Showing papers on "Liquid-propellant rocket published in 2000"

Rocket and Spacecraft Propulsion: Principles, Practice and New Developments

Abstract: Foreward Authors Preface Acknowledgements Principles of Rocket Propulsion The Thermal Rocket Engine Liquid Propellant Rocket Engines Solid Propellant Rocket Motors Launch Vehicle Dynamics Electric Propulsion Advanced Thermal Rockets Appendix 1: Current Launcher portfolio Appendix 2: Ariane 5 and the Space Shuttle Appendix 3: Orbital motion Bibliography Index.

Third Way of Development of Single-Stage-to-Orbit Propulsion

Abstract: Considering a numberofoptionsfora space-launch vehicle-propulsion system,between advancedrocketmotors and airbreathers,inparticular,thermallyintegrated rocket-basedcombined-cycles,anewcyclethathasbeengiven thenameKLIN (meaning wedgein Russian ), is proposed asthe“ third way.” Itconsists of a combination of a liquid rocket engine and a deep-cooled turbojet (with oxygen addition to atmospheric air ). Currently, the KLIN concept is considered for application with a vertical takeoff horizontal landing vehicle, with the turbojet and the rocket engine optimized individually, with a possibility for incorporating an aerospike-type nozzle. Retaining a rocket trajectoryup toMach 3,theturbojetandrocketengineareassumed tooperatetogetherfromtakeoffwitha gradual reduction in the deep-cooled turbojet output, e nally terminating the turbojet at Mach 6. It can be shown that the KLIN can be manufactured with available or foreseeable technology, and provides a combination of engine weight and specie c impulse that yields twice the payload mass fraction, in addition to other advantages, compared with a rocket-engine-operated vehicle.

A Design Tool for Liquid Rocket Engine Injectors

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

Development Status of the NASA MC-1 (Fastrac) Engine

Abstract: The MC-1 (formerly known as the Fastrac 60K) Engine is being developed for the X-34 technology demonstrator vehicle. It is a pump-fed liquid rocket engine with fixed thrust operating at one rated power level of 60,000 lbf vacuum thrust using a 15:1 area ratio nozzle (slightly higher for the 30:1 flight nozzle). Engine system development testing of the MC-1 has been ongoing since 24 Oct 1998. To date, 48 tests have been conducted on three engines using three separate test stands. This paper will provide some details of the engine, the tests conducted, and the lessons learned to date.

Parallel unsteady turbo-pump simulations for liquid rocket engines

Abstract: This paper reports the progress being made towards complete turbo-pump simulation capability for liquid rocket engines. The Space Shuttle Main Engine (SSME) turbo-pump impeller is used as a test case for the performance evaluation of the MPI, hybrid MPI/Open-MP, and MLP versions of the INS3D code. Then, a computational model of a turbo-pump has been developed for the shuttle upgrade program. Relative motion of the grid system for rotor-stator interaction was obtained by employing overset grid techniques. Unsteady computations for SSME turbo-pump, which contains 101 zones with 31 million grid points, are carried on Origin 2000 systems at NASA Ames Research Center. The approach taken for these simulations, and the performance of the parallel versions of the code are presented.

Rocket with high pressure fueling module

Abstract: A rocket with a high pressure propellant module, comprising a toy bottle rocket that uses gases generated from the reaction between baking soda and vinegar A mechanism controls the release of the gas formed by the chemical reaction and allows the gas to reach a high pressure inside the rocket prior to release

Hybrid rocket propulsion system including array of hybrid or fluid attitude-control rocket engines

Abstract: This propulsion system of a rocket motor assembly includes an array of attitude-control rocket engines, one or more oxidizer-fluid sources, one or more ignition-fluid sources, and, optionally, one or more primary rocket engines. Each of the attitude-control rocket engines has a respective combustion chamber and is offset from the longitudinal axis of the rocket motor assembly so that when a selected one or group of the attitude-control rocket engines is fired, the flight path of the assembly is diverted and/or the rocket assembly spins. The oxidizer-fluid and ignition-fluid sources are in operative communication with the attitude-control rocket engines to respectively permit oxidizer fluid and ignition fluid to be supplied to selected ones or groups of the attitude-control rocket engines. Optionally, a portion of the ignition fluid from the ignition-fluid source can be cooled and used to pressurize the oxidizer-fluid source.

Modeling and Robust Analysis of a Liquid Rocket Engine

Abstract: Stability and dynamic performance of liquid-propellant rocket engines (LPRE) are two of the fundamental issues in the enginevehicle integration process. This analysis requires the construction of a detailed model, trying to capture the most realistic phenomena involved, which generally include several sources of uncertainties. In this paper, a methodology for robust modeling and stability analysis is presented. Firstly, the linear models of the LPRE components are obtained by modeling the various physical processes, at a nominal regime of operation. Afterwards, the Laplace transform is applied to derive a block diagram representation of the linear LPRE. The stability study and dynamic analysis are carried out taking in account the uncertainties in parameters of the plant. The robust stability is assured via the Generalized Kharitonov's Theorem; and the robust frequency and step responses are obtained with the use of specialized MATLAB toolboxes. The robust performance of the system in the time domain is obtained in terms of the response to step function input, while taking into account the plant uncertainties, also known as robust step response. A practical application is illustrated by analyzing a simple pressurefed LPRE system. Copyright © 2000 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved. INTRODUCTION During the engine-vehicle integration phase, the knowledge of the engine dynamics can avoid serious instability problems, caused by the dynamic coupling of the propellant feed system, the rocket engine, and the vehicle structure. The dynamic analysis helps to characterize many aspects of the engine system operation. It requires the construction of a mathematical model to describe approximately the most important phenomena of the actual system. The engine itself contains several sources of intense pressure fluctuations due to turbulent flow in feed lines, fluttering of pump wheel blades, vibrations of control valves, and unsteady motion in the combustion chamber and gas generator. The coupling of these oscillations with the natural frequencies of the system structure represents often a source of instability, sometimes observed with catastrophic consequences. In general, solving the dynamic equations is not a simple task, since the more realistic models are time-variant, non-linear and of high order, and include uncertainties used to represent imprecise physical parameters or unknown operating conditions. Even for linear time-invariant models, this problem can become quite difficult to solve when the parameters deviate from the nominal values in a known rated regime. Traditional numerical methods of analysis are not suited for these uncertain models since they require a great amount of computational effort. Besides, the direct approach of solving the coupled differential

Catalyst Bed Instability Within the USFE H2O2/JP-8 Rocket Engine

Abstract: Orbital Sciences Corporation has been awarded a contract by NASA's Marshall Space Flight Center, in cooperation with the U.S. Air Force Research Laboratory's Military Space Plane Technology Program Office, for the Upper Stage Flight Experiment (USFE) program. Orbital is designing, developing, and will flight test a new low-cost, 10,000 lbf hydrogen peroxide/ JP-8 pressure fed liquid rocket. During combustion chamber tests at NASA Stennis Space Center (SSC) of the USFE engine, the catalyst bed showed a low frequency instability occurring as the H202 flow reached about 1/3 its design rate. This paper reviews the USFE catalyst bed and combustion chamber and its operation, then discusses the dynamics of the instability. Next the paper describes the dynamic computer model used to recreate the instability. The model was correlated to the SSC test data, and used to investigate possible solutions to the problem. The combustion chamber configuration which solved the instability is shown, and the subsequent stable operation presented.

Development of Metal Matrix Composites for NASA'S Advanced Propulsion Systems

Abstract: The state-of-the-art development of several aluminum and copper based Metal Matrix Composites (MMC) for NASA's advanced propulsion systems will be presented. The presentation's goal is to provide an overview of NASA-Marshall Space Flight Center's planned and on-going activities in MMC for advanced liquid rocket engines such as the X-33 vehicle's Aerospike and X-34 Fastrac engine. The focus will be on lightweight and environmental compatibility with oxygen and hydrogen of key MMC materials, within each NASA's new propulsion application, that will provide a high payoff for NASA's reusable launch vehicle systems and space access vehicles. Advanced MMC processing techniques such as plasma spray, centrifugal casting, pressure infiltration casting will be discussed. Development of a novel 3D printing method for low cost production of composite preform, and functional gradient MMC to enhanced rocket engine's dimensional stability will be presented.

Liquid-propellant rocket reheat engine

Abstract: rocketry. SUBSTANCE: engine has chamber, fuel and oxidizer booster pumps, turbo-pump unit, gas generator, thrust regulator with programmed reversal actuator and propellant component ratio throttle valve. Provision is made for programmed starting and ignition system which includes fuel reservoir connected to low-flow loop of thrust regulator and fuel ampoule whose outlet is connected with fuel injectors of gas generator through start-and-shut-off valve. This reservoir is also connected to self- contained ignition injectors located in combustion chamber of engine through starting fuel ampoule. Used as oxidizer booster-pump is screw pump driven by gas turbine; its working medium is oxidizing gas which is exhausted to outlet manifold of booster pump. Feed pipe line of gas turbine of prepump is provided with heat exchanger where has intended for pressurization of rocket tanks is preheated. EFFECT: possibility of working of engine on low-toxic non-self-ignition components of propellant; reduced dynamic action of engine on rocket in the course of start; enhanced efficiency of pressurization system. 8 cl, 1 dwg

The 260: The Largest Solid Rocket Motor Ever Tested

Abstract: Aerojet in the mid 1960s, under contract to NASA, built and static hot fire tested the largest solid rocket motor (SRM) in history for the purpose of demonstrating the feasibility of utilizing large SRMs for space exploration. This program successfully fabricated two high strength steel chambers, loaded each with approximately 1,68 million pounds of propellant, and static test fired these giants with their nozzles up from an underground silo located adjacent to the Florida everglades. Maximum thrust and total impulse in excess of 5,000,000 lbf and 3,470,000,000 lbf-sec were achieved. Flames from the second firing, conducted at night, were seen over eighty miles away. For comparative purposes: the thrust developed was nearly 100 times that of a Minuteman III second stage and the 260 in.-dia cross-section was over 3 times that of the Space Shuttle SRM.

Ring chamber of liquid-propellant rocket engine

Abstract: mechanical engineering; engines using for their operation cryogenic components, such as oxygen and hydrogen. SUBSTANCE: ring chamber of rocket engine has nozzle, combustion chamber with mixing head arranged inside nozzle and fuel and oxidizer supply manifold. Combustion chamber has profiled inner wall placed along longitudinal axis of chamber. Cooled cylinder is installed inside combustion chamber. One end of cylinder is connected with mixing head, and the other, with central part of nozzle, forming together with profiled inner wall of combustion chamber a ring critical section. Space of fuel supply manifold is divided into three parts isolated from each other. One part of space is connected with fuel supply space of mixing head, and two other parts, with spaces of cooling duct of ring combustion chamber and cylinder which are connected to each other. EFFECT: provision of high specific thrust pulse at minimum overall dimensions of chamber. 2 cl, 4 dwg

Liquid cryogenic propellant rocket engine with closed loop of turbine drive of turbopump unit (versions)

Abstract: rocketry; cryogenic liquid propellant rocket engines. SUBSTANCE: liquid-propellant engine has chamber with injector assembly and regenerative cooling line, turbopump unit with oxidizer pump (liquid oxygen) and fuel (hydrocarbon fuel) pump whose outlet mains are connected with injector assembly of chamber and closed loop of turbine drive of turbopump unit which includes circulating pump, regenerative cooling line heat exchanger-heater using gas as heat-transfer agent which are connected in series; gas is obtained in gas generator fed from outlet mains of oxidizer and fuel pumps; heat transfer agent from outlet of above-mentioned heat exchanger-heater is introduced into outlet main of oxidizer pump; turbopump unit includes also turbine, heat exchanger-condenser mounted on outlet main of oxidizer pump; outlet of heat exchanger-condenser is connected with circulating pump inlet through flow regulator. Control of engine is effected by means of component ratio regulator fitted in gas generator fuel supply main. Another version of liquid-propellant engine consists in availability of fuel pump which simultaneously performs function of circulating pump. Besides that, fuel pump outlet is connected with injector assembly of chamber and with chamber regenerative cooling line through controllable flow divider. Systems of these engines include starting systems with starting pump connected in parallel with circulating pump. Pump is connected with autonomous drive through split coupling. After starting the engine, starting pump is switched off. EFFECT: enhanced reliability of engine due to obtaining optimal characteristics in thermal factor and strength characteristics; low cost of engine due to usage of low-cost materials. 14 cl, 2 dwg

Liquid-propellant rocket engine

Abstract: spacecraft; rocket engines. SUBSTANCE: proposed liquid-propellant rocket engine is designed for use in cryogenic stages of launch vehicles and as cruise engines of spacecraft. It includes combustion chamber with regenerative cooling duct, evaporator, fuel and oxidizer pumps, and turbine. Outlet of pump of one of components (fuel or oxidizer) is connected by main line with inlet of coolant line of evaporator, and outlet of evaporator at the same line is connected with inlet of turbine. Outlet of pump of other component communicates with combustion chamber. Engine includes also condenser, intermediate coolant source with controllable valve, intermediate coolant circulating pump with intermediate coolant turbine. Inlet of evaporator at heat carrier line is connected with outlet of intermediate coolant turbine. Outlet of evaporator at heat carrier line is connected with intermediate coolant source by means of controllable valve, and with inlet of intermediate coolant circulating pump. Outlet of intermediate coolant circulating pump is connected with inlet of combustion chamber regenerative cooling duct, and outlet of duct is connected with inlet of intermediate coolant turbine. Inlet of condenser at cooling agent line is connected with outlet of pump of one of components, and outlet of the same line is connected with combustion chamber. Inlet and outlet of condenser at heat carrier line are connected, respectively, with outlet of turbine and inlet of pump of the component. EFFECT: improved efficiency of liquid-propellant rocket engine, enlarged operating capabilities. 1 dwg

Current and Future Rocket Propulsion Testing at NASA Stennis Space Center

Abstract: Year 2000 has been an active one for large-scale propulsion testing at the NASA John C. Stennis Space Center. This paper highlights several of the current-year test programs conducted at the Stennis Space Center (SSC) including the X-33 Aerospike Engine, Ultra Low Cost Engine (ULCE) program, and the Hybrid Sounding Rocket (HYSR) program. Future directions in propulsion test are also introduced including the development of a large-scale Rocket Based Combined Cycle (RBCC) test facility.

Chamber swinging unit of reheat liquid propellant rocket engine

Abstract: rocket manufacturing. SUBSTANCE: proposed unit has bellows with guard rings placed between bellows corrugations. Support rings are hermetically connected with gas duct and combustion chamber. Cardan ring installed outside the bellows is connected with support tings by hinge joints with load bearing brackets. Screen installed inside bellows consists of two telescopic cylindrical shells fitted one into the other with clearance. Cylindrical shells cantilever mounted on support rings form chamber. This chamber is connected with cooling working medium supply main line through service elements made in support rings and is connected with bellows unit space through clearance between shells. Casing is installed outside the guard rings. This casing is made in form of metal cylindrical spiral whose ends are connected with support rings. EFFECT: simplified design and supply of controls, provision of control of rocket thrust vector within wide range of control forces with chamber turned through required angles. 8 cl, 2 dwg

Method for orbit injection of payload by multifunctional launch vehicle of combination arrangement with cruise liquid- propellant rocket engine installations (lrei), multifunctional launch vehicle with cruise lpei and method for its development

Abstract: rocket-space engineering, applicable in development of transportation systems designed for earth-orbit injection of various space vehicles. SUBSTANCE: a combination arrangement of launch vehicle is formed with a lower polypod pack of similar rocket pods. The cruise LPEIs of the central and side rocket pods are started, and the LPEIs operate in accordance with the injection program. The LPEI thrust is reduced for adjustment of inertia and aerodynamic loads on the launch vehicle. The side rocket pods are mounted on the central one symmetrically relative to its longitudinal axis in the sectors formed by the rolling planes of the central pod LPEI. After manufacture of the central rocket pod it is subjected to ground and flight-design tests, as part of the tandem arrangement launch vehicle inclusive. Then, this pod is used for formation of the above lower polypod pack. EFFECT: expanded range of masses of orbit injected payloads. 8 cl, 11 dwg

Liquid-propellant rocket power plant

Abstract: FIELD: space crafts; boost units, launch vehicle stages. SUBSTANCE: plant has tanks of oxidant and fuel with at least one cryogenic component, components feed turbopump sets, combustion with cooling jacket whose inner space is connected to fuel feed turbopump set, and supercharging gas storage cylinder. Moreover, turbocompressor set for circulating gas and cooler are introduced into plant. Supercharging gas storage cylinder is placed in communication with main lines delivering gas into turbopump sets delivering oxidant and fuel and into turbocompressor set providing circulation of gas. Outlet of sets communicate with cooler one outlet of which communicates with combustion chamber cooling jacket through compressor of turbocompressor providing circulation of gas. Outlet of jacket is connected with into main lines delivering gas into turbopump sets providing delivery of oxidant, fuel and turbocompressor set. Other outlet and inlet of cooler communicate with inner space of combustion chamber and with outlet of pump of turbopump set delivering oxidant. EFFECT: enhanced reliability and efficiency of plant, enlarged operating capabilities. 1 dwg

Two-Phase Flow Behavior in a Liquid Propellant Tank.

Abstract: To give appropriate assesment of propellant management systems for space application, the behavior of two-phase flows under low gravity conditions was both experimentally and numerically studied. By use of a numerical analysis method with C-CUP(CIP Combined Unified Procedure)scheme coupled with Level Set method and CSF(continuum Surface Force)model, the flow fields in a liquid rocket propellant tank were simulated. Main concern was placed on the generation of a deep dip of the gas phase. The numerical results were compared with corresponding experimental data obtained in a drop tower, and they showed good agreement. The generation of the dip was clearly revealed, and the phenomenon of gas suction was properly reproduced in the simulation. The effect of gravity force on the dip generation was studied, and it was found that the dip generation was alleviated with increase in the gravity force.

Evaluation of Vortex Chamber Concepts for Liquid Rocket Engine Applications

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

Swinging assembly of liquid-propellant rocket engine

Abstract: liquid-propellant rocket engines. SUBSTANCE: swinging assembly has chamber and pipeline carrying engine products of gas generation and running from turbine to chamber. The latter is mounted relative to pipeline for spatial displacement about point in vicinity of or on chamber axis at preset solid angle. Movable spherical sealing joint incorporating head and shell is provided between chamber and pipeline. Either head or shell has spherical sealing surface and other one has seat accommodating annular sealing member held tight against spherical sealing surface, for example, by means of spring; this member is made of sealing material such as polyfluoroethylene possessing low coefficient of friction on mentioned spherical sealing surface; head and shell are provided with through channels directly communicating on inlet side with gas-generation product delivery pipeline and on outlet side, with engine chamber. EFFECT: facilitated manufacture, reduced cost and size of combustion-chamber swinging assembly. 9 cl, 6 dwg