case. The discrepancy between the full mixing and turbulent mixing analyses increased with increasing heat transfer. Both analysis methods predicted identical channel resistances at the coolant inlet, but in the stratified analysis the thermal resistance was negligible. The implications are significant. Neglect of thermal stratification could lead to underpredictions in nozzle wall temperatures. Even worse, testing at subscale conditions may be inadequate for modeling conditions that would exist in a full scale engine.

  10. Fuel injector nozzle for an internal combustion engine

    DOEpatents

    Cavanagh, Mark S.; Urven, Jr., Roger L.; Lawrence, Keith E.

    2011-03-22

    A direct injection fuel injector includes a nozzle tip having a plurality of passages allowing fluid communication between an inner nozzle tip surface portion and an outer nozzle tip surface portion and directly into a combustion chamber of an internal combustion engine. A first group of the passages have inner surface apertures located substantially in a first common plane. A second group of the passages have inner surface apertures located substantially in at least a second common plane substantially parallel to the first common plane. The second group has more passages than the first group.

  11. Fuel Injector Nozzle For An Internal Combustion Engine

    DOEpatents

    Cavanagh, Mark S.; Urven, Jr.; Roger L.; Lawrence, Keith E.

    2006-04-25

    A direct injection fuel injector includes a nozzle tip having a plurality of passages allowing fluid communication between an inner nozzle tip surface portion and an outer nozzle tip surface portion and directly into a combustion chamber of an internal combustion engine. A first group of the passages have inner surface apertures located substantially in a first common plane. A second group of the passages have inner surface apertures located substantially in at least a second common plane substantially parallel to the first common plane. The second group has more passages than the first group.

  12. Fuel injector nozzle for an internal combustion engine

    DOEpatents

    Cavanagh, Mark S.; Urven, Jr., Roger L.; Lawrence, Keith E.

    2008-11-04

    A direct injection fuel injector includes a nozzle tip having a plurality of passages allowing fluid communication between an inner nozzle tip surface portion and an outer nozzle tip surface portion and directly into a combustion chamber of an internal combustion engine. A first group of the passages have inner surface apertures located substantially in a first common plane. A second group of the passages have inner surface apertures located substantially in at least a second common plane substantially parallel to the first common plane. The second group has more passages than the first group.

  13. Fuel injector nozzle for an internal combustion engine

    DOEpatents

    Cavanagh, Mark S.; Urven, Jr., Roger L.; Lawrence, Keith E.

    2007-11-06

    A direct injection fuel injector includes a nozzle tip having a plurality of passages allowing fluid communication between an inner nozzle tip surface portion and an outer nozzle tip surface portion and directly into a combustion chamber of an internal combustion engine. A first group of the passages have inner surface apertures located substantially in a first common plane. A second group of the passages have inner surface apertures located substantially in at least a second common plane substantially parallel to the first common plane. The second group has more passages than the first group.

  14. Metal Matrix Composites for Rocket Engine Applications

    NASA Technical Reports Server (NTRS)

    McDonald, Kathleen R.; Wooten, John R.

    2000-01-01

    This document is from a presentation about the applications of Metal Matrix Composites (MMC) in rocket engines. Both NASA and the Air Force have goals which would reduce the costs and the weight of launching spacecraft. Charts show the engine weight distribution for both reuseable and expendable engine components. The presentation reviews the operating requirements for several components of the rocket engines. The next slide reviews the potential benefits of MMCs in general and in use as materials for Advanced Pressure Casting. The next slide reviews the drawbacks of MMCs. The reusable turbopump housing is selected to review for potential MMC application. The presentation reviews solutions for reusable turbopump materials, pointing out some of the issues. It also reviews the development of some of the materials.

  15. Fabry-Perot interferometer development for rocket engine plume spectroscopy

    NASA Technical Reports Server (NTRS)

    Bickford, R. L.; Madzsar, G.

    1990-01-01

    This paper describes a new rugged high-resolution Fabry-Perot interferometer (FPI) designed for rocket engine plume spectroscopy, which is capable of detecting spectral signatures of eroding engine components during rocket engine tests and/or flight operations. The FPI system will make it possible to predict and to respond to the incipient rocket engine failures and to indicate the presence of rocket components degradation. The design diagram of the FPI spectrometer is presented.

  16. Jet-Engine Exhaust Nozzle With Thrust-Directing Flaps

    NASA Technical Reports Server (NTRS)

    Wing, David J.

    1996-01-01

    Convergent/divergent jet-engine exhaust nozzle has cruciform divergent passage containing flaps that move to deflect flow of exhaust in either or both planes perpendicular to main fore-and-aft axis of undeflected flow. Prototype of thrust-vector-control nozzles installed in advanced, high-performance airplanes to provide large pitching (usually, vertical) and yawing (usually, horizontal) attitude-control forces independent of attitude-control forces produced by usual aerodynamic control surfaces.

  17. Jet-Engine Exhaust Nozzle With Thrust-Directing Flaps

    NASA Technical Reports Server (NTRS)

    Wing, David J.

    1996-01-01

    Convergent/divergent jet-engine exhaust nozzle has cruciform divergent passage containing flaps that move to deflect flow of exhaust in either or both planes perpendicular to main fore-and-aft axis of undeflected flow. Prototype of thrust-vector-control nozzles installed in advanced, high-performance airplanes to provide large pitching (usually, vertical) and yawing (usually, horizontal) attitude-control forces independent of attitude-control forces produced by usual aerodynamic control surfaces.

  18. Rocket nozzle expansion ratio analysis for dual-fuel earth-to-orbit vehicles

    NASA Technical Reports Server (NTRS)

    Martin, James A.

    1989-01-01

    Results are reported from a recent study of the effects of Space Shuttle Main Engine expansion ratio modifications, in the cases of both single-stage and two-stage systems. Two-position nozzles were employed; after varying the lower expansion ratio while the higher was held constant at 120, the lower expansion ratio was held constant at 40 or 60 while the higher expansion ratio was varied. The expansion ratios for minimum vehicle dry mass are different for single-stage and two-stage systems. For two-stage systems, a single expansion ratio of 77.5 provides a lower dry mass than any two-position nozzle.

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

    NASA Technical Reports Server (NTRS)

    Keba, John E.

    1999-01-01

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

  20. Flow processes in overexpanded chemical rocket nozzles. Part 2: Side loads due to asymmetric separation

    NASA Technical Reports Server (NTRS)

    Schmucker, R. H.

    1984-01-01

    Methods for measuring the lateral forces, occurring as a result of asymmetric nozzle flow separation, are discussed. The effect of some parameters on the side load is explained. A new method was developed for calculation of the side load. The values calculated are compared with side load data of the J-2 engine. Results are used for predicting side loads of the space shuttle main engine.

  1. Analysis and Results from a Flush Airdata Sensing System in Close Proximity to Firing Rocket Nozzles

    NASA Technical Reports Server (NTRS)

    Ali, Aliyah N.; Borrer, Jerry L.

    2013-01-01

    This paper presents information regarding the nosecap Flush Airdata Sensing (FADS) system on Orion’s Pad Abort 1 (PA-1) vehicle. The purpose of the nosecap FADS system was to test whether or not useful data could be obtained from a FADS system if it was placed in close proximity to firing rocket nozzles like the Attitude Control Motor (ACM) nozzles on the PA-1 Launch Abort System. The nosecap FADS system used pressure measurements from a series of pressure ports which were arranged in a cruciform pattern and flush with the surface of the vehicle to estimate values of angle of attack, angle of sideslip, Mach number, impact pressure, and freestream static pressure. This paper will present the algorithms employed by the FADS system along with the development of the calibration datasets and a comparison of the final results to the Best Estimated Trajectory (BET) data for PA-1. Also presented in this paper is a Computational Fluid Dynamics (CFD) study to explore the impact of the ACM on the nosecap FADS system. The comparison of the nosecap FADS system results to the BET and the CFD study showed that more investigation is needed to quantify the impact of the firing rocket motors on the FADS system.

  2. Large liquid rocket engine transient performance simulation system

    NASA Technical Reports Server (NTRS)

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

    1989-01-01

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

  3. AJ26 rocket engine testing news briefing

    NASA Technical Reports Server (NTRS)

    2010-01-01

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

  4. AJ26 rocket engine testing news briefing

    NASA Image and Video Library

    2010-02-24

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

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

    NASA Technical Reports Server (NTRS)

    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.

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

    NASA Astrophysics Data System (ADS)

    Jiang, Tsung Leo; Chiang, Wei-Tang; Jang, Shyh-Dihng

    1995-05-01

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

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

    SciTech Connect

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

    1995-05-01

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

  8. Thermographic Nondestructive Evaluation of the Space Shuttle Main Engine Nozzle

    NASA Technical Reports Server (NTRS)

    Walker, James L.; Lansing, Matthew D.; Russell, Samuel S.; Caraccioli, Paul; Whitaker, Ann F. (Technical Monitor)

    2000-01-01

    The methods and results presented in this summary address the thermographic identification of interstitial leaks in the Space Shuttle Main Engine nozzles. A highly sensitive digital infrared camera is used to record the minute cooling effects associated with a leak source, such as a crack or pinhole, hidden within the nozzle wall by observing the inner "hot wall" surface as the nozzle is pressurized. These images are enhanced by digitally subtracting a thermal reference image taken before pressurization, greatly diminishing background noise. The method provides a nonintrusive way of localizing the tube that is leaking and the exact leak source position to within a very small axial distance. Many of the factors that influence the inspectability of the nozzle are addressed; including pressure rate, peak pressure, gas type, ambient temperature and surface preparation.

  9. Air breathing engine/rocket trajectory optimization

    NASA Technical Reports Server (NTRS)

    Smith, V. K., III

    1979-01-01

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

  10. Nuclear thermal rocket engine operation and control

    SciTech Connect

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

    1993-06-01

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

  11. Combustion dynamics in liquid rocket engines

    NASA Technical Reports Server (NTRS)

    Mclain, W. H.

    1971-01-01

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

  12. Design and Evaluation of Dual-Expander Aerospike Nozzle Upper Stage Engine

    DTIC Science & Technology

    2014-09-18

    the original cycle model in preparation for optimizing the engine’s specific impulse and thrust-to-weight ratio. The changes to the model allowed...specific impulse comparable to the RL10. Parametric studies using the new model verified the Dual-Expander Aerospike Nozzle architecture conforms to rocket...design delivers 35,000 lbf of vacuum thrust at 469.4 seconds of vacuum specific impulse with a thrust-to-weight ratio of 127.2 in an engine that is one

  13. Radiation effect on rocket engine performance

    NASA Technical Reports Server (NTRS)

    Chiu, Huei-Huang

    1988-01-01

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

  14. AFRPL Graphite Performance Prediction Program. Improved Capability for the Design and Ablation Performance Prediction of Advanced Air Force Solid Propellant Rocket Nozzles

    DTIC Science & Technology

    1976-12-01

    characterization test matrix. 3.2.1.1 Rocket Motor Environments Rocket motor environments were based on three advanced MX propellants , namely, * XLDB * HTPB * PEG...107 41 Predicted Nozzle Respoise to HTPB Propellant , 60.0 Seconds .... ............ 108 42 Nozzle Geometry, Rocketdyne Condor Nozzle...species as a function of temperature for an HTPB propellant is shown in Figure 8. This solution represents the kinetically controlled ablation of edge

  15. Physical attachment of sensors in rocket engines

    NASA Astrophysics Data System (ADS)

    Blue, Craig A.; Robson, Mark T.; Jiang, Y.-Q.; Lin, Ray Y.

    The bonding of sensor elements to rocket engine components by means of rapid radiation heating (RRH) drastically reduces the processing time required; neither is vacuum processing needed, as in conventional bonding processes. RRH bonding involves the melting of a brazing alloy by the heat that is radiatively delivered from a source to the sample surface, followed by sample conduction; the process is therefore inapplicable in cases where the sample is inherently a poor thermal conductor, as in the case of ceramics.

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

    SciTech Connect

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

    1989-06-01

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

  17. Exhaust Nozzle for a Multitube Detonative Combustion Engine

    NASA Technical Reports Server (NTRS)

    Bratkovich, Thomas E.; Williams, Kevin E.; Bussing, Thomas R. A.; Lidstone, Gary L.; Hinkey, John B.

    2004-01-01

    An improved type of exhaust nozzle has been invented to help optimize the performances of multitube detonative combustion engines. The invention is applicable to both air-breathing and rocket engines used to propel some aircraft and spacecraft, respectively. In a detonative combustion engine, thrust is generated through the expulsion of combustion products from a detonation process in which combustion takes place in a reaction zone coupled to a shock wave. The combustion releases energy to sustain the shock wave, while the shock wave enhances the combustion in the reaction zone. The coupled shockwave/reaction zone, commonly referred to as a detonation, propagates through the reactants at very high speed . typically of the order of several thousands of feet per second (of the order of 1 km/s). The very high speed of the detonation forces combustion to occur very rapidly, thereby contributing to high thermodynamic efficiency. A detonative combustion engine of the type to which the present invention applies includes multiple parallel cylindrical combustion tubes, each closed at the front end and open at the rear end. Each tube is filled with a fuel/oxidizer mixture, and then a detonation wave is initiated at the closed end. The wave propagates rapidly through the fuel/oxidizer mixture, producing very high pressure due to the rapid combustion. The high pressure acting on the closed end of the tube contributes to forward thrust. When the detonation wave reaches the open end of the tube, it produces a blast wave, behind which the high-pressure combustion products are expelled from the tube. The process of filling each combustion tube with a detonable fuel/oxidizer mixture and then producing a detonation repeated rapidly to obtain repeated pulses of thrust. Moreover, the multiple combustion tubes are filled and fired in a repeating sequence. Hence, the pressure at the outlet of each combustion tube varies cyclically. A nozzle of the present invention channels the

  18. Nozzle

    DOEpatents

    Chen, Alexander G.; Cohen, Jeffrey M.

    2009-06-16

    A fuel injector has a number of groups of nozzles. The groups are generally concentric with an injector axis. Each nozzle defines a gas flowpath having an outlet for discharging a fuel/air mixture jet. There are means for introducing the fuel to the air. One or more groups of the nozzles are oriented to direct the associated jets skew to the injector axis.

  19. Analysis of a Radioisotope Thermal Rocket Engine

    NASA Technical Reports Server (NTRS)

    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.

  20. Quantifying Instability Sources in Liquid Rocket Engines

    NASA Technical Reports Server (NTRS)

    Farmer, Richard C.; Cheng, Gary C.

    2000-01-01

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

  1. Software for Collaborative Engineering of Launch Rockets

    NASA Technical Reports Server (NTRS)

    Stanley, Thomas Troy

    2003-01-01

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

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

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

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

  3. Experimental Altitude Performance of JP-4 Fuel and Liquid-Oxygen Rocket Engine with an Area Ratio of 48

    NASA Technical Reports Server (NTRS)

    Fortini, Anthony; Hendrix, Charles D.; Huff, Vearl N.

    1959-01-01

    The performance for four altitudes (sea-level, 51,000, 65,000, and 70,000 ft) of a rocket engine having a nozzle area ratio of 48.39 and using JP-4 fuel and liquid oxygen as a propellant was evaluated experimentally by use of a 1000-pound-thrust engine operating at a chamber pressure of 600 pounds per square inch absolute. The altitude environment was obtained by a rocket-ejector system which utilized the rocket exhaust gases as the pumping fluid of the ejector. Also, an engine having a nozzle area ratio of 5.49 designed for sea level was tested at sea-level conditions. The following table lists values from faired experimental curves at an oxidant-fuel ratio of 2.3 for various approximate altitudes.

  4. The prediction of nonlinear three dimensional combustion instability in liquid rockets with conventional nozzles

    NASA Technical Reports Server (NTRS)

    Powell, E. A.; Zinn, B. T.

    1973-01-01

    An analytical technique is developed to solve nonlinear three-dimensional, transverse and axial combustion instability problems associated with liquid-propellant rocket motors. The Method of Weighted Residuals is used to determine the nonlinear stability characteristics of a cylindrical combustor with uniform injection of propellants at one end and a conventional DeLaval nozzle at the other end. Crocco's pressure sensitive time-lag model is used to describe the unsteady combustion process. The developed model predicts the transient behavior and nonlinear wave shapes as well as limit-cycle amplitudes and frequencies typical of unstable motor operation. The limit-cycle amplitude increases with increasing sensitivity of the combustion process to pressure oscillations. For transverse instabilities, calculated pressure waveforms exhibit sharp peaks and shallow minima, and the frequency of oscillation is within a few percent of the pure acoustic mode frequency. For axial instabilities, the theory predicts a steep-fronted wave moving back and forth along the combustor.

  5. Thrust chamber life prediction. Volume 1: Mechanical and physical properties of high performance rocket nozzle materials

    NASA Technical Reports Server (NTRS)

    Esposito, J. J.; Zabora, R. F.

    1975-01-01

    Pertinent mechanical and physical properties of six high conductivity metals were determined. The metals included Amzirc, NARloy Z, oxygen free pure copper, electroformed copper, fine silver, and electroformed nickel. Selection of these materials was based on their possible use in high performance reusable rocket nozzles. The typical room temperature properties determined for each material included tensile ultimate strength, tensile yield strength, elongation, reduction of area, modulus of elasticity, Poisson's ratio, density, specific heat, thermal conductivity, and coefficient of thermal expansion. Typical static tensile stress-strain curves, cyclic stress-strain curves, and low-cycle fatigue life curves are shown. Properties versus temperature are presented in graphical form for temperatures from 27.6K (-410 F) to 810.9K (1000 F).

  6. High altitude chemically reacting gas particle mixtures. Volume 3: Computer code user's and applications manual. [rocket nozzle and orbital plume flow fields

    NASA Technical Reports Server (NTRS)

    Smith, S. D.

    1984-01-01

    A users manual for the RAMP2 computer code is provided. The RAMP2 code can be used to model the dominant phenomena which affect the prediction of liquid and solid rocket nozzle and orbital plume flow fields. The general structure and operation of RAMP2 are discussed. A user input/output guide for the modified TRAN72 computer code and the RAMP2F code is given. The application and use of the BLIMPJ module are considered. Sample problems involving the space shuttle main engine and motor are included.

  7. Key technology for reusable rocket engine turbopump

    NASA Astrophysics Data System (ADS)

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

    2002-03-01

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

  8. Tests of a D vented thrust deflecting nozzle behind a simulated turbofan engine

    NASA Technical Reports Server (NTRS)

    Watson, T. L.

    1982-01-01

    A D vented thrust deflecting nozzle applicable to subsonic V/STOL aircraft was tested behind a simulated turbofan engine in the verticle thrust stand. Nozzle thrust, fan operating characteristics, nozzle entrance conditions, and static pressures were measured. Nozzle performance was measured for variations in exit area and thrust deflection angle. Six core nozzle configurations, the effect of core exit axial location, mismatched core and fan stream nozzle pressure ratios, and yaw vane presence were evaluated. Core nozzle configuration affected performance at normal and engine out operating conditions. Highest vectored nozzle performance resulted for a given exit area when core and fan stream pressure were equal. Its is concluded that high nozzle performance can be maintained at both normal and engine out conditions through control of the nozzle entrance Mach number with a variable exit area.

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

    NASA Technical Reports Server (NTRS)

    Wang, Ten-See

    2005-01-01

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

  10. Jet engine nozzle exit configurations and associated systems and methods

    NASA Technical Reports Server (NTRS)

    Mengle, Vinod G. (Inventor)

    2011-01-01

    Nozzle exit configurations and associated systems and methods are disclosed. An aircraft system in accordance with one embodiment includes a jet engine exhaust nozzle having an internal flow surface and an exit aperture, with the exit aperture having a perimeter that includes multiple projections extending in an aft direction. Aft portions of individual neighboring projections are spaced apart from each other by a gap, and a geometric feature of the multiple can change in a monotonic manner along at least a portion of the perimeter.

  11. Jet Engine Nozzle Exit Configurations and Associated Systems and Methods

    NASA Technical Reports Server (NTRS)

    Mengle, Vinod G. (Inventor)

    2013-01-01

    Nozzle exit configurations and associated systems and methods are disclosed. An aircraft system in accordance with one embodiment includes a jet engine exhaust nozzle having an internal flow surface and an exit aperture, with the exit aperture having a perimeter that includes multiple projections extending in an aft direction. Aft portions of individual neighboring projections are spaced apart from each other by a gap, and a geometric feature of the multiple can change in a monotonic manner along at least a portion of the perimeter.

  12. Rocket Engine Innovations Advance Clean Energy

    NASA Technical Reports Server (NTRS)

    2012-01-01

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

  13. Reusable Rocket Engine Turbopump Health Management System

    NASA Technical Reports Server (NTRS)

    Surko, Pamela

    1994-01-01

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

  14. Parallelization of Rocket Engine System Software (Press)

    NASA Technical Reports Server (NTRS)

    Cezzar, Ruknet

    1996-01-01

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

  15. Parallelization of Rocket Engine System Software (Press)

    NASA Technical Reports Server (NTRS)

    Cezzar, Ruknet

    1996-01-01

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

  16. GRCop-84 Developed for Rocket Engines

    NASA Technical Reports Server (NTRS)

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

    2001-01-01

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

  17. Engine-Level Simulation of Liquid Rocket Combustion Instabilities

    DTIC Science & Technology

    2013-01-01

    Chapter 3. DATES COVERED (From - To) January 2013-July 2013 4. TITLE AND SUBTITLE Engine -Level Simulation of Liquid Rocket Combustion Instabilities...ABSTRACT A numerical investigation into combustion instability in liquid rocket engines is undertaken using large eddy simulations (LES). HPCMP resources...have been applied to demonstrate the ability to simulate combustion instability in liquid rocket engines and to gain further understanding of these

  18. Future Modeling Needs in Pulse Detonation Rocket Engine Design

    NASA Technical Reports Server (NTRS)

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

    2001-01-01

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

  19. Rocket engine diagnostics using qualitative modeling techniques

    NASA Technical Reports Server (NTRS)

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

    1992-01-01

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

  20. Rocket engine diagnostics using qualitative modeling techniques

    NASA Technical Reports Server (NTRS)

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

    1992-01-01

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

  1. Life extending control for rocket engines

    NASA Technical Reports Server (NTRS)

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

    1992-01-01

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

  2. Rocket engine diagnostics using qualitative modeling techniques

    NASA Technical Reports Server (NTRS)

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

    1992-01-01

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

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

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

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

  4. Advanced oxygen-hydrocarbon rocket engine study

    NASA Technical Reports Server (NTRS)

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

    1981-01-01

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

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

    NASA Technical Reports Server (NTRS)

    Harmon, T. J.; Roschak, E.

    1993-01-01

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

  6. Liquid fuel injection elements for rocket engines

    NASA Technical Reports Server (NTRS)

    Cox, George B., Jr. (Inventor)

    1993-01-01

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

  7. Engine protection system for recoverable rocket booster

    NASA Technical Reports Server (NTRS)

    Shelby, Jr., Jerry A. (Inventor)

    1994-01-01

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

  8. Rocket engine hot-spot detector

    NASA Astrophysics Data System (ADS)

    Collamore, F. N.

    1985-04-01

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

  9. Liquid fuel injection elements for rocket engines

    NASA Astrophysics Data System (ADS)

    Cox, George B., Jr.

    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 herein 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.

  10. Liquid fuel injection elements for rocket engines

    NASA Astrophysics Data System (ADS)

    Cox, George B., Jr.

    1993-11-01

    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.

  11. A Rocket Engine Design Expert System

    NASA Technical Reports Server (NTRS)

    Davidian, Kenneth J.

    1989-01-01

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

  12. A rocket engine design expert system

    NASA Technical Reports Server (NTRS)

    Davidian, Kenneth J.

    1989-01-01

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

  13. Acoustic/infrasonic rocket engine signatures

    NASA Astrophysics Data System (ADS)

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

    2003-09-01

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

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

    NASA Astrophysics Data System (ADS)

    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.

  15. Boundary layer development as a function of chamber pressure in the NASA Lewis 1030:1 area ratio rocket nozzle

    NASA Technical Reports Server (NTRS)

    Smith, Tamara A.

    1988-01-01

    Through the use of theoretical predictions of fluid properties and experimental heat transfer and thrust measurements, the zones of laminar, transitional, and turbulent boundary layer flow were defined for the NASA Lewis 1030:1 area ratio rocket nozzle. Tests were performed on the nozzle at chamber pressures from 350 to 100 psia. For these conditions, the throat diameter Reynolds numbers varied from 300,000 to 1 million. The propellants used were gaseous hydrogen and gaseous oxygen. Thrust measurements and nozzle outer wall temperature measurements were taken during the 3-sec test runs. Comparison of experimental heat transfer and thrust data with the corresponding predictions from the Two-Dimensional Kinetics (TDK) nozzle analysis program indicated laminar flow in the nozzle at a throat diameter Reynolds number of 320,000 or chamber pressure of 360 psia. Comparison of experimental and predicted heat transfer data indicated transitional flow up to and including a chamber pressure of 1000 psia. Predicted values of the axisymmetric acceleration parameter within the convergent and divergent nozzle were consistent with the above results. Based upon an extrapolation of the heat transfer data and predicted distributions of the axisymmetric acceleration parameter, transitional flow was predicted up to a throat diameter Reynolds number of 220,000 or 2600-psia chamber pressure. Above 2600-psia chamber pressure, fully developed turbulent flow was predicted.

  16. Boundary layer development as a function of chamber pressure in the NASA Lewis 1030:1 area ratio rocket nozzle

    NASA Technical Reports Server (NTRS)

    Smith, Tamara A.

    1988-01-01

    Through the use of theoretical predictions of fluid properties and experimental heat transfer and thrust measurements, the zones of laminar, transitional, and turbulent boundary layer flow were defined for the NASA Lewis 1039:1 area ratio rocket nozzle. Tests were performed on the nozzle at chamber pressures from 350 to 100 psia. For these conditions, the throat diameter Reynolds numbers varied from 300,000 to 1 million. The propellants used were gaseous hydrogen and gaseous oxygen. Thrust measurements and nozzle outer wall temperature measurements were taken during the 3-sec test runs. Comparison of experimental heat transfer and thrust data with the corresponding predictions from the Two-Dimensional Kinetics (TDK) nozzle analysis program indicated laminar flow in the nozzle at a throat diameter Reynolds number of 320,000 or chamber pressure of 360 psia. Comparison of experimental and predicted heat transfer data indicated transitional flow up to and including a chamber pressure of 1000 psia. Predicted values of the axisymmetric acceleration parameter within the convergent and divergent nozzle were consistent with the above results. Based upon an extrapolation of the heat transfer data and predicted distributions of the axisymmetric acceleration parameter, transitional flow was predicted up to a throat diameter Reynolds number of 220,000 or 2600-psia chamber pressure. Above 2600-psia chamber pressure, fully developed turbulent flow was predicted.

  17. NASA Lewis Thermal Barrier Feasibility Investigated for Use in Space Shuttle Solid-Rocket Motor Nozzle-to-Case Joints

    NASA Technical Reports Server (NTRS)

    Steinetz, Bruce M.; Dunlap, Patrick H., Jr.

    1999-01-01

    Assembly joints of modern solid-rocket motor cases are usually sealed with conventional O-ring seals. The 5500 F combustion gases produced by rocket motors are kept a safe distance away from the seals by thick layers of insulation and by special compounds that fill assembly split-lines in the insulation. On limited occasions, NASA has observed charring of the primary O-rings of the space shuttle solid-rocket nozzle-assembly joints due to parasitic leakage paths opening up in the gap-fill compounds during rocket operation. Thus, solid-rocket motor manufacturer Thiokol approached the NASA Lewis Research Center about the possibility of applying Lewis braided-fiber preform seal as a thermal barrier to protect the O-ring seals. This thermal barrier would be placed upstream of the primary O-rings in the nozzle-to-case joints to prevent hot gases from impinging on the O-ring seals (see the following illustration). The illustration also shows joints 1 through 5, which are potential sites where the thermal barrier could be used.

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

    NASA Astrophysics Data System (ADS)

    St. Germain, Brad David

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

  19. Robust Strategy for Rocket Engine Health Monitoring

    NASA Technical Reports Server (NTRS)

    Santi, L. Michael

    2001-01-01

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

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

    NASA Technical Reports Server (NTRS)

    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.

  1. Catalyst Development for Hydrogen Peroxide Rocket Engines

    NASA Technical Reports Server (NTRS)

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

    1999-01-01

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

  2. Catalyst Development for Hydrogen Peroxide Rocket Engines

    NASA Technical Reports Server (NTRS)

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

    1999-01-01

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

  3. Fiber-reinforced ceramic composites for Earth-to-orbit rocket engine turbines

    NASA Technical Reports Server (NTRS)

    Brockmeyer, Jerry W.; Schnittgrund, Gary D.

    1990-01-01

    Fiber reinforced ceramic matrix composites (FRCMC) are emerging materials systems that offer potential for use in liquid rocket engines. Advantages of these materials in rocket engine turbomachinery include performance gain due to higher turbine inlet temperature, reduced launch costs, reduced maintenance with associated cost benefits, and reduced weight. This program was initiated to assess the state of FRCMC development and to propose a plan for their implementation into liquid rocket engine turbomachinery. A complete range of FRCMC materials was investigated relative to their development status and feasibility for use in the hot gas path of earth-to-orbit rocket engine turbomachinery. Of the candidate systems, carbon fiber-reinforced silicon carbide (C/SiC) offers the greatest near-term potential. Critical hot gas path components were identified, and the first stage inlet nozzle and turbine rotor of the fuel turbopump for the liquid oxygen/hydrogen Space Transportation Main Engine (STME) were selected for conceptual design and analysis. The critical issues associated with the use of FRCMC were identified. Turbine blades were designed, analyzed and fabricated. The Technology Development Plan, completed as Task 5 of this program, provides a course of action for resolution of these issues.

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

  13. Plume flowfield analysis of the shuttle primary Reaction Control System (RCS) rocket engine

    NASA Technical Reports Server (NTRS)

    Hueser, J. E.; Brock, F. J.

    1990-01-01

    A solution was generated for the physical properties of the Shuttle RCS 4000 N (900 lb) rocket engine exhaust plume flowfield. The modeled exhaust gas consists of the five most abundant molecular species, H2, N2, H2O, CO, and CO2. The solution is for a bare RCS engine firing into a vacuum; the only additional hardware surface in the flowfield is a cylinder (=engine mount) which coincides with the nozzle lip outer corner at X = 0, extends to the flowfield outer boundary at X = -137 m and is coaxial with the negative symmetry axis. Continuum gas dynamic methods and the Direct Simulation Monte Carlo (DSMC) method were combined in an iterative procedure to produce a selfconsistent solution. Continuum methods were used in the RCS nozzle and in the plume as far as the P = 0.03 breakdown contour; the DSMC method was used downstream of this continuum flow boundary. The DSMC flowfield extends beyond 100 m from the nozzle exit and thus the solution includes the farfield flow properties, but substantial information is developed on lip flow dynamics and thus results are also presented for the flow properties in the vicinity of the nozzle lip.

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

    NASA Image and Video Library

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

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

    SciTech Connect

    Bolonkin, A.

    1991-01-01

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

  16. Spectroscopy and multivariate analyses applications related to solid rocket nozzle bondline

    NASA Technical Reports Server (NTRS)

    Arendale, W. F.; Hatcher, Richard; Benson, Brian; Workman, Gary L.

    1991-01-01

    Chemical composition and molecular orientation define the properties of materials. Information related to chemical composition and molecular configuration is obtained by various forms of spectroscopy. Software algorithms developed for multivariate analyses, expert systems, and Artificial Intelligence (AI) are used to conduct repetitive operations. The techniques are believed to be of particular significance toward achieving TQM objectives. The objective was to obtain information related to the quality of the bondline in the solid rocket motor, SRM, nozzle. Hysol 934 NA, a room temperature curing epoxide resin, is used as the bonding agent. A good bond requires that the adhesive be placed on a properly prepared metal surface, the adhesives Part A and B be mixed in appropriate ratio from material within shelf life specifications. Spectroscopic data was obtained for surfaces prepared according to specifications, contaminated metal surfaces, samples of the epoxide adhesive at times that represent shelf aging from 3 months to 2 years, several mix ratio of A to B, and curing material. Temperature was found to be a significant factor. The study concentrated on pot life and mix ratio.

  17. Computation of the space shuttle solid rocket booster nozzle start-up transient flow

    SciTech Connect

    Cline, M.C.; Wilmoth, R.G.

    1984-01-01

    The first NASA Space Shuttle flight (STS-1) produced an overpressure wave that exceeded preflight predictions by as much as 5 to 1. This second overpressure wave occurred just after the solid rocket booster (SRB) igniter wave. To understand this overpressure phenomenon, a numerical simulation effort was undertaken. Both the SRB static firing test and STS-1 geometries were studied for two-dimensional, inviscid and viscous flow. The inviscid calculations did not produce significant second overpressure waves. However, the viscous calculations did produce second overpressure waves that qualitatively agree with experiment. These overpressure waves were present in both the static firing test and STS-1 geometries. This second overpressure wave is generated by the motion of the boundary layer separation point and the subsequent radial motion of the exhaust jet during the start-up of the SRB nozzle flow. The presence of the mobile launch platform exhaust hole wall amplifies this wave, but does not appear to be the source of any additional overpressure waves. The lack of good quantitative agreement between theory and experiment indicates that other overpressure sources, not accounted for by this simulation, may be present.

  18. Linear Approximation to Optimal Control Allocation for Rocket Nozzles with Elliptical Constraints

    NASA Technical Reports Server (NTRS)

    Orr, Jeb S.; Wall, Johnm W.

    2011-01-01

    In this paper we present a straightforward technique for assessing and realizing the maximum control moment effectiveness for a launch vehicle with multiple constrained rocket nozzles, where elliptical deflection limits in gimbal axes are expressed as an ensemble of independent quadratic constraints. A direct method of determining an approximating ellipsoid that inscribes the set of attainable angular accelerations is derived. In the case of a parameterized linear generalized inverse, the geometry of the attainable set is computationally expensive to obtain but can be approximated to a high degree of accuracy with the proposed method. A linear inverse can then be optimized to maximize the volume of the true attainable set by maximizing the volume of the approximating ellipsoid. The use of a linear inverse does not preclude the use of linear methods for stability analysis and control design, preferred in practice for assessing the stability characteristics of the inertial and servoelastic coupling appearing in large boosters. The present techniques are demonstrated via application to the control allocation scheme for a concept heavy-lift launch vehicle.

  19. Large-displacement structural durability analyses of simple bend specimen emulating rocket nozzle liners

    NASA Technical Reports Server (NTRS)

    Arya, Vinod K.; Halford, Gary R.

    1994-01-01

    Large-displacement elastic and elastic-plastic, finite-element stress-strain analyses of an oxygen-tree high-conductivity (OFHC) copper plate specimen were performed using an updated Lagrangian formulation. The plate specimen is intended for low-cost experiments that emulate the most important thermomechanical loading and failure modes of a more complex rocket nozzle. The plate, which is loaded in bending at 593 C, contains a centrally located and internally pressurized channel. The cyclic crack initiation lives were estimated using the results from the analyses and isothermal strain-controlled low-cycle fatigue data for OFHC copper. A comparison of the predicted and experimental cyclic lives showed that an elastic analysis predicts a longer cyclic life than that observed in experiments by a factor greater than 4. The results from elastic-plastic analysis for the plate bend specimen, however, predicted a cyclic life in close agreement with experiment, thus justifying the need for the more rigorous stress-strain analysis.

  20. Large-displacement structural durability analyses of simple bend specimen emulating rocket nozzle liners

    NASA Astrophysics Data System (ADS)

    Arya, Vinod K.; Halford, Gary R.

    1994-06-01

    Large-displacement elastic and elastic-plastic, finite-element stress-strain analyses of an oxygen-tree high-conductivity (OFHC) copper plate specimen were performed using an updated Lagrangian formulation. The plate specimen is intended for low-cost experiments that emulate the most important thermomechanical loading and failure modes of a more complex rocket nozzle. The plate, which is loaded in bending at 593 C, contains a centrally located and internally pressurized channel. The cyclic crack initiation lives were estimated using the results from the analyses and isothermal strain-controlled low-cycle fatigue data for OFHC copper. A comparison of the predicted and experimental cyclic lives showed that an elastic analysis predicts a longer cyclic life than that observed in experiments by a factor greater than 4. The results from elastic-plastic analysis for the plate bend specimen, however, predicted a cyclic life in close agreement with experiment, thus justifying the need for the more rigorous stress-strain analysis.

  1. Test Stand at the Rocket Engine Test Facility

    NASA Image and Video Library

    1973-02-21

    The thrust stand in the Rocket Engine Test Facility at the National Aeronautics and Space Administration (NASA) Lewis Research Center in Cleveland, Ohio. The Rocket Engine Test Facility was constructed in the mid-1950s to expand upon the smaller test cells built a decade before at the Rocket Laboratory. The $2.5-million Rocket Engine Test Facility could test larger hydrogen-fluorine and hydrogen-oxygen rocket thrust chambers with thrust levels up to 20,000 pounds. Test Stand A, seen in this photograph, was designed to fire vertically mounted rocket engines downward. The exhaust passed through an exhaust gas scrubber and muffler before being vented into the atmosphere. Lewis researchers in the early 1970s used the Rocket Engine Test Facility to perform basic research that could be utilized by designers of the Space Shuttle Main Engines. A new electronic ignition system and timer were installed at the facility for these tests. Lewis researchers demonstrated the benefits of ceramic thermal coatings for the engine’s thrust chamber and determined the optimal composite material for the coatings. They compared the thermal-coated thrust chamber to traditional unlined high-temperature thrust chambers. There were more than 17,000 different configurations tested on this stand between 1973 and 1976. The Rocket Engine Test Facility was later designated a National Historic Landmark for its role in the development of liquid hydrogen as a propellant.

  2. CFD Simulation of Liquid Rocket Engine Injectors

    NASA Technical Reports Server (NTRS)

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

    2001-01-01

    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

  3. Enriched Storable Oxidizers for Rocket Engines

    NASA Technical Reports Server (NTRS)

    Sackheim, R. L.; Herdy, J. R., Jr.

    2010-01-01

    The name "enriched storable oxidizers" (ESOs) has been coined for a family of optimized mixtures of between two and four oxidizer fluids. For most applications, the constituents of these mixtures would be nitrogen tetroxide (N2O4), nitrous oxide (N2O), and nitrogen dioxide (NO2); in some applications, the mixtures might include inhibited red fuming nitric acid [IRFN (which consists of red fuming nitric acid to which some hydrogen fluoride is added to reduce its corrosive effect]. The optimum proportions of these constituents would be different for different applications. ESOs were originally proposed for use in spacecraft and launch-rocket propulsion systems: ESOs could be especially useful in advanced spacecraft propulsion systems that could operate in multiple modes. ESOs might also be useful in special terrestrial applications that could include ramjet and scramjet aircraft engines.

  4. Parallelization of Rocket Engine Simulator Software (PRESS)

    NASA Technical Reports Server (NTRS)

    Cezzar, Ruknet

    1997-01-01

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

  5. Software Estimates Costs of Testing Rocket Engines

    NASA Technical Reports Server (NTRS)

    2002-01-01

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

  6. Software Estimates Costs of Testing Rocket Engines

    NASA Technical Reports Server (NTRS)

    2002-01-01

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

  7. Orbit transfer rocket engine technology program: Advanced engine study

    NASA Technical Reports Server (NTRS)

    Erickson, C. M.

    1992-01-01

    In Task D.6 of the Advanced Engine Study, three primary subtasks were accomplished: (1) design of parametric data; (2) engine requirement variation studies; and (3) vehicle study/engine study coordination. Parametric data were generated for vacuum thrusts ranging from 7500 lbf to 50,000 lbf, nozzle expansion ratios from 600 to 1200, and engine mixture ratios from 5:1 to 7:1. Failure Modes and Effects Analysis (FMEA) was used as a departure point for these parametric analyses. These data are intended to assist in definition and trade studies. In the Engine Requirements Variation Studies, the individual effects of increasing the throttling ratio from 10:1 to 20:1 and requiring the engine to operate at a maximum mixture ratio of 12:1 were determined. Off design engine balances were generated at these extreme conditions and individual component operating requirements analyzed in detail. Potential problems were identified and possible solutions generated. In the Vehicle Study/Engine Study coordination subtask, vehicle contractor support was provided as needed, addressing a variety of issues uncovered during vehicle trade studies. This support was primarily provided during Technical Interchange Meetings (TIM) in which Space Exploration Initiative (SEI) studies were addressed.

  8. Metallized Gelled Propellants: Oxygen/RP-1/Aluminum Rocket Engine Calorimeter Heat Transfer Measurements and Analysis

    NASA Technical Reports Server (NTRS)

    Palaszewski, Bryan

    1997-01-01

    A set of analyses was conducted to determine the heat transfer characteristics of metallized gelled liquid propellants in a rocket engine. The analyses used the data from experiments conducted with a small 30- to 40-lbf thrust engine composed of a modular injector, igniter, chamber and nozzle. The fuels used were traditional liquid RP-1 and gelled RP-1 with 0-wt %, 5-wt%, and 55-wt% loadings of aluminum with silicon dioxide gellant, and gaseous oxygen as the oxidizer. Heat transfer was computed based on measurements using calorimeter rocket chamber and nozzle hardware with a total of 31 cooling channels. A gelled fuel coating formed in the 0-, 5- and 55-wt% engines, and the coating was composed of unburned gelled fuel and partially combusted RP-1. The coating caused a large decrease in calorimeter engine heat flux in the last half of the chamber for the 0- and 5-wt% RP-1/Al. This heat flux reduction effect was analyzed by comparing engine runs and the changes in the heat flux during a run as well as from run to run. Heat transfer and time-dependent heat flux analyses and interpretations are provided. The 5- and 55-wt% RP-1/Al fueled engines had the highest chamber heat fluxes, with the 5-wt% fuel having the highest throat flux. This result is counter to the predicted result, where the 55 wt% fuel has the highest combustion and throat temperature, and therefore implies that it would deliver the highest throat heat flux. The 5-wt% RP-1/Al produced the most influence on the engine heat transfer and the heat flux reduction was caused by the formation of a gelled propellant layer in the chamber and nozzle.

  9. Axisymmetric computational fluid dynamics analysis of a film/dump-cooled rocket nozzle plume

    NASA Technical Reports Server (NTRS)

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

    1993-01-01

    Prediction of convective base heating rates for a new launch vehicle presents significant challenges to analysts concerned with base environments. The present effort seeks to augment classical base heating scaling techniques via a detailed investigation of the exhaust plume shear layer of a single H2/O2 Space Transportation Main Engine (STME). Use of fuel-rich turbine exhaust to cool the STME nozzle presented concerns regarding potential recirculation of these gases to the base region with attendant increase in the base heating rate. A pressure-based full Navier-Stokes computational fluid dynamics (CFD) code with finite rate chemistry is used to predict plumes for vehicle altitudes of 10 kft and 50 kft. Levels of combustible species within the plume shear layers are calculated in order to assess assumptions made in the base heating analysis.

  10. Axisymmetric computational fluid dynamics analysis of a film/dump-cooled rocket nozzle plume

    NASA Astrophysics Data System (ADS)

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

    1993-06-01

    Prediction of convective base heating rates for a new launch vehicle presents significant challenges to analysts concerned with base environments. The present effort seeks to augment classical base heating scaling techniques via a detailed investigation of the exhaust plume shear layer of a single H2/O2 Space Transportation Main Engine (STME). Use of fuel-rich turbine exhaust to cool the STME nozzle presented concerns regarding potential recirculation of these gases to the base region with attendant increase in the base heating rate. A pressure-based full Navier-Stokes computational fluid dynamics (CFD) code with finite rate chemistry is used to predict plumes for vehicle altitudes of 10 kft and 50 kft. Levels of combustible species within the plume shear layers are calculated in order to assess assumptions made in the base heating analysis.

  11. Advanced Methods for Aircraft Engine Thrust and Noise Benefits: Nozzle-Inlet Flow Analysis

    NASA Technical Reports Server (NTRS)

    Morgan, Morris H.; Gilinsky, Mikhail; Patel, Kaushal; Coston, Calvin; Blankson, Isaiah M.

    2003-01-01

    The research is focused on a wide regime of problems in the propulsion field as well as in experimental testing and theoretical and numerical simulation analyses for advanced aircraft and rocket engines. Results obtained are based on analytical methods, numerical simulations and experimental tests at the NASA LaRC and Hampton University computer complexes and experimental facilities. The main objective of this research is injection, mixing and combustion enhancement in propulsion systems. The sub-projects in the reporting period are: (A) Aero-performance and acoustics of Telescope-shaped designs. The work included a pylon set application for SCRAMJET. (B) An analysis of sharp-edged nozzle exit designs for effective fuel injection into the flow stream in air-breathing engines: triangular-round and diamond-round nozzles. (C) Measurement technique improvements for the HU Low Speed Wind Tunnel (HU LSWT) including an automatic data acquisition system and a two component (drag-lift) balance system. In addition, a course in the field of aerodynamics was developed for the teaching and training of HU students.

  12. Advanced Methods for Aircraft Engine Thrust and Noise Benefits: Nozzle-Inlet Flow Analysis

    NASA Technical Reports Server (NTRS)

    Morgan, Morris H.; Gilinsky, Mikhail; Patel, Kaushal; Coston, Calvin; Blankson, Isaiah M.

    2003-01-01

    The research is focused on a wide regime of problems in the propulsion field as well as in experimental testing and theoretical and numerical simulation analyses for advanced aircraft and rocket engines. Results obtained are based on analytical methods, numerical simulations and experimental tests at the NASA LaRC and Hampton University computer complexes and experimental facilities. The main objective of this research is injection, mixing and combustion enhancement in propulsion systems. The sub-projects in the reporting period are: (A) Aero-performance and acoustics of Telescope-shaped designs. The work included a pylon set application for SCRAMJET. (B) An analysis of sharp-edged nozzle exit designs for effective fuel injection into the flow stream in air-breathing engines: triangular-round and diamond-round nozzles. (C) Measurement technique improvements for the HU Low Speed Wind Tunnel (HU LSWT) including an automatic data acquisition system and a two component (drag-lift) balance system. In addition, a course in the field of aerodynamics was developed for the teaching and training of HU students.

  13. Fiber reinforced superalloys for rocket engines

    NASA Technical Reports Server (NTRS)

    Petrasek, Donald W.; Stephens, Joseph R.

    1989-01-01

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

  14. Software Estimates Costs of Testing Rocket Engines

    NASA Technical Reports Server (NTRS)

    Smith, C. L.

    2003-01-01

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

  15. Fiber reinforced superalloys for rocket engines

    NASA Technical Reports Server (NTRS)

    Petrasek, Donald W.; Stephens, Joseph R.

    1988-01-01

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

  16. Performance of a 2D-CD nonaxisymmetric exhaust nozzle on a turbojet engine at altitude

    NASA Technical Reports Server (NTRS)

    Straight, D. M.; Cullom, R. R.

    1982-01-01

    Baseline thrust and cooling data obtained with a 2D-CD versatile research exhaust nozzle mounted on a turbojet engine in an altitude chamber are presented. The tests covered a range of nozzle pressure ratios, nozzle pressure ratios, nozzle throat areas, and internal expansion area ratios. The thrust data obtained show good agreement with theory and scale model results after correcting the data for leakage and bypass cooling flows. Additional work is needed to improve predictability of cooling performance.

  17. Problems of the mathematical description of rocket engines as plants

    NASA Astrophysics Data System (ADS)

    Kiforenko, B. N.

    2012-09-01

    Mathematical models of liquid-propellant, nuclear, and electric rocket engines are presented that more fully describe thrust generation than the classical models do. The optimal control of engine thrust is analyzed within the framework of Mayer's general variational problem. It is shown that the control of a rocket engine satisfying the necessary optimality conditions belongs to the boundary arc of the feasible control set between the point of maximum thrust and the point of maximum exhaust velocity

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

    NASA Technical Reports Server (NTRS)

    Linne, Diane L.

    1993-01-01

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

  19. Transpiration Cooled Throat for Hydrocarbon Rocket Engines

    DTIC Science & Technology

    1991-12-01

    system and was identified by parame- ter and sequence. 4.4 NOZZLE TESTING 4.4.1 -alg-reter Noizle Testiag 4.4.1.A Cold Flow Water Tests The...Coolant (RP-1) 31 4.3.7 Propellant Feed System 31 4.3.8 Instrumentation 33 4.4 Nozzle Testing 38 4.4.1 Calorimeter Nozzle Testing 38 4.4.2...Thermocouple Platelet Stack 24 13 Fabrication Cycle 28 14 Carbon Deposition Test Assembly Installed in Test Bay 30 15 Propellant System Schematic 32 16

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

    SciTech Connect

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

    1992-12-31

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

  1. Telemetry Boards Interpret Rocket, Airplane Engine Data

    NASA Technical Reports Server (NTRS)

    2009-01-01

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

  2. NARC Rayon Replacement Program for the Space Shuttle Reusable Solid Rocket Motor Nozzle: Screening Summary

    NASA Technical Reports Server (NTRS)

    Cook, R. V.; Fairbourn, M. W.; Wendel, G. M.

    2000-01-01

    Thiokol Corporation and NASA MSFC are jointly developing a replacement for North American Rayon Corporation (NARC) Aerospace Grade Rayon (1650/720 continuous filament), the precursor for the Carbon Cloth Phenolic (CCP) ablatives used in the Space Shuttle Reusable Solid Rocket Motor (RSRM) Nozzles. NARC discontinued production of Aerospace Grade Rayon in September 1997. NASA maintains a stockpile of NARC Rayon to support RSRM production through the summer of 2005. The program plan for selection and qualification of a replacement for NARC rayon was approved in August 1998. Screening activities began in February 1999. The intent of this paper is to provide a summary of the data generated during the screening phase of the NARC Rayon Replacement Program. Twelve cellulose based fibers (rayon and lyocell) were evaluated. These fibers were supplied by three independent vendors. Many of these fibers were carbonized by two independent carbonizers. Each candidate was tested according to standard acceptance test methods at each step of the manufacturing process. Additional testing was performed with the candidate CCPS, including hot fire tests, Process studies and mechanical and thermal characterization. Six of the twelve fiber candidates tested were dropped at the conclusion of Phase 1. The reasons for the elimination of these candidates included; difficulties in processing the material in the whitegoods, carbon and CCP forms; poor composite mechanical performance; and future availability concerns. The remaining six fibers demonstrated enough promise to merit continued evaluation and optimization of the CCP fabrication process. Note: Certain CCP data falls under the restrictions of US export laws, (ITAR, etc.) and will not be included in this paper.

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

    NASA Technical Reports Server (NTRS)

    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.

  4. A minimum cost tolerance allocation method for rocket engines and robust rocket engine design

    NASA Technical Reports Server (NTRS)

    Gerth, Richard J.

    1993-01-01

    Rocket engine design follows three phases: systems design, parameter design, and tolerance design. Systems design and parameter design are most effectively conducted in a concurrent engineering (CE) environment that utilize methods such as Quality Function Deployment and Taguchi methods. However, tolerance allocation remains an art driven by experience, handbooks, and rules of thumb. It was desirable to develop and optimization approach to tolerancing. The case study engine was the STME gas generator cycle. The design of the major components had been completed and the functional relationship between the component tolerances and system performance had been computed using the Generic Power Balance model. The system performance nominals (thrust, MR, and Isp) and tolerances were already specified, as were an initial set of component tolerances. However, the question was whether there existed an optimal combination of tolerances that would result in the minimum cost without any degradation in system performance.

  5. Acoustic investigation of the engine-over-the-wing concept using a D-shaped nozzle

    NASA Technical Reports Server (NTRS)

    Reshotko, M.; Friedman, R.

    1973-01-01

    Small-model experiments were conducted of the engine-over-the-wing concept using a D-shaped nozzle in order to determine the static-lift and acoustic characteristics at two wing-flap positions. Configurations were tested with the flow attached and unattached to the upper surface of the flaps. Attachment was obtained with a nozzle flow deflector. In both cases, high frequency noise shielding by the wing was obtained. Configurations using the D-shaped nozzle are compared with corresponding ones using a circular nozzle. With flow attached to the flaps, the static lift and acoustic results are almost the same for both nozzles. Without the nozzle flow deflector, (unattached flap flow), the D-nozzle is considerably noisier than a circular nozzle in the low and middle frequencies.

  6. Injector for liquid fueled rocket engine

    NASA Technical Reports Server (NTRS)

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

    2000-01-01

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

  7. High Torque Nozzle Development for the 2.75-Inch Rocket System

    DTIC Science & Technology

    2001-03-01

    significantly different from the NOSIH-AA-2 propellant used in the Mk66 motor. The Thiokol propellant was an AP/ HTPB /AI composite propellant . Metal Tab... IGNITER SALT ROD NOZZLE Replaced by High Torque Nozzle WARHEAD ADAPTER NSWC/IH-AA-2 PROPELLANT STABILIZING ROD WITH SHAPE MEMORY RING Figure 15. Test Motor...attributable to the nozzle. The fired motors were disassembled and visually inspected for anomalies such as hot spots , ruptures and bulges with no

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

    NASA Technical Reports Server (NTRS)

    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.

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

    NASA Astrophysics Data System (ADS)

    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.

  10. Analysis and Results from a Flush Airdata Sensing (FADS) System in Close Proximity to Firing Rocket Nozzles

    NASA Technical Reports Server (NTRS)

    Ali, Aliyah N.; Borrer, Jerry L.

    2013-01-01

    This presentation presents information regarding the nose-cap flush airdata sensing (FADS) system on Orion's Pad Abort 1 (PA-1) vehicle. The purpose of the nose-cap FADS system was to test whether or not useful data could be obtained from a FADS system if it was placed in close proximity to firing rockets nozzles like the attitude control motor (ACM) nozzles on the PA-1 launch abort system (LAS). The nose-cap FADS systems use pressure measurements from a series of pressure ports which are arranged in a cruciform pattern and flush with the surface of the vehicle to estimate values of angle of attack, angle of side-slip, Mach number, impact pressure and free-stream static pressure.

  11. High-speed schlieren imaging of rocket exhaust plumes

    NASA Astrophysics Data System (ADS)

    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.

  12. Rocketdyne/Westinghouse nuclear thermal rocket engine modeling

    NASA Technical Reports Server (NTRS)

    Glass, James F.

    1993-01-01

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

  13. Fiber optics in liquid propellant rocket engine environments

    NASA Technical Reports Server (NTRS)

    Delcher, R.; Dinnsen, D.; Barkhoudarian, S.

    1991-01-01

    Fiber optics have recently been seen to offer several major benefits in liquid-fuel rocket engine applications. Fiber-optic sensors can provide measurements that cannot be made with conventional techniques. Fiber optics also can reduce harness weight, provide lightning immunity, and increase frequency response. This paper discusses the results of feasibility testing optical fibers in simulated liquid-fuel rocket engine environments. The environments included cryogenic and high temperatures, and high vibration levels.

  14. Rocketdyne/Westinghouse nuclear thermal rocket engine modeling

    NASA Astrophysics Data System (ADS)

    Glass, James F.

    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.

  15. Preburner of Staged Combustion Rocket Engine

    NASA Technical Reports Server (NTRS)

    Yost, M. C.

    1978-01-01

    A regeneratively cooled LOX/hydrogen staged combustion assembly system with a 400:1 expansion area ratio nozzle utilizing an 89,000 Newton (20,000 pound) thrust regeneratively cooled thrust chamber and 175:1 tubular nozzle was analyzed, assembled, and tested. The components for this assembly include two spark/torch oxygen-hydrogen igniters, two servo-controlled LOX valves, a preburner injector, a preburner combustor, a main propellant injector, a regeneratively cooled combustion chamber, a regeneratively cooled tubular nozzle with an expansion area ratio of 175:1, an uncooled heavy-wall steel nozzle with an expansion area ratio of 400:1, and interconnecting ducting. The analytical effort was performed to optimize the thermal and structural characteristics of each of the new components and the ducting, and to reverify the capabilities of the previously fabricated components. The testing effort provided a demonstration of the preburner/combustor chamber operation, chamber combustion efficiency and stability, and chamber and nozzle heat transfer.

  16. Improvement of Rocket Engine Plume Analysis Techniques

    NASA Technical Reports Server (NTRS)

    Smith, S. D.

    1982-01-01

    A nozzle plume flow field code was developed. The RAMP code which was chosen as the basic code is of modular construction and has the following capabilities: two phase with two phase transonic solution; a two phase, reacting gas (chemical equilibrium reaction kinetics), supersonic inviscid nozzle/plume solution; and is operational for inviscid solutions at both high and low altitudes. The following capabilities were added to the code: a direct interface with JANNAF SPF code; shock capturing finite difference numerical operator; two phase, equilibrium/frozen, boundary layer analysis; a variable oxidizer to fuel ratio transonic solution; an improved two phase transonic solution; and a two phase real gas semiempirical nozzle boundary layer expansion.

  17. Nonlinear Longitudinal Mode Instability in Liquid Propellant Rocket Engine Preburners

    NASA Technical Reports Server (NTRS)

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

    2004-01-01

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

  18. Rocket U: Engineers Push New Limits

    NASA Image and Video Library

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

  19. Effect of Correlation on Multi-Engine Rocket Propulsion Systems

    NASA Technical Reports Server (NTRS)

    Baker, J. R.; Breneman, J. E.

    1989-01-01

    A matter of great concern in the design and operation of multi-engine rocket propulsion systems is the effect of the premature shutdown of one engine on the vehicle. This probability that a premature shutdown will cause a vehicle loss is termed correlation. Based on airbreathing experiences as well as rocket engine data the best estimate of this correlation is made and then applied to the overall multi-engine reliability problem to demonstrate its potential effect. At this point, follow-on analyses are pointed out that illustrate how any potential failures that may cause a correlatable event can be eliminated; thus bringing this correlation to almost 0.

  20. Stepped nozzle

    DOEpatents

    Sutton, G.P.

    1998-07-14

    An insert is described which allows a supersonic nozzle of a rocket propulsion system to operate at two or more different nozzle area ratios. This provides an improved vehicle flight performance or increased payload. The insert has significant advantages over existing devices for increasing nozzle area ratios. The insert is temporarily fastened by a simple retaining mechanism to the aft end of the diverging segment of the nozzle and provides for a multi-step variation of nozzle area ratio. When mounted in place, the insert provides the nozzle with a low nozzle area ratio. During flight, the retaining mechanism is released and the insert ejected thereby providing a high nozzle area ratio in the diverging nozzle segment. 5 figs.

  1. Stepped nozzle

    DOEpatents

    Sutton, George P.

    1998-01-01

    An insert which allows a supersonic nozzle of a rocket propulsion system to operate at two or more different nozzle area ratios. This provides an improved vehicle flight performance or increased payload. The insert has significant advantages over existing devices for increasing nozzle area ratios. The insert is temporarily fastened by a simple retaining mechanism to the aft end of the diverging segment of the nozzle and provides for a multi-step variation of nozzle area ratio. When mounted in place, the insert provides the nozzle with a low nozzle area ratio. During flight, the retaining mechanism is released and the insert ejected thereby providing a high nozzle area ratio in the diverging nozzle segment.

  2. Propellant-Flow-Actuated Rocket Engine Igniter

    NASA Technical Reports Server (NTRS)

    Wollen, Mark

    2013-01-01

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

  3. Full-Scale Investigation of Several Jet-Engine Noise-Reduction Nozzles

    NASA Technical Reports Server (NTRS)

    Coles, Willard D; Callaghan, Edmund E

    1958-01-01

    A number of noise-suppression nozzles were tested on full-scale engines. In general, these nozzles achieved noise reduction by the mixing interference of adjacent jets, that is, by using multiple-slot-nozzles. Several of the nozzles achieved reductions in sound power of approximately 5 decibels (nearly 70 percent) with small thrust losses (approx. 1 percent). The maximum sound-pressure level was reduced by as much as 18 decibels in particular frequency bands. Some of the nozzles showed considerable spatial asymmetry; that is, the sound field was not rotationally symmetrical. A method of calculating the limiting frequency effected by such nozzles is presented. Furthermore data are shown that appear to indicate that further reductions in sound power will not be easily achieved from nozzles using mixing interference as a means of noise suppression

  4. Ablative material testing for low-pressure, low-cost rocket engines

    NASA Technical Reports Server (NTRS)

    Richter, G. Paul; Smith, Timothy D.

    1995-01-01

    The results of an experimental evaluation of ablative materials suitable for the production of light weight, low cost rocket engine combustion chambers and nozzles are presented. Ten individual specimens of four different compositions of silica cloth-reinforced phenolic resin materials were evaluated for comparative erosion in a subscale rocket engine combustion chamber. Gaseous hydrogen and gaseous oxygen were used as propellants, operating at a nominal chamber pressure of 1138 kPa (165 psi) and a nominal mixture ratio (O/F) of 3.3. These conditions were used to thermally simulate operation with RP-1 and liquid oxygen, and achieved a specimen throat gas temperature of approximately 2456 K (4420 R). Two high-density composition materials exhibited high erosion resistance, while two low-density compositions exhibited approximately 6-75 times lower average erosion resistance. The results compare favorably with previous testing by NASA and provide adequate data for selection of ablatives for low pressure, low cost rocket engines.

  5. Space Shuttle main engine nozzle-steerhorn dynamics

    NASA Technical Reports Server (NTRS)

    Kiefling, L.

    1981-01-01

    On two occasions during the Space Shuttle main engine development, the LH2 feedline (called the steerhorn, because of its shape) failed during the cutoff transient. A dynamic test was undertaken, and an analytical model was developed and correlated to the dynamic test. Detailed models of the tube bundle were required to obtain the equivalent shell coefficients. All-shell models of the nozzle wall were found better than beam-shell models. The most difficult part of the structure to simulate was the felt-metal pad between the feedline and its mount, which introduced nonlinear stiffness and damping and led to the use of separate low amplitude and high amplitude models. The total structure was found to have 400 modes in the frequency range of interest, 0 to 500 Hz. Good test analysis correlation was obtained and a modified feedline configuration was found to demonstrate a 40% reduction of response stress from the original configuration.

  6. An Object Model for a Rocket Engine Numerical Simulator

    NASA Technical Reports Server (NTRS)

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

    1998-01-01

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

  7. An Object Model for a Rocket Engine Numerical Simulator

    NASA Technical Reports Server (NTRS)

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

    1998-01-01

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

  8. High-Area-Ratio Rocket Nozzle at High Combustion Chamber Pressure: Experimental and Analytical Validation

    NASA Technical Reports Server (NTRS)

    Jankovsky, Robert S.; Smith, Timothy D.; Pavli, Albert J.

    1999-01-01

    Experimental data were obtained on an optimally contoured nozzle with an area ratio of 1025:1 and on a truncated version of this nozzle with an area ratio of 440:1. The nozzles were tested with gaseous hydrogen and liquid oxygen propellants at combustion chamber pressures of 1800 to 2400 psia and mixture ratios of 3.89 to 6.15. This report compares the experimental performance, heat transfer, and boundary layer total pressure measurements with theoretical predictions of the current Joint Army, Navy, NASA, Air Force (JANNAF) developed methodology. This methodology makes use of the Two-Dimensional Kinetics (TDK) nozzle performance code. Comparisons of the TDK-predicted performance to experimentally attained thrust performance indicated that both the vacuum thrust coefficient and the vacuum specific impulse values were approximately 2.0-percent higher than the turbulent prediction for the 1025:1 configurations, and approximately 0.25-percent higher than the turbulent prediction for the 440:1 configuration. Nozzle wall temperatures were measured on the outside of a thin-walled heat sink nozzle during the test fittings. Nozzle heat fluxes were calculated front the time histories of these temperatures and compared with predictions made with the TDK code. The heat flux values were overpredicted for all cases. The results range from nearly 100 percent at an area ratio of 50 to only approximately 3 percent at an area ratio of 975. Values of the integral of the heat flux as a function of nozzle surface area were also calculated. Comparisons of the experiment with analyses of the heat flux and the heat rate per axial length also show that the experimental values were lower than the predicted value. Three boundary layer rakes mounted on the nozzle exit were used for boundary layer measurements. This arrangement allowed total pressure measurements to be obtained at 14 different distances from the nozzle wall. A comparison of boundary layer total pressure profiles and analytical

  9. Time-Frequency Analysis of Rocket Nozzle Wall Pressures During Start-up Transients

    NASA Technical Reports Server (NTRS)

    Baars, Woutijn J.; Tinney, Charles E.; Ruf, Joseph H.

    2011-01-01

    Surveys of the fluctuating wall pressure were conducted on a sub-scale, thrust- optimized parabolic nozzle in order to develop a physical intuition for its Fourier-azimuthal mode behavior during fixed and transient start-up conditions. These unsteady signatures are driven by shock wave turbulent boundary layer interactions which depend on the nozzle pressure ratio and nozzle geometry. The focus however, is on the degree of similarity between the spectral footprints of these modes obtained from transient start-ups as opposed to a sequence of fixed nozzle pressure ratio conditions. For the latter, statistically converged spectra are computed using conventional Fourier analyses techniques, whereas the former are investigated by way of time-frequency analysis. The findings suggest that at low nozzle pressure ratios -- where the flow resides in a Free Shock Separation state -- strong spectral similarities occur between fixed and transient conditions. Conversely, at higher nozzle pressure ratios -- where the flow resides in Restricted Shock Separation -- stark differences are observed between the fixed and transient conditions and depends greatly on the ramping rate of the transient period. And so, it appears that an understanding of the dynamics during transient start-up conditions cannot be furnished by a way of fixed flow analysis.

  10. Comparison of theoretical and experimental thrust performance of a 1030:1 area ratio rocket nozzle at a chamber pressure of 2413 kN/m2 (350 psia)

    NASA Technical Reports Server (NTRS)

    Smith, Tamara A.; Pavli, Albert J.; Kacynski, Kenneth J.

    1987-01-01

    The joint Army. Navy, NASA. Air Force (JANNAF) rocket engine peformnace prediction procedure is based on the use of various reference computer programs. One of the reference programs for nozzle analysis is the Two-Dimensional Kinetics (TDK) Program. The purpose of this report is to calibrate the JANNAF procedure incorporated into the December l984 version of the TDK program for the high-area-ratio rocket engine regime. The calibration was accomplished by modeling the performance of a 1030:1 rocket nozzle tested at NASA Lewis Research Center. A detailed description of the experimental test conditions and TDK input parameters is given. The results show that the computer code predicts delivered vacuum specific impulse to within 0.12 to 1.9 percent of the experimental data. Vacuum thrust coefficient predictions were within + or - 1.3 percent of experimental results. Predictions of wall static pressure were within approximately + or - 5 percent of the measured values. An experimental value for inviscid thrust was obtained for the nozzle extension between area ratios of 427.5 and 1030 by using an integration of the measured wall static pressures. Subtracting the measured thrust gain produced by the nozzle between area ratios of 427.5 and 1030 from the inviscid thrust gain yielded experimental drag decrements of 10.85 and 27.00 N (2.44 and 6.07 lb) for mixture ratios of 3.04 and 4.29, respectively. These values correspond to 0.45 and 1.11 percent of the total vacuum thrust. At a mixture ratio of 4.29, the TDK predicted drag decrement was 16.59 N (3.73 lb), or 0.71 percent of the predicted total vacuum thrust.

  11. Some unusual oscillating bearing applications in liquid propellant rocket engines

    NASA Technical Reports Server (NTRS)

    Daniels, C. M.

    1975-01-01

    Liquid propellant rocket engines have some unique applications for highly loaded oscillating-motion bearings operating at cryogenic temperatures and vacuum environment conditions. Among these are (1) thrust attitude control gimbal bearings that permit pivoting of the entire rocket engine and (2) bearings in the linkages of ducting bellows joints that articulate when the engine gimbals. The main rocket engines for first, second, and third stages of the Saturn vehicle have thrust vector control gimbals which use PTFE liner bearing materials and ducting linkage bearings using solid film dry lubricants, both to reduce friction and provide required wear life. Data on the frictional and wear characteristics of these materials, while operating under varying temperature and atmospheric pressure conditions, are presented. A discussion of the bearing designs for flexible bellows joint linkages is also presented.

  12. Reusable rocket engine intelligent control system framework design, phase 2

    NASA Technical Reports Server (NTRS)

    Nemeth, ED; Anderson, Ron; Ols, Joe; Olsasky, Mark

    1991-01-01

    Elements of an advanced functional framework for reusable rocket engine propulsion system control are presented for the Space Shuttle Main Engine (SSME) demonstration case. Functional elements of the baseline functional framework are defined in detail. The SSME failure modes are evaluated and specific failure modes identified for inclusion in the advanced functional framework diagnostic system. Active control of the SSME start transient is investigated, leading to the identification of a promising approach to mitigating start transient excursions. Key elements of the functional framework are simulated and demonstration cases are provided. Finally, the advanced function framework for control of reusable rocket engines is presented.

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

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

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

  14. Blood Pump Development Using Rocket Engine Flow Simulation Technology

    NASA Technical Reports Server (NTRS)

    Kiris, Cetin C.; Kwak, Dochan

    2002-01-01

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

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

    NASA Technical Reports Server (NTRS)

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

    2001-01-01

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

  16. Transmittance and Radiance Computations for Rocket Engine Plume Environments

    NASA Technical Reports Server (NTRS)

    Tejwani, Gopal D.

    2003-01-01

    Emission and absorption characteristics of several atmospheric and combustion species have been studied and are presented with reference to rocket engine plume environments. The effects of clous, rain, and fog on plume radiance/transmittance has also been studied.Preliminary results for the radiance from the exhaust plume of the space shuttle main engine are shown and discussed.

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

    NASA Technical Reports Server (NTRS)

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

    1998-01-01

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

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

    NASA Astrophysics Data System (ADS)

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

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

  19. Interior flow and near-nozzle spray development in a marine-engine diesel fuel injector

    NASA Astrophysics Data System (ADS)

    Hult, J.; Simmank, P.; Matlok, S.; Mayer, S.; Falgout, Z.; Linne, M.

    2016-04-01

    A consolidated effort at optically characterising flow patterns, in-nozzle cavitation, and near-nozzle jet structure of a marine diesel fuel injector is presented. A combination of several optical techniques was employed to fully transparent injector models, compound metal-glass and full metal injectors. They were all based on a common real-scale dual nozzle hole geometry for a marine two-stroke diesel engine. In a stationary flow rig, flow velocities in the sac-volume and nozzle holes were measured using PIV, and in-nozzle cavitation visualized using high-resolution shadowgraphs. The effect of varying cavitation number was studied and results compared to CFD predictions. In-nozzle cavitation and near-nozzle jet structure during transient operation were visualized simultaneously, using high-speed imaging in an atmospheric pressure spray rig. Near-nozzle spray formation was investigated using ballistic imaging. Finally, the injector geometry was tested on a full-scale marine diesel engine, where the dynamics of near-nozzle jet development was visualized using high-speed shadowgraphy. The range of studies focused on a single common geometry allows a comprehensive survey of phenomena ranging from first inception of cavitation under well-controlled flow conditions to fuel jet structure at real engine conditions.

  20. Advanced oxygen-hydrocarbon rocket engine study

    NASA Technical Reports Server (NTRS)

    Obrien, C. J.

    1979-01-01

    A consistent engine system data base was generated for defining advantages and disadvantages, system performance and operating limits, engine parametric data, and technology requirements for candidate high pressure LO2/HC engine systems. Optimum LO2/HC engine power cycles were synthetized and representative conceptual engine design generated for a specified advanced surface to orbit transportation system.

  1. Experimental evaluation of heat transfer on a 1030:1 area ratio rocket nozzle

    NASA Technical Reports Server (NTRS)

    Kacynski, Kenneth J.; Pavli, Albert J.; Smith, Tamara A.

    1987-01-01

    A 1030:1 carbon steel, heat-sink nozzle was tested. The test conditions included a nominal chamber pressure of 2413 kN/sq m and a mixture ratio range of 2.78 to 5.49. The propellants were gaseous oxygen and gaseous hydrogen. Outer wall temperature measurements were used to calculate the inner wall temperature and the heat flux and heat rate to the nozzle at specified axial locations. The experimental heat fluxes were compared to those predicted by the Two-Dimensional Kinetics (TDK) computer model analysis program. When laminar boundary layer flow was assumed in the analysis, the predicted values were within 15% of the experimental values for the area ratios of 20 to 975. However, when turbulent boundary layer conditions were assumed, the predicted values were approximately 120% higher than the experimental values. A study was performed to determine if the conditions within the nozzle could sustain a laminar boundary layer. Using the flow properties predicted by TDK, the momentum-thickness Reynolds number was calculated, and the point of transition to turbulent flow was predicted. The predicted transition point was within 0.5 inches of the nozzle throat. Calculations of the acceleration parameter were then made to determine if the flow conditions could produce relaminarization of the boundary layer. It was determined that if the boundary layer flow was inclined to transition to turbulent, the acceleration conditions within the nozzle would tend to suppress turbulence and keep the flow laminar-like.

  2. Computational Fluid Dynamics Modeling of a Supersonic Nozzle and Integration into a Variable Cycle Engine Model

    NASA Technical Reports Server (NTRS)

    Connolly, Joseph W.; Friedlander, David; Kopasakis, George

    2014-01-01

    This paper covers the development of an integrated nonlinear dynamic simulation for a variable cycle turbofan engine and nozzle that can be integrated with an overall vehicle Aero-Propulso-Servo-Elastic (APSE) model. A previously developed variable cycle turbofan engine model is used for this study and is enhanced here to include variable guide vanes allowing for operation across the supersonic flight regime. The primary focus of this study is to improve the fidelity of the model's thrust response by replacing the simple choked flow equation convergent-divergent nozzle model with a MacCormack method based quasi-1D model. The dynamic response of the nozzle model using the MacCormack method is verified by comparing it against a model of the nozzle using the conservation element/solution element method. A methodology is also presented for the integration of the MacCormack nozzle model with the variable cycle engine.

  3. Computational Fluid Dynamics Modeling of a Supersonic Nozzle and Integration into a Variable Cycle Engine Model

    NASA Technical Reports Server (NTRS)

    Connolly, Joseph W.; Friedlander, David; Kopasakis, George

    2015-01-01

    This paper covers the development of an integrated nonlinear dynamic simulation for a variable cycle turbofan engine and nozzle that can be integrated with an overall vehicle Aero-Propulso-Servo-Elastic (APSE) model. A previously developed variable cycle turbofan engine model is used for this study and is enhanced here to include variable guide vanes allowing for operation across the supersonic flight regime. The primary focus of this study is to improve the fidelity of the model's thrust response by replacing the simple choked flow equation convergent-divergent nozzle model with a MacCormack method based quasi-1D model. The dynamic response of the nozzle model using the MacCormack method is verified by comparing it against a model of the nozzle using the conservation element/solution element method. A methodology is also presented for the integration of the MacCormack nozzle model with the variable cycle engine.

  4. Advanced oxygen-hydrocarbon rocket engine study

    NASA Technical Reports Server (NTRS)

    Obrien, C. J.; Salkeld, R.

    1980-01-01

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

  5. Mixer nozzle aeroacoustic characteristics for the energy efficient engine

    NASA Technical Reports Server (NTRS)

    Gliebe, P. R.; Sandusky, G. T.; Chamberlin, R.

    1981-01-01

    Aeroacoustic tests are conducted on scale model mixer nozzle configurations, a separate flow nozzle, and a baseline conical nozzle in an anechoic chamber free-jet facility to investigate exhaust system designs. Far-field acoustic data and exhaust jet plume aerodynamic data are obtained for various combinations of exhaust jet velocity and simulated flight speed, with corrected microphone data at a constant 12.2 m arc distance, and scaled acoustic data at a sideline distance of 457 m. It is found that jet plume aerodynamic and acoustic characteristics of lobed mixer exhaust systems are similar to those of a conical nozzle operating at the same specific thrust, although differences occur at high frequencies, where the sound pressure level of the mixer nozzle is 1-5 dB higher than that of a conical nozzle. In addition, no direct correlation is found between exit plane turbulence levels and plume development or acoustic characteristics for mixer exhaust configurations.

  6. Use of Soft Computing Technologies For Rocket Engine Control

    NASA Technical Reports Server (NTRS)

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

    2003-01-01

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

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

    NASA Technical Reports Server (NTRS)

    2003-01-01

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

  8. IR signature study of aircraft engine for variation in nozzle exit area

    NASA Astrophysics Data System (ADS)

    Baranwal, Nidhi; Mahulikar, Shripad P.

    2016-01-01

    In general, jet engines operate with choked nozzle during take-off, climb and cruise, whereas unchoking occurs while landing and taxiing (when engine is not running at full power). Appropriate thrust in an aircraft in all stages of the flight, i.e., take-off, climb, cruise, descent and landing is achieved through variation in the nozzle exit area. This paper describes the effect on thrust and IR radiance of a turbojet engine due to variation in the exit area of a just choked converging nozzle (Me = 1). The variations in the nozzle exit area result in either choking or unchoking of a just choked converging nozzle. Results for the change in nozzle exit area are analyzed in terms of thrust, mass flow rate and specific fuel consumption. The solid angle subtended (Ω) by the exhaust system is estimated analytically, for the variation in nozzle exit area (Ane), as it affects the visibility of the hot engine parts from the rear aspect. For constant design point thrust, IR radiance is studied from the boresight (ϕ = 0°, directly from the rear side) for various percentage changes in nozzle exit area (%ΔAne), in the 1.9-2.9 μm and 3-5 μm bands.

  9. Investigations of Rocket Engine Combustion Emissions During ACCENT

    NASA Astrophysics Data System (ADS)

    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.

  10. Effect of nozzle orifice geometry on spray, combustion, and emission characteristics under diesel engine conditions.

    SciTech Connect

    Som, S.; Longman, D. E; Ramirez, A. I.; Aggarwal, S. K.

    2011-03-01

    Diesel engine performance and emissions are strongly coupled with fuel atomization and spray processes, which in turn are strongly influenced by injector flow dynamics. Modern engines employ micro-orifices with different orifice designs. It is critical to characterize the effects of various designs on engine performance and emissions. In this study, a recently developed primary breakup model (KH-ACT), which accounts for the effects of cavitation and turbulence generated inside the injector nozzle is incorporated into a CFD software CONVERGE for comprehensive engine simulations. The effects of orifice geometry on inner nozzle flow, spray, and combustion processes are examined by coupling the injector flow and spray simulations. Results indicate that conicity and hydrogrinding reduce cavitation and turbulence inside the nozzle orifice, which slows down primary breakup, increasing spray penetration, and reducing dispersion. Consequently, with conical and hydroground nozzles, the vaporization rate and fuel air mixing are reduced, and ignition occurs further downstream. The flame lift-off lengths are the highest and lowest for the hydroground and conical nozzles, respectively. This can be related to the rate of fuel injection, which is higher for the hydroground nozzle, leading to richer mixtures and lower flame base speeds. A modified flame index is employed to resolve the flame structure, which indicates a dual combustion mode. For the conical nozzle, the relative role of rich premixed combustion is enhanced and that of diffusion combustion reduced compared to the other two nozzles. In contrast, for the hydroground nozzle, the role of rich premixed combustion is reduced and that of non-premixed combustion is enhanced. Consequently, the amount of soot produced is the highest for the conical nozzle, while the amount of NOx produced is the highest for the hydroground nozzle, indicating the classical tradeoff between them.

  11. Development and Test of a Rocket Engine Using Environmentally Friendly Propellants

    NASA Technical Reports Server (NTRS)

    Webster, Kristi

    2009-01-01

    Develop and test a rocket engine that operates on environmentally friendly propellants; Liquid Oxygen (LOX) and Liquid Methane (LCH4). Due to modifications the rocket engine designed last summer (KJ_REX) is not the same rocket thruster tested this summer, but very similar. The new modified rocket thruster was built for NASA by Orion Propulsion Inc. (OPI), Huntsville, AL.

  12. Grooved Fuel Rings for Nuclear Thermal Rocket Engines

    NASA Technical Reports Server (NTRS)

    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.

  13. Advanced solid rocket motor nozzle thrust vector control flexseal development status

    NASA Astrophysics Data System (ADS)

    Prins, William S.; Meyer, Scott A.; Cox, Paul D.

    1992-07-01

    The advanced solid rocket motor (ASRM) flexseal development status is reviewed focusing on design goals and requirements, design configuration, analysis activities, and verification tests. It is concluded that the ASRM flexseal incorporates flight-proven materials in an innovative design configuration. Variable thickness shims and efficient packaging of the flexseal make it possible to achieve a significant weight reduction. A flexseal insulator design derived from strategic solid rocket motor experience will provide the necessary bearing thermal protection while minimizing vectoring torque variability.

  14. 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

    NASA Astrophysics Data System (ADS)

    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.

  15. High altitude chemically reacting gas particle mixtures. Volume 2: Program manual for RAMP2. [rocket nozzle and orbital plume flow fields

    NASA Technical Reports Server (NTRS)

    Smith, S. D.

    1984-01-01

    All of the elements used in the Reacting and Multi-Phase (RAMP2) computer code are described in detail. The code can be used to model the dominant phenomena which affect the prediction of liquid and solid rocket nozzle and orbital plume flow fields.

  16. Thermal hydraulic design analysis of ternary carbide fueled square-lattice honeycomb nuclear rocket engine

    SciTech Connect

    Furman, Eric M.; Anghaie, Samim

    1999-01-22

    A computational analysis is conducted to determine the optimum thermal-hydraulic design parameters for a square-lattice honeycomb nuclear rocket engine core that will incorporate ternary carbide based uranium fuels. Recent studies at the Innovative Nuclear Space Power and Propulsion Institute (INSPI) have demonstrated the feasibility of processing solid solution, ternary carbide fuels such as (U, Zr, Nb)C, (U, Zr, Ta)C, (U, Zr, Hf)C and (U, Zr, W)C. The square-lattice honeycomb design provides high strength and is amenable to the processing complexities of these ultrahigh temperature fuels. A parametric analysis is conducted to examine how core geometry, fuel thickness and the propellant flow area effect the thermal performance of the nuclear rocket engine. The principal variables include core size (length and diameter) and fuel element dimensions. The optimum core configuration requires a balance between high specific impulse and thrust level performance, and maintaining the temperature and strength limits of the fuel. A nuclear rocket engine simulation code is developed and used to examine the system performance as well as the performance of the main reactor core components. The system simulation code was originally developed for analysis of NERVA-Derivative and Pratt and Whitney XNR-2000 nuclear thermal rockets. The code is modified and adopted to the square-lattice geometry of the new fuel design. Thrust levels ranging from 44,500 to 222,400 N (10,000 to 50,000 lbf) are considered. The average hydrogen exit temperature is kept at 2800 K, which is well below the melting point of these fuels. For a nozzle area ratio of 300 and a thrust chamber pressure of 4.8 Mpa (700 psi), the specific impulse is 930 s. Hydrogen temperature and pressure distributions in the core and the fuel maximum temperatures are calculated.

  17. Thermal hydraulic design analysis of ternary carbide fueled square-lattice honeycomb nuclear rocket engine

    NASA Astrophysics Data System (ADS)

    Furman, Eric M.; Anghaie, Samim

    1999-01-01

    A computational analysis is conducted to determine the optimum thermal-hydraulic design parameters for a square-lattice honeycomb nuclear rocket engine core that will incorporate ternary carbide based uranium fuels. Recent studies at the Innovative Nuclear Space Power and Propulsion Institute (INSPI) have demonstrated the feasibility of processing solid solution, ternary carbide fuels such as (U, Zr, Nb)C, (U, Zr, Ta)C, (U, Zr, Hf)C and (U, Zr, W)C. The square-lattice honeycomb design provides high strength and is amenable to the processing complexities of these ultrahigh temperature fuels. A parametric analysis is conducted to examine how core geometry, fuel thickness and the propellant flow area effect the thermal performance of the nuclear rocket engine. The principal variables include core size (length and diameter) and fuel element dimensions. The optimum core configuration requires a balance between high specific impulse and thrust level performance, and maintaining the temperature and strength limits of the fuel. A nuclear rocket engine simulation code is developed and used to examine the system performance as well as the performance of the main reactor core components. The system simulation code was originally developed for analysis of NERVA-Derivative and Pratt & Whitney XNR-2000 nuclear thermal rockets. The code is modified and adopted to the square-lattice geometry of the new fuel design. Thrust levels ranging from 44,500 to 222,400 N (10,000 to 50,000 lbf) are considered. The average hydrogen exit temperature is kept at 2800 K, which is well below the melting point of these fuels. For a nozzle area ratio of 300 and a thrust chamber pressure of 4.8 Mpa (700 psi), the specific impulse is 930 s. Hydrogen temperature and pressure distributions in the core and the fuel maximum temperatures are calculated.

  18. Development of low cost fabrication techniques for large solid rocket nozzles

    NASA Technical Reports Server (NTRS)

    Warga, J. J.

    1971-01-01

    Property measurements and fabrication characteristics were determined and the performance in subscale (Minuteman Wing 2 second stage) motors was evaluated. It was demonstrated that the incorporation of low cost fabrication techniques in a full scale 260 in. nozzle could result in savings of $149,000 when compared with an identical design using tape-wrapped components throughout.

  19. Advanced Methods for Aircraft Engine Thrust and Noise Benefits: Nozzle-Inlet Flow Analysis

    NASA Technical Reports Server (NTRS)

    Morgan, Morris H.; Gilinsky, Mikhail M.

    2001-01-01

    Three connected sub-projects were conducted under reported project. Partially, these sub-projects are directed to solving the problems conducted by the HU/FM&AL under two other NASA grants. The fundamental idea uniting these projects is to use untraditional 3D corrugated nozzle designs and additional methods for exhaust jet noise reduction without essential thrust lost and even with thrust augmentation. Such additional approaches are: (1) to add some solid, fluid, or gas mass at discrete locations to the main supersonic gas stream to minimize the negative influence of strong shock waves forming in propulsion systems; this mass addition may be accompanied by heat addition to the main stream as a result of the fuel combustion or by cooling of this stream as a result of the liquid mass evaporation and boiling; (2) to use porous or permeable nozzles and additional shells at the nozzle exit for preliminary cooling of exhaust hot jet and pressure compensation for non-design conditions (so-called continuous ejector with small mass flow rate; and (3) to propose and analyze new effective methods fuel injection into flow stream in air-breathing engines. Note that all these problems were formulated based on detailed descriptions of the main experimental facts observed at NASA Glenn Research Center. Basically, the HU/FM&AL Team has been involved in joint research with the purpose of finding theoretical explanations for experimental facts and the creation of the accurate numerical simulation technique and prediction theory for solutions for current problems in propulsion systems solved by NASA and Navy agencies. The research is focused on a wide regime of problems in the propulsion field as well as in experimental testing and theoretical and numerical simulation analysis for advanced aircraft and rocket engines. The F&AL Team uses analytical methods, numerical simulations, and possible experimental tests at the Hampton University campus. We will present some management activity

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

    NASA Technical Reports Server (NTRS)

    Veres, Joseph P.; Lavelle, Thomas M.

    2000-01-01

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

  1. Advanced subsonic Technology Noise Reduction Element Separate Flow Nozzle Tests for Engine Noise Reduction Sub-Element

    NASA Technical Reports Server (NTRS)

    Saiyed, Naseem H.

    2000-01-01

    Contents of this presentation include: Advanced Subsonic Technology (AST) goals and general information; Nozzle nomenclature; Nozzle schematics; Photograph of all baselines; Configurations tests and types of data acquired; and Engine cycle and plug geometry impact on EPNL.

  2. Parametric Model of an Aerospike Rocket Engine

    NASA Technical Reports Server (NTRS)

    Korte, J. J.

    2000-01-01

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

  3. Parametric Model of an Aerospike Rocket Engine

    NASA Technical Reports Server (NTRS)

    Korte, J. J.

    2000-01-01

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

  4. Thermal Analysis of a Carbon Fiber Rope Barrier for Use in the Reusable Solid Rocket Motor Nozzle Joint-2

    NASA Technical Reports Server (NTRS)

    Clayton, J. Louie; Phelps, Lisa (Technical Monitor)

    2001-01-01

    This study provides for development and verification of analysis methods used to assess performance of a carbon fiber rope (CFR) thermal barrier system that is currently being qualified for use in Reusable Solid Rocket Motor (RSRM) nozzle joint-2. Modeled geometry for flow calculations considers the joint to be vented with the porous CFR barriers placed in the "open' assembly gap. Model development is based on a 1-D volume filling approach where flow resistances (assembly gap and CFRs) are defined by serially connected internal flow and the porous media "Darcy" relationships. Combustion gas flow rates are computed using the volume filling code by assuming a lumped distribution total joint fill volume on a per linear circumferential inch basis. Gas compressibility, friction and heat transfer are included in the modeling. Gas-to-wall heat transfer is simulated by concurrent solution of the compressible flow equations and a large thermal 2-D finite element (FE) conduction grid. The derived numerical technique loosely couples the FE conduction matrix with the compressible gas flow equations, Free constants that appear in the governing equations are calibrated by parametric model comparison to hot fire subscale test results. The calibrated model is then used to make full-scale motor predictions using RSRM aft dome environments. Model results indicate that CFR thermal barrier systems will provide a thermally benign and controlled pressurization environment for the RSRM nozzle joint-2 primary seal activation.

  5. Thermal Analysis of a Carbon Fiber Rope Barrier for Use in the Reusable Solid Rocket Motor Nozzle Joint-2

    NASA Technical Reports Server (NTRS)

    Clayton, J. Louie; Phelps, Lisa (Technical Monitor)

    2001-01-01

    This study provides for development and verification of analysis methods used to assess performance of a carbon fiber rope (CFR) thermal barrier system that is currently being qualified for use in Reusable Solid Rocket Motor (RSRM) nozzle joint-2. Modeled geometry for flow calculations considers the joint to be vented with the porous CFR barriers placed in the "open' assembly gap. Model development is based on a 1-D volume filling approach where flow resistances (assembly gap and CFRs) are defined by serially connected internal flow and the porous media "Darcy" relationships. Combustion gas flow rates are computed using the volume filling code by assuming a lumped distribution total joint fill volume on a per linear circumferential inch basis. Gas compressibility, friction and heat transfer are included in the modeling. Gas-to-wall heat transfer is simulated by concurrent solution of the compressible flow equations and a large thermal 2-D finite element (FE) conduction grid. The derived numerical technique loosely couples the FE conduction matrix with the compressible gas flow equations, Free constants that appear in the governing equations are calibrated by parametric model comparison to hot fire subscale test results. The calibrated model is then used to make full-scale motor predictions using RSRM aft dome environments. Model results indicate that CFR thermal barrier systems will provide a thermally benign and controlled pressurization environment for the RSRM nozzle joint-2 primary seal activation.

  6. Thermal Analysis of a Carbon Fiber Rope Barrier for Use in the Reusable Solid Rocket Motor Nozzle Joint-2

    NASA Technical Reports Server (NTRS)

    Clayton, J. Louie

    2002-01-01

    This study provides development and verification of analysis methods used to assess performance of a carbon fiber rope (CFR) thermal barrier system that is currently being qualified for use in Reusable Solid Rocket Motor (RSRM) nozzle joint-2. Modeled geometry for flow calculations considers the joint to be vented with the porous CFR barriers placed in the 'open' assembly gap. Model development is based on a 1-D volume filling approach where flow resistances (assembly gap and CFRs) are defined by serially connected internal flow and the porous media 'Darcy' relationships. Combustion gas flow rates are computed using the volume filling code by assuming a lumped distribution total joint fill volume on a per linear circumferential inch basis. Gas compressibility, friction and heat transfer are included in the modeling. Gas-to-wall heat transfer is simulated by concurrent solution of the compressible flow equations and a large thermal 2-D finite element (FE) conduction grid. The derived numerical technique loosely couples the FE conduction matrix with the compressible gas flow equations. Free constants that appear in the governing equations are calibrated by parametric model comparison to hot fire subscale test results. The calibrated model is then used to make full-scale motor predictions using RSRM aft dome environments. Model results indicate that CFR thermal barrier systems will provide a thermally benign and controlled pressurization environment for the RSRM nozzle joint-2 primary seal activation.

  7. Rocket engine heat transfer and material technology for commercial applications

    NASA Technical Reports Server (NTRS)

    Hiltabiddle, J.; Campbell, J.

    1974-01-01

    Liquid fueled rocket engine combustion, heat transfer, and material technology have been utilized in the design and development of compact combustion and heat exchange equipment intended for application in the commercial field. An initial application of the concepts to the design of a compact steam generator to be utilized by electrical utilities for the production of peaking power is described.

  8. Rocket engine heat transfer and material technology for commercial applications

    NASA Technical Reports Server (NTRS)

    Hiltabiddle, J.; Campbell, J.

    1974-01-01

    Liquid fueled rocket engine combustion, heat transfer, and material technology have been utilized in the design and development of compact combustion and heat exchange equipment intended for application in the commercial field. An initial application of the concepts to the design of a compact steam generator to be utilized by electrical utilities for the production of peaking power is described.

  9. Cryostatless high temperature supercurrent bearings for rocket engine turbopumps

    NASA Technical Reports Server (NTRS)

    Rao, Dantam K.; Dill, James F.

    1989-01-01

    The rocket engine systems examined include SSME, ALS, and CTV systems. The liquid hydrogen turbopumps in the SSME and ALS vehicle systems are identified as potentially attractive candidates for development of Supercurrent Bearings since the temperatures around the bearings is about 30 K, which is considerably lower than the 95 K transition temperatures of HTS materials. At these temperatures, the current HTS materials are shown to be capable of developing significantly higher current densities. This higher current density capability makes the development of supercurrent bearings for rocket engines an attractive proposition. These supercurrent bearings are also shown to offer significant advantages over conventional bearings used in rocket engines. They can increase the life and reliability over rolling element bearings because of noncontact operation. They offer lower power loss over conventional fluid film bearings. Compared to conventional magnetic bearings, they can reduce the weight of controllers significantly, and require lower power because of the use of persistent currents. In addition, four technology areas that require further attention have been identified. These are: Supercurrent Bearing Conceptual Design Verification; HTS Magnet Fabrication and Testing; Cryosensors and Controller Development; and Rocket Engine Environmental Compatibility Testing.

  10. Machine learning of fault characteristics from rocket engine simulation data

    NASA Technical Reports Server (NTRS)

    Ke, Min; Ali, Moonis

    1990-01-01

    Transformation of data into knowledge through conceptual induction has been the focus of our research described in this paper. We have developed a Machine Learning System (MLS) to analyze the rocket engine simulation data. MLS can provide to its users fault analysis, characteristics, and conceptual descriptions of faults, and the relationships of attributes and sensors. All the results are critically important in identifying faults.

  11. Selecting and using materials for a nuclear rocket engine reactor

    NASA Astrophysics Data System (ADS)

    Lanin, Anatolii G.; Fedik, Ivan I.

    2011-03-01

    This paper provides a historical account of how the nuclear rocket engine reactor was created and discusses the problem of selecting materials for a gas environment at a temperature of up to 3100 K and energy release of 30 MW per liter.

  12. Liquid rocket engine fluid-cooled combustion chambers

    NASA Technical Reports Server (NTRS)

    1972-01-01

    A monograph on the design and development of fluid cooled combustion chambers for liquid propellant rocket engines is presented. The subjects discussed are (1) regenerative cooling, (2) transpiration cooling, (3) film cooling, (4) structural analysis, (5) chamber reinforcement, and (6) operational problems.

  13. Pretest uncertainty analysis for chemical rocket engine tests

    NASA Technical Reports Server (NTRS)

    Davidian, Kenneth J.

    1987-01-01

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

  14. Dual-mode Operation of a Rocket-Ramjet Combined Cycle Engine

    NASA Astrophysics Data System (ADS)

    Tomioka, Sadatake; Tani, Koichiro; Masumoto, Ryo; Ueda, Shuuichi

    One-dimensional evaluation of Ramjet-mode operation was carried out on a rocket-ramjet combined cycle engine model. For simplicity, instantaneous mixing between the airflow and rocket exhaust, instantaneous heat release, and pressure recovery by a normal-shock wave were assumed. Shock wave location was so decided that the heat release at the injection (heat addition) location was to thermally-choke the combustion gas flow. By changing the injection location, it was shown that a further downstream injection resulted in a further thrust production and a further fuel flow rate requirement for choking, and a lesser specific impulse. Balancing the thrust production and the specific impulse in terms of the launch vehicle acceleration performance should be pursued. The total pressure loss within the engine model was dominated by the shock wave location, not depended on injection location and fuel flow rate, so that having shock wave penetration to further upstream location was beneficial both for thrust production in the engine and at the external nozzle.

  15. Development of carbon-carbon nozzle extension for liquid fuel rocket motors

    NASA Astrophysics Data System (ADS)

    Sokolovsky, M. I.; Petukhov, S. N.; Semyonov, Yu. P.; Sokolov, B. A.

    2008-12-01

    Successful experience of RSC “Energy” and SPA “Iskra” in the development of carbon-carbon extension for oxygen-kerosene liquid fuel rocket motor has been summarized. Methodological approach that served to completion of carbon-carbon extension development in full and at comparatively small expenses has been described. Results of practical application of carbon-carbon extension for liquid fuel rocket motor 11D58M have been presented within the framework of International Space Program “Sea Launch”.

  16. Coefficients of discharge of fuel-injection nozzles for compression-ignition engines

    NASA Technical Reports Server (NTRS)

    Gelalles, A G

    1932-01-01

    This report presents the results of an investigation to determine the coefficients of discharge of nozzles with small, round orifices of the size used with high-speed compression-ignition engines. The injection pressures and chamber back pressures employed were comparable to those existing in compression-ignition engines during injection. The construction of the nozzles was varied to determine the effect of the nozzle design on the coefficient. Tests were also made with nozzles assembled in an automatic injection valve, both with a plain and with a helically grooved stem. It was found that a smooth passage before the orifice is requisite for high flow efficiency. A beveled leading edge before the orifice gave a higher coefficient of discharge than a rounded edge. The results with the nozzles assembled in an automatic injection valve having a plain stem duplicated those with the nozzles assembled at the end of a straight tube of constant diameter. Lower coefficients were obtained with the nozzles assembled in an injection valve having a helically grooved stem. When the coefficients of nozzles of any one geometrical shape were plotted against values of corresponding Reynold's numbers for the orifice diameters and rates of flow tested, it was found that experimental points were distributed along a single curve.

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

    NASA Technical Reports Server (NTRS)

    Williams, Richard; Tulpule, Sharayu; Hawman, Michael

    1990-01-01

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

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

    NASA Technical Reports Server (NTRS)

    1973-01-01

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

  19. Additive Manufacturing a Liquid Hydrogen Rocket Engine

    NASA Technical Reports Server (NTRS)

    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.

  20. Linear quadratic servo control of a reusable rocket engine

    NASA Technical Reports Server (NTRS)

    Musgrave, Jeffrey L.

    1991-01-01

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

  1. Ground test facility for SEI nuclear rocket engines

    SciTech Connect

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

    1992-08-01

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

  2. Reusable rocket engine turbopump condition monitoring

    NASA Technical Reports Server (NTRS)

    Hampson, M. E.

    1984-01-01

    Significant improvements in engine readiness with reductions in maintenance costs and turn-around times can be achieved with an engine condition monitoring systems (CMS). The CMS provides health status of critical engine components, without disassembly, through monitoring with advanced sensors. Engine failure reports over 35 years were categorized into 20 different modes of failure. Rotor bearings and turbine blades were determined to be the most critical in limiting turbopump life. Measurement technologies were matched to each of the failure modes identified. Three were selected to monitor the rotor bearings and turbine blades: the isotope wear detector and fiberoptic deflectometer (bearings), and the fiberoptic pyrometer (blades). Signal processing algorithms were evaluated for their ability to provide useful health data to maintenance personnel. Design modifications to the Space Shuttle Main Engine (SSME) high pressure turbopumps were developed to incorporate the sensors. Laboratory test fixtures have been designed for monitoring the rotor bearings and turbine blades in simulated turbopump operating conditions.

  3. Heated-Pressure-Ball Monopropellant Rocket Engine

    NASA Technical Reports Server (NTRS)

    Greene, William D.

    2005-01-01

    A recent technology disclosure presents a concept for a monopropellant thermal spacecraft thruster that would feature both the simplicity of a typical prior pressure-fed propellant supply system and the smaller mass and relative compactness of a typical prior pump-fed system. The source of heat for this thruster would likely be a nuclear- fission reactor. The propellant would be a cryogenic fluid (a liquefied low-molecular-weight gas) stored in a tank at a low pressure. The propellant would flow from the tank, through a feedline, into three thick-walled spherical tanks, denoted pressure balls, that would be thermally connected to the reactor. Valves upstream and downstream of the pressure balls would be operated in a three-phase cycle in which propellant would flow into one pressure ball while the fluid underwent pressurization through heating in another ball and pressurized propellant was discharged from the remaining ball into the reactor. After flowing through the reactor, wherein it would be further heated, the propellant would be discharged through an exhaust nozzle to generate thrust. A fraction of the pressurized gas from the pressure balls would be diverted to maintain the desired pressure in the tank.

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

    NASA Technical Reports Server (NTRS)

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

    1977-01-01

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

  5. Rocket engine thrust chamber heat transfer calculation and analysis

    NASA Technical Reports Server (NTRS)

    Saha, H.

    1974-01-01

    A parametric study of the heat transfer rate along the wall of a rocket nozzle is presented. The influences of different parameters; laminar and turbulent Lewis number, mixture ratio, initial wall temperature distribution, and eddy viscosity, were considered. The numerical evaluation of these influences on heat transfer rate was done by using three different compressible, reacting laminar and turbulent boundary layer computer programs; MABL (Mass Addition Boundary Layer Program), MABL-KE (MABL program is modified to include turbulent kinetic energy equation), and BLIMP (Boundary Layer Integral Matrix Procedure). This study also provided an excellent opportunity to evaluate the efficiencies of these three computer programs and to suggest one of them for future computational purposes.

  6. Nozzle geometry and forward velocity effects on noise for CTOL engine-over-the-wing concept

    NASA Technical Reports Server (NTRS)

    Vonglahn, U. H.; Goodykoontz, J. H.; Wagner, J. M.

    1973-01-01

    Acoustic shielding benefits for jet noise of engine-over-the-wing for conventional aircraft (CTOL) application were studied with and without forward velocity for various small-scale nozzles. These latter included convergent, bypass and mixer, with and without forward ejector, nozzles. A 13-inch free jet was used to provide forward velocity. Farfield noise data were obtained for subsonic jet velocities from 650 to 980 ft/sec and forward velocities from zero to 360 ft/sec. The studies showed that although shielding benefits were obtained with all nozzles, the greatest benefits were obtained with mixer nozzles. The absolute magnitude of the jet noise shielding benefits with forward velocity was similar to the variation in nozzle-only noise with forward velocity.

  7. NASA Tests Rocket Engine for Commercial Vehicle

    NASA Image and Video Library

    NASA's John C. Stennis Space Center in Mississippi conducted a successful test firing Wednesday of the liquid-fuel AJ26 engine that will power the first stage of Orbital Sciences Corp.'s Taurus II ...

  8. Reusable rocket engine turbopump condition monitoring

    NASA Technical Reports Server (NTRS)

    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.

  9. Composite material application for liquid rocket engines

    NASA Technical Reports Server (NTRS)

    Heubner, S. W.

    1982-01-01

    With increasing emphasis on improving engine thrust-to-weight ratios to provide improved payload capabilities, weight reductions achievable by the use of composites have become attractive. Of primary significance is the weight reduction offered by composites, although high temperature properties and cost reduction were also considered. The potential for application of composites to components of Earth-to-orbit hydrocarbon engines and orbit-to-orbit LOX/H2 engines was assessed. The components most likely to benefit from the application of composites were identified, as were the critical technology areas where developed would be required. Recommendations were made and a program outlined for the design, fabrication, and demonstration of specific engine components.

  10. Liquid rocket engine combustion stabilization devices

    NASA Technical Reports Server (NTRS)

    1974-01-01

    Combustion instability, which results from a coupling of the combustion process and the fluid dynamics of the engine system, was investigated. The design of devices which reduce coupling (combustion chamber baffles) and devices which increase damping (acoustic absorbers) are described. Included in the discussion are design criteria and recommended practices, structural and mechanical design, thermal control, baffle geometry, baffle/engine interactions, acoustic damping analysis, and absorber configurations.

  11. Developing Avionics Hardware and Software for Rocket Engine Testing

    NASA Technical Reports Server (NTRS)

    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.

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

    NASA Technical Reports Server (NTRS)

    Binder, Michael

    1993-01-01

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

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

    NASA Astrophysics Data System (ADS)

    Binder, Michael

    1993-07-01

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

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

    NASA Astrophysics Data System (ADS)

    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.

  15. Analysis of nozzle effect on pulsed detonation engine performance based on laser absorption spectroscopy with Doppler frequency shift

    NASA Astrophysics Data System (ADS)

    Huang, Xiao-long; Li, Ning; Weng, Chun-sheng; Lv, Xiao-jing

    2016-10-01

    An optical experiment system of tunable diode laser absorption spectroscopy is designed for valveless gas-liquid PDE to reveal the mechanism of nozzle improved the thrust performance. The velocity of detonation exhaust with non-nozzle, convergent nozzle, divergent nozzle and convergent-divergent nozzle is tested by laser Doppler velocimetry. The results indicate that laser Doppler method can accurately infer the instantaneous flow velocity, especially the velocity platform where contributes more to the engine impulse. The maximum value is 1222.66 m/s, 1128.52 m/s, 1338.64 m/s and 1296.93 m/s, the time of duration which the velocity is greater than 400m/s is 8.51ms, 7.58ms, 5.83ms and 17.62ms of the velocity under the condition of non-nozzle, convergent nozzle, divergent nozzle and convergent-divergent nozzle respectively.

  16. Transient Simulation of the Integrated Powerhead Demonstrator (IPD) Rocket Engine

    NASA Technical Reports Server (NTRS)

    Guidos, Mike; Seymour, Dave

    2003-01-01

    The IPD rocket engine is the product of a joint Air Force/NASA program to demonstrate the concept of a full-flow staged combustion power cycle. In this type of rocket engine cycle, both the entire fuel flow and the entire oxidizer flow are combusted in one of two pre-bumers and used to drive two turbopumps, thus utilizing the flow work of the total propellant flow. The basis of the IPD engine program relies on integrating newly developed and pre-existing hardware to demonstrate the component and material technologies to make this concept feasible, while simultaneously saving development time and costs. To provide insight to the project team and contractors during engine development and test phases, the engine system was modeled at MSFC using the Rocket Engine Transient Simulation (ROCETS) software to analyze system performance and determine component integration issues. The ROCETS software is used extensively a MSFC to perform steady-state power-balances and transient simulations of thermodynamic power and general fluid systems. The software is favored for its capability to solve large systems of non-linear equations, its librarie of fluid properties and flow devices, its flexibility to modify existing code to improve the physics-derived approximations of real fluid thermodynamic behavior, and the ability to add unique system constraints. The purpose of this paper is to present the methodology used to model the IPD engine system, detail the pitfalls encountered with the software, and explain the approximations made to more accurately represent engine component and fluid behavior. Engine system performance output from the model will be presented and explained in comparison with real fluid and component behavior.

  17. Reusable Rocket Engine Operability Modeling and Analysis

    NASA Technical Reports Server (NTRS)

    Christenson, R. L.; Komar, D. R.

    1998-01-01

    This paper describes the methodology, model, input data, and analysis results of a reusable launch vehicle engine operability study conducted with the goal of supporting design from an operations perspective. Paralleling performance analyses in schedule and method, this requires the use of metrics in a validated operations model useful for design, sensitivity, and trade studies. Operations analysis in this view is one of several design functions. An operations concept was developed given an engine concept and the predicted operations and maintenance processes incorporated into simulation models. Historical operations data at a level of detail suitable to model objectives were collected, analyzed, and formatted for use with the models, the simulations were run, and results collected and presented. The input data used included scheduled and unscheduled timeline and resource information collected into a Space Transportation System (STS) Space Shuttle Main Engine (SSME) historical launch operations database. Results reflect upon the importance not only of reliable hardware but upon operations and corrective maintenance process improvements.

  18. Turnaround Time Modeling for Conceptual Rocket Engines

    NASA Technical Reports Server (NTRS)

    Nix, Michael; Staton, Eric J.

    2004-01-01

    Recent years have brought about a paradigm shift within NASA and the Space Launch Community regarding the performance of conceptual design. Reliability, maintainability, supportability, and operability are no longer effects of design; they have moved to the forefront and are affecting design. A primary focus of this shift has been a planned decrease in vehicle turnaround time. Potentials for instituting this decrease include attacking the issues of removing, refurbishing, and replacing the engines after each flight. less, it is important to understand the operational affects of an engine on turnaround time, ground support personnel and equipment. One tool for visualizing this relationship involves the creation of a Discrete Event Simulation (DES). A DES model can be used to run a series of trade studies to determine if the engine is meeting its requirements, and, if not, what can be altered to bring it into compliance. Using DES, it is possible to look at the ways in which labor requirements, parallel maintenance versus serial maintenance, and maintenance scheduling affect the overall turnaround time. A detailed DES model of the Space Shuttle Main Engines (SSME) has been developed. Trades may be performed using the SSME Processing Model to see where maintenance bottlenecks occur, what the benefits (if any) are of increasing the numbers of personnel, or the number and location of facilities, in addition to trades previously mentioned, all with the goal of optimizing the operational turnaround time and minimizing operational cost. The SSME Processing Model was developed in such a way that it can easily be used as a foundation for developing DES models of other operational or developmental reusable engines. Performing a DES on a developmental engine during the conceptual phase makes it easier to affect the design and make changes to bring about a decrease in turnaround time and costs.

  19. FDNS code to predict wall heat fluxes or wall temperatures in rocket nozzles

    NASA Technical Reports Server (NTRS)

    Karr, Gerald R.

    1993-01-01

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

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

    NASA Technical Reports Server (NTRS)

    Morris, C. I.

    2003-01-01

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

  1. Computational analysis of liquid hypergolic propellant rocket engines

    NASA Technical Reports Server (NTRS)

    Krishnan, A.; Przekwas, A. J.; Gross, K. W.

    1992-01-01

    The combustion process in liquid rocket engines depends on a number of complex phenomena such as atomization, vaporization, spray dynamics, mixing, and reaction mechanisms. A computational tool to study their mutual interactions is developed to help analyze these processes with a view of improving existing designs and optimizing future designs of the thrust chamber. The focus of the article is on the analysis of the Variable Thrust Engine for the Orbit Maneuvering Vehicle. This engine uses a hypergolic liquid bipropellant combination of monomethyl hydrazine as fuel and nitrogen tetroxide as oxidizer.

  2. Exhaust nozzle control and core engine fuel control for turbofan engine

    SciTech Connect

    Camp, V.T.

    1981-10-13

    This control serves to optimize thrust during steady state and transient operation modes of a turbofan engine of the mixed flow type by adjusting or trimming the exhaust nozzle area as a function of fan pressure ratio and fan rotor speed and by adjusting or trimming the core engine fuel flow as a function of fan rotor speed and/or turbine inlet temperature. The control serves to enhance stability by assuring airflow in the engine and its inlet is within a given value avoiding inlet buzz and high distortion to the engine and avoiding even transient operation in conditions that might cause compressor flow instability or stall. Fuel flow is adjusted or trimmed as a function of fan rotor speed or turbine inlet temperature limits depending on which is calling for the least amount of fuel.

  3. Advanced nuclear rocket engine mission analysis

    SciTech Connect

    Ramsthaler, J.; Farbman, G.; Sulmeisters, T.; Buden, D.; Harris, P.

    1987-12-01

    The use of a derivative of the NERVA engine developed from 1955 to 1973 was evluated for potential application to Air Force orbital transfer and maneuvering missions in the time period 1995 to 2020. The NERVA stge was found to have lower life cycle costs (LCC) than an advanced chemical stage for performing low earth orbit (LEO) to geosynchronous orbit (GEO0 missions at any level of activity greater than three missions per year. It had lower life cycle costs than a high performance nuclear electric engine at any level of LEO to GEO mission activity. An examination of all unmanned orbital transfer and maneuvering missions from the Space Transportation Architecture study (STAS 111-3) indicated a LCC advantage for the NERVA stage over the advanced chemical stage of fifteen million dollars. The cost advanced accured from both the orbital transfer and maneuvering missions. Parametric analyses showed that the specific impulse of the NERVA stage and the cost of delivering material to low earth orbit were the most significant factors in the LCC advantage over the chemical stage. Lower development costs and a higher thrust gave the NERVA engine an LCC advantage over the nuclear electric stage. An examination of technical data from the Rover/NERVA program indicated that development of the NERVA stage has a low technical risk, and the potential for high reliability and safe operation. The data indicated the NERVA engine had a great flexibility which would permit a single stage to perform all Air Force missions.

  4. Investigation of installation effects on twin-engine convergent-divergent nozzles

    NASA Technical Reports Server (NTRS)

    Bare, E. A.; Berrier, B. L.

    1983-01-01

    An investigation was conducted in the Langley 16-Foot Transonic Tunnel to determine installation effects on convergent-divergent nozzles applicable to twin-engine reduced-power supersonic cruise aircraft. Tests were conducted at Mach numbers from 0.50 to 1.20, angles of attack from -5 deg to 9 deg, and at nozzle pressure ratios from jet off (1.0) to 8.0. The effects of empennage arrangement, nozzle length, and afterbody closure on total and component drag coefficients were investigated.

  5. Effect of empennage arrangement on single-engine nozzle/afterbody static pressures at transonic speeds

    NASA Technical Reports Server (NTRS)

    Henderson, William P.; Burley, James R., II

    1987-01-01

    An investigation has been conducted in the Langley 16-Foot Transonic Tunnel to determine the effects on empennage arrangement on single-engine nozzle/afterbody static pressures. Tests were done at Mach numbers from 0.60 to 1.20, nozzle pressure ratios from 1.0 (jet off) to 8.0. and angles of attack from -3 to 9 deg (at jet off conditions), depending on Mach number. Three empennage arrangements (aft, staggered, and forward) were investigated. Extensive measurements were made of static pressure on the nozzle/afterbody in the vicinity of the tail surfaces.

  6. Variable area nozzle for gas turbine engines driven by shape memory alloy actuators

    NASA Technical Reports Server (NTRS)

    Rey, Nancy M. (Inventor); Miller, Robin M. (Inventor); Tillman, Thomas G. (Inventor); Rukus, Robert M. (Inventor); Kettle, John L. (Inventor); Dunphy, James R. (Inventor); Chaudhry, Zaffir A. (Inventor); Pearson, David D. (Inventor); Dreitlein, Kenneth C. (Inventor); Loffredo, Constantino V. (Inventor)

    2001-01-01

    A gas turbine engine includes a variable area nozzle having a plurality of flaps. The flaps are actuated by a plurality of actuating mechanisms driven by shape memory alloy (SMA) actuators to vary fan exist nozzle area. The SMA actuator has a deformed shape in its martensitic state and a parent shape in its austenitic state. The SMA actuator is heated to transform from martensitic state to austenitic state generating a force output to actuate the flaps. The variable area nozzle also includes a plurality of return mechanisms deforming the SMA actuator when the SMA actuator is in its martensitic state.

  7. Expert System Architecture for Rocket Engine Numerical Simulators: A Vision

    NASA Technical Reports Server (NTRS)

    Mitra, D.; Babu, U.; Earla, A. K.; Hemminger, Joseph A.

    1998-01-01

    Simulation of any complex physical system like rocket engines involves modeling the behavior of their different components using mostly numerical equations. Typically a simulation package would contain a set of subroutines for these modeling purposes and some other ones for supporting jobs. A user would create an input file configuring a system (part or whole of a rocket engine to be simulated) in appropriate format understandable by the package and run it to create an executable module corresponding to the simulated system. This module would then be run on a given set of input parameters in another file. Simulation jobs are mostly done for performance measurements of a designed system, but could be utilized for failure analysis or a design job such as inverse problems. In order to use any such package the user needs to understand and learn a lot about the software architecture of the package, apart from being knowledgeable in the target domain. We are currently involved in a project in designing an intelligent executive module for the rocket engine simulation packages, which would free any user from this burden of acquiring knowledge on a particular software system. The extended abstract presented here will describe the vision, methodology and the problems encountered in the project. We are employing object-oriented technology in designing the executive module. The problem is connected to the areas like the reverse engineering of any simulation software, and the intelligent systems for simulation.

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

    NASA Technical Reports Server (NTRS)

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

    1998-01-01

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

  9. Advanced Methods for Aircraft Engine Thrust and Noise Benefits: Nozzle-Inlet Flow Analysis

    NASA Technical Reports Server (NTRS)

    Gilinsky, Mikhail; Morgan, Morris H.; Povitsky, Alex; Schkolnikov, Natalia; Njoroge, Norman; Coston, Calvin; Blankson, Isaiah M.

    2001-01-01

    The Fluid Mechanics and Acoustics Laboratory at Hampton University (HU/FM&AL) jointly with the NASA Glenn Research Center has conducted four connected subprojects under the reporting project. Basically, the HU/FM&AL Team has been involved in joint research with the purpose of theoretical explanation of experimental facts and creation of accurate numerical simulation techniques and prediction theory for solution of current problems in propulsion systems of interest to the NAVY and NASA agencies. This work is also supported by joint research between the NASA GRC and the Institute of Mechanics at Moscow State University (IM/MSU) in Russia under a CRDF grant. The research is focused on a wide regime of problems in the propulsion field as well as in experimental testing and theoretical and numerical simulation analyses for advanced aircraft and rocket engines. The FM&AL Team uses analytical methods, numerical simulations and possible experimental tests at the Hampton University campus. The fundamental idea uniting these subprojects is to use nontraditional 3D corrugated and composite nozzle and inlet designs and additional methods for exhaust jet noise reduction without essential thrust loss and even with thrust augmentation. These subprojects are: (1) Aeroperformance and acoustics of Bluebell-shaped and Telescope-shaped designs; (2) An analysis of sharp-edged nozzle exit designs for effective fuel injection into the flow stream in air-breathing engines: triangular-round, diamond-round and other nozzles; (3) Measurement technique improvement for the HU Low Speed Wind Tunnel; a new course in the field of aerodynamics, teaching and training of HU students; experimental tests of Mobius-shaped screws: research and training; (4) Supersonic inlet shape optimization. The main outcomes during this reporting period are: (l) Publications: The AIAA Paper #00-3170 was presented at the 36th AIAA/ASME/SAE/ASEE Joint Propulsion Conference, 17-19 June, 2000, Huntsville, AL. The AIAA

  10. Advanced Methods for Aircraft Engine Thrust and Noise Benefits: Nozzle-Inlet Flow Analysis

    NASA Technical Reports Server (NTRS)

    Gilinsky, Mikhail; Morgan, Morris H.; Povitsky, Alex; Schkolnikov, Natalia; Njoroge, Norman; Coston, Calvin; Blankson, Isaiah M.

    2001-01-01

    The Fluid Mechanics and Acoustics Laboratory at Hampton University (HU/FM&AL) jointly with the NASA Glenn Research Center has conducted four connected subprojects under the reporting project. Basically, the HU/FM&AL Team has been involved in joint research with the purpose of theoretical explanation of experimental facts and creation of accurate numerical simulation techniques and prediction theory for solution of current problems in propulsion systems of interest to the NAVY and NASA agencies. This work is also supported by joint research between the NASA GRC and the Institute of Mechanics at Moscow State University (IM/MSU) in Russia under a CRDF grant. The research is focused on a wide regime of problems in the propulsion field as well as in experimental testing and theoretical and numerical simulation analyses for advanced aircraft and rocket engines. The FM&AL Team uses analytical methods, numerical simulations and possible experimental tests at the Hampton University campus. The fundamental idea uniting these subprojects is to use nontraditional 3D corrugated and composite nozzle and inlet designs and additional methods for exhaust jet noise reduction without essential thrust loss and even with thrust augmentation. These subprojects are: (1) Aeroperformance and acoustics of Bluebell-shaped and Telescope-shaped designs; (2) An analysis of sharp-edged nozzle exit designs for effective fuel injection into the flow stream in air-breathing engines: triangular-round, diamond-round and other nozzles; (3) Measurement technique improvement for the HU Low Speed Wind Tunnel; a new course in the field of aerodynamics, teaching and training of HU students; experimental tests of Mobius-shaped screws: research and training; (4) Supersonic inlet shape optimization. The main outcomes during this reporting period are: (l) Publications: The AIAA Paper #00-3170 was presented at the 36th AIAA/ASME/SAE/ASEE Joint Propulsion Conference, 17-19 June, 2000, Huntsville, AL. The AIAA

  11. Deimos Methane-Oxygen Rocket Engine Test Results

    NASA Astrophysics Data System (ADS)

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

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

  12. Rocket Engine Plume Diagnostics at Stennis Space Center

    NASA Technical Reports Server (NTRS)

    Tejwani, Gopal D.; Langford, Lester A.; VanDyke, David B.; McVay, Gregory P.; Thurman, Charles C.

    2003-01-01

    The Stennis Space Center has been at the forefront of development and application of exhaust plume spectroscopy to rocket engine health monitoring since 1989. Various spectroscopic techniques, such as emission, absorption, FTIR, LIF, and CARS, have been considered for application at the engine test stands. By far the most successful technology h a been exhaust plume emission spectroscopy. In particular, its application to the Space Shuttle Main Engine (SSME) ground test health monitoring has been invaluable in various engine testing and development activities at SSC since 1989. On several occasions, plume diagnostic methods have successfully detected a problem with one or more components of an engine long before any other sensor indicated a problem. More often, they provide corroboration for a failure mode, if any occurred during an engine test. This paper gives a brief overview of our instrumentation and computational systems for rocket engine plume diagnostics at SSC. Some examples of successful application of exhaust plume spectroscopy (emission as well as absorption) to the SSME testing are presented. Our on-going plume diagnostics technology development projects and future requirements are discussed.

  13. Conjugate (solid/fluid) computational fluid dynamics analysis of the space shuttle solid rocket motor nozzle/case and case field joints

    NASA Technical Reports Server (NTRS)

    Doran, D.; Keeton, L. W.; Dionne, P. J.; Singhal, A. K.

    1989-01-01

    Three-dimensional, conjugate (solid/fluid) heat transfer analyses of new designs of the Solid Rocket Motor (SRM) nozzle/case and case field joints are described. The main focus was to predict the consequences of multiple rips (or debonds) in the ambient cure adhesive packed between the nozzle/case joint surfaces and the bond line between the mating field joint surfaces. The models calculate the transient temperature responses of the various materials neighboring postulated flow/leakpaths into, past, and out from the nozzle/case primary O-ring cavity and case field capture O-ring cavity. These results were used to assess if the design was failsafe (i.e., no potential O-ring erosion) and reusable (i.e., no excessive steel temperatures). The models are adaptions and extensions of the general purpose PHOENICS fluid dynamics code. A non-orthogonal coordinate system was employed and 11,592 control cells for the nozzle/case and 20,088 for the case field joints are used with non-uniform distribution. Physical properties of both fluid and solids are temperature dependent. A number of parametric studies were run for both joints with results showing temperature limits for reuse for the steel case on the nozzle joint being exceeded while the steel case temperatures for the field joint were not. O-ring temperatures for the nozzle joint predicted erosion while for the field joint they did not.

  14. Nozzle Side Load Testing and Analysis at Marshall Space Flight Center

    NASA Technical Reports Server (NTRS)

    Ruf, Joseph H.; McDaniels, David M.; Brown, Andrew M.

    2009-01-01

    Realistic estimates of nozzle side loads, the off-axis forces that develop during engine start and shutdown, are important in the design cycle of a rocket engine. The estimated magnitude of the nozzle side loads has a large impact on the design of the nozzle shell and the engine s thrust vector control system. In 2004 Marshall Space Flight Center (MSFC) began developing a capability to quantify the relative magnitude of side loads caused by different types of nozzle contours. The MSFC Nozzle Test Facility was modified to measure nozzle side loads during simulated nozzle start. Side load results from cold flow tests on two nozzle test articles, one with a truncated ideal contour and one with a parabolic contour are provided. The experimental approach, nozzle contour designs and wall static pressures are also discussed

  15. Multiobjective Optimization of Rocket Engine Pumps Using Evolutionary Algorithm

    NASA Technical Reports Server (NTRS)

    Oyama, Akira; Liou, Meng-Sing

    2001-01-01

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

  16. Method of fabricating a rocket engine combustion chamber

    NASA Technical Reports Server (NTRS)

    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.

  17. Investigation of a nozzle instability on an F100 engine equipped with a digital electronic engine control

    NASA Technical Reports Server (NTRS)

    Burcham, F. W., Jr.; Zeller, J. R.

    1984-01-01

    An instability in the nozzle of the F100 engine, equipped with a digital electronic engine control (DEEC), was observed during a flight evaluation on an F-15 aircraft. The instability occurred in the upper left hand corner (ULMC) of the flight envelope during augmentation. The instability was not predicted by stability analysis, closed-loop simulations of the the engine, or altitude testing of the engine. The instability caused stalls and augmentor blowouts. The nozzle instability and the altitude testing are described. Linear analysis and nonlinear digital simulation test results are presented. Software modifications on further flight test are discussed.

  18. Rocket engine coaxial injector liquid/gas interface flow phenomena

    NASA Astrophysics Data System (ADS)

    Mayer, Wolfgang; Kruelle, Gerd

    1992-07-01

    Coaxial injectors are used for injecting and mixing propellants in cryogenic rocket engines. Theoretical and experimental studies are reported which show the significance for atomization and mixing of the physical processes involved in the coaxial injector flow. The impact of internal fluid jet motions on surface irritation is demonstrated. A model is presented which calculates droplet atomization quantities such as frequency, droplet diameter, and liquid core shape.

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

    NASA Technical Reports Server (NTRS)

    1979-01-01

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

  20. Dynamic response at altitude of a turbojet engine with variable area exhaust nozzle

    NASA Technical Reports Server (NTRS)

    Delio, Gene J; Rosenweig, Solomon

    1952-01-01

    The dynamic characteristics of turbojet engine with variable exhaust nozzle area were investigated over a range of altitudes and flight Mach numbers. These characteristics generalize to standard static level condition. Data resulting from approximate step disturbances in either manipulated variable suggested functional relationships from which functions can be derived. The constants of the transfer functions are listed for five combinations of engine speed and exhaust nozzle area. The minimum data needed to completely define the transfer functions were: experimentally determined dynamic characteristics, such as engine time constant and initial rise ratio, resulting from either manipulated variable; static characteristics determined from steady-state performance curves.