Showing papers on "Liquid-propellant rocket published in 1998"

Liquid-Propellant Rocket Engine Injector Dynamics

Abstract: The dynamic characteristics of liquid-rocket injectors in the presence of intense combustion-chamber and propellant feed-line oscillations are discussed. Liquid-propellant injectors always function in nonsteady e ow environments and are therefore considered as a dynamic component of an engine. In addition to its main function of injecting propellant and preparing a combustible mixture, an injector simultaneously acts as a sensitive element that may generate and modify e ow oscillations because of its intrinsic unsteadiness and interactions with the combustion-chamber and feed-system dynamics. This paper also addresses nonlinear effects of nonstationary processes occurring in injectors. Various mechanisms for driving self-pulsations in both liquid and gas ‐liquid injectors are summarized systematically. A vital problem of bifurcational unsteadiness of injector operation is considered.

Mixing and Spray Formation in Coaxial Jets

Abstract: The e ow regimes and e ne-scale structure of the mixture resulting from the destabilization of a dense, slow jet by a fast light annular jet are discussed. From the primary shear instability between the streams, it is shown that the momentum ratio M = of the annular to the inner stream is the key parameter 2 2 r u /r u 2 2 1 1 on which the inner potential core length and the condition of a recirculation transition depend. The instability analysis shows that the initial wavelength of the disturbances l is proportional to the vorticity thickness d of the fast stream at the nozzle exit l ; d(r1/r 2) 1/2 ; this further provides, by an original scenario, an estimate of the droplets’ size formed by the capillary instability of the sheet’ s rim formed from the initial interface disturbances. Liquid viscosity is not expected to play any signie cant role in practical conditions in liquid rocket propellant engines. These e ndings are put in relation with a selected review of known or yet unexplained results on the subject. HERE is a frequent need to realize a uniform mixture from two initially segregated streams in many practical instances. A signie cant example is the case of liquid rocket propellant engines because the length of the combustion chamber is limited by the ability of the injecting devices to fragment and mix the reactants down to a sufe cient level of homogeneity for evaporation and combustion to completion at the desired distance from the injector outlet. For technological reasons one of the reactants is usually available in the liquid state and the other in the gas phase; for historical reasons the coaxial geometry is commonly used to merge the two streams. 1 The relative e ow rates of the reactants must be adjusted so that the global stoichiometry is at least respected, or such that the gas phase is in excess for all of the liquid to vaporize and burn, this implies that the gas stream is usually much more rapid than the liquid stream at the injector outlet. This is the situation encountered in H 2/O2 engines, where a slow, dense liquid oxygen (LOX) stream in the central jet of the coaxial injector is surrounded by a fast, light, gaseous hydrogen annular stream. At the root of the interpenetration process between the two phases is a strong shear destabilizing the central liquid jet, which further fragments into a more or less uniform spray. This process, which is contrasted with the case of a simple jet issuing in a quiescent environment, 2 is known as airblast atomization. 3

Initial Growth Rate and Visual Characteristics of a Round Jet into a Sub- to Supercritical Environment of Relevance to Rocket, Gas Turbine, and Diesel Engines

Abstract: : The combustion chamber temperature and pressure in many liquid rocket, gas turbine, and diesel engines are quite high and can reach levels above the critical point for the injected fuels and/or oxidizers. A high pressure chamber is used to investigate and understand the nature of the interaction between the injected fluid and the environment under such conditions. Pure N2 He, and O2 fluids are injected. Several chamber media are selected including, N2, He, and mixtures of CO+N2. The effects of chamber pressure ranging from a subcritical (i.e. relative pressure, P(sub r) = P/P (sub injectant critical 1) value at a supercritical chamber temperature (relative temperature T sub r = T/T sub injectant critical >1) are photographically observed and documented near the injector hole exit region using a CCD camera illuminated by a short-duration back-lit strobe light. At low subcritical chamber pressures, the jets exhibit surface irregularities that amplify downstream, looking intact, shiny, but wavy (sinuous) on the surface that eventually break up into irregularly-shaped small entities.

Cycles for Earth-to-Orbit Propulsion

Abstract: The reduction of Earth-to-orbit launch costs in conjunction with an increase in launcher reliability and operational efe ciency are the key demands on future space transportation systems. Results of various system analyses indicate that these demands can be met with future single-stage-to-orbit (SSTO) vehicles using advanced technologies for both structure and propulsion systems. This paper will provide a classie cation and description of all turbopump feed liquid rocket engine cycles, followed by a combined vehicle/propulsion analysis of the design parameters for the propulsion systems and their associated thermodynamic cycles for the launch of future SSTO vehicles. Existing and projected rocket engine cycles capable of SSTO missions will be presented.

Non-Linear Interactions in Liquid-Propellant Rocket Engine Injectors

Abstract: Experimental results and theoretical explanations of such non-linear phenomena as self-pulsation and non-periodical unsteadiness due to hydraulic instability in liquid propellant flow both in jet and swirl injectors as well as interaction of liquid and gaseous propellant flows in gas-liquid injectors are described and discussed. It is shown that additionally to the abilities of Liquid-Propellant Rocket Engine (LRE) injector to serve as a sensor, amplifier, phase shifter and actuator in LRE as a dynamic system, in some circumstances it can also be a generator of oscillations that can influence both secondary heat release and combustion chamber pressure oscillations, and change stationary characteristics of mixing zone that cause the dynamic characteristics of combustion. Self-pulsation motion was found in liquid injectors: jet injectors with enlargements in their flow path, swirl injectors when using compressible or boiling liquid or having excess radial component of liquid velocity inside the vortex chamber. The latter is typical for jet engine fuel injectors and can lead to cavitation erosion of interior injector passages. Most powerful self-pulsation occurs due to interaction of liquid and gaseous propellant flows in gas-liquid injectors. Such self-pulsation motion was discovered in all types of injectors used in LRE: shear/coaxial, shear/swirl with central LOX and exterior hydrogen stages, and both swirl/swirl coaxial injector with central liquid flow. Dependence of boundaries of self-pulsatio n regime on both operational parameters and principal design features are presented and showed strong influence on real operational conditions. Influence of self-pulsatio n motion on atomization and mixing is presented. Positive and negative consequences of self-pulsation are observed. The phenomena of non-periodica l unsteadiness of injection process in LRE is discovered and studied. It presents temporal changes of flow angle, thickness of spray at the injector exit and cause so called "vibro-activity of combustion", that can cause bifurcations of the operational process in LRE. So, designers of LRE injectors must be informed not only about the linear dynamic characteristics of propellant injectors they use hi LRE but also about the non-linear properties that strongly depends on real parameters of the combustion chamber and hardly can be modelled in ambient conditions of cold flow tests.

Modification and verification testing of a russian nk-33 rocket engine for reusable and restartable applications

Abstract: Flight-qualified liquid propellant rocket engines from the Russian lunar launch program were received at Aerojet, modified to include reusable and restartable features with modern instrumentation and controls, and test fired to verify the modifications. The NK-33 liquid oxygen/kerosene propellant rocket engine was designed and manufactured by Samara State Scientific and Production Enterprise "TRUD" (now known as N.D. Kuznetsov Samara Scientific and Technical Company) of Samara, Russia, for the Soviet N-l launch vehicle. This staged combustion engine delivered high pressure (2109 psia chamber pressure) and high performance (331 seconds vacuum delivered specific impulse) that had never been available in the West for an hydrocarbon engine. Aerojet imported 36 NK-33 engines, along with 9 NK-43 engines, an upper stage version of the same engine, from N.D. Kuznetsov SSTC. The first of the NK33 engines was modified for use on the Kistler K-l launch vehicle. Modifications included replacement of pyrotechnic initiated valves with solenoid actuated valves; replacement of electromechan ical actuators for thrust and mixture ratio control; redesign of purge supply systems; replacement of solid propellant for the turbopump start spinup and the main chamber igniters; redesign and replacement of the thrust frame for addition of a gimbal and thrust vector control mount; addition of valves, pyrotechnics, and plumbing to restart the engine; and replacement of instrumentation and wiring harnesses. This engine was successfully test fired at Aerojet to verify

Liquid tripropellant rocket engine coaxial injector

Abstract: A tripropellant coaxial pintle injector for a rocket engine which enables smooth transitioning between two types of propellants that are adapted to be alternatively mixed with a third type of propellant in a combustion chamber of a rocket engine. The tripropellant coaxial injector includes a pintle for expelling a first propellant into a combustion chamber in a generally radial direction and two or more concentric metering sleeves which define orifices for expelling two or more different propellants into the combustion chamber in an axial direction thus, such a configuration allows for alternative mixing of two propellants with a third propellant in the combustion chamber of a rocket engine.

Stable-combustion oxidizer for hybrid rockets

Abstract: A hybrid heater vaporization system (280, 380) including hybrid heaters (250, 350) can be used to vaporize liquid oxygen as it enters a motor (200, 300) to improve the combustion characteristics of a hybrid rocket (11). The motor (200, 300) preferably includes hybrid fuel both in a substantially cylindrical portion (216, 316) in the substantially cylindrical portion of the motor and a substantially multi-toroidal shaped portion (217, 316) in the forward end of the motor (200, 300). The vaporization system (280, 380) of the present invention finally makes hybrid rockets (11) practical for aerospace applications.

Liquid-Propellant Rocket Engine Health-Monitoring Techniques

Abstract: Health monitoring of liquid-propellant rocket engines (LRE) is one of the key technologies for improving the safety of existing engines and developing reliable next-generation engines. Extensive research has been done on the health monitoring of the Space Shuttle Main Engine and next-generation reusable LRE. A brief overview of these research projects is presented. Research advances on the health monitoring of the Long March Main Engine YF-20B are described in detail. The failure mode simulation and analysis of the YF-20B engine are introduced. A component module-based diagnosis method is developed, and a fuzzy hypersphere neural network is demonstrated for the fault detection and isolation of the engine. A real-time verie cation system for the health-monitoring algorithms and system was constructed and applied in the research. I. Introduction B ECAUSE of increasingly stringent requirements for the safety, reliability, and operational capabilities of space vehicles and their launch systems in the past 30 years along with the feasibility provided by the progress of modern science and technology, the health-monitoring techniques for liquid-propellant rocket engines (LRE) have undergone signie cant developments. In earlier stages of development, health-monitoring technology was applied in the ground test process for large-scale LRE. In the 1970s, expendable LREs, such as Atlas and Titan, were monitored by redlines, which are limits or thresholds on some important operating parameters, 1 and automatic test data analysis systems were progressively developed for these engines in early 1990s. 2 The partially reusable rocket engine Space Shuttle Main Engine (SSME), developed in the 1970s, is monitored by a condition-monitoring system that is a simple function of redlines. 3 In the 1980s the SAFD (System for Anomaly and Failure Detection ) was developed for SSME ground tests to improve the monitoring performance of the redline system. 1

Propellant Management and Conditioning within the X-34 Main Propulsion System

Abstract: The X-34 hypersonic flight vehicle is currently under development by Orbital Sciences Corporation (Orbital). The Main Propulsion ystem as been designed around the liquid propellant Fastrac rocket engine currently under development at NASA Marshall Space Flight Center. This paper presents analyses of the MPS subsystems used to manage the liquid propellants. These subsystems include the propellant tanks, the tank vent/relief subsystem, and the dump/fill/drain subsystem. Analyses include LOX tank chill and fill time estimates, LOX boil-off estimates, propellant conditioning simulations, and transient propellant dump simulations.

Advanced Cryogenic Solid Hybrid Rocket Engine Developments: Concept and Test Results

Abstract: : ORBITEC has conducted considerable RD (2) solid hydrogen/gaseous oxygen; (3) solid methane/gaseous oxygen; and (4) solid methane-aluminum/gaseous oxygen. The primary focus of this paper is on solid oxygen/gaseous hydrogen. Work achieved to date includes: (1) a total of over 50 solid oxygen test firings; (2) establishment of regression rate data for the different propellant combinations, where the rates can be a factor of 20 to 40 times higher than conventional HTPB-based hybrids; (3) achievement of burn durations from 1 to 30 seconds; and (4) engine chamber pressures as high as 250 psi The potential applications include. research devices to test high-energy density matter (HEDM); hybrid rocket launch vehicle upper stages; or orbit transfer vehicles. During a current sponsored USAF Research Laboratory (RL, Edwards Air Force Base, CA) project, ORBITEC is to design, develop and test a larger, SOX/LH2 flight-type engine that will have throttling and O/F ratio control.

Transpiration Cooling of Rocket Thrust Chamber with Liquid Oxygen

Abstract: The study on thermal management of platelet transpiration cooling with liquid oxygen is made for the thrust chamber of liquid rocket engine. The analysis of heat transfer is performed using the Barfle and Leadon correlation to determine the heat flux blockage for transpiration cooled wall A heat transfer model for platelet and liquid oxygen in the wall is used to obtain the thermal penetration depth and the temperature profiles of the platelet and the coolant Both steady and unsteady processes being considered, the analytical solution and numerical solution are obtained for them separately. The effects of the blowing ratio on the temperature of the chamber wall are discussed in detail

Experiment and acoustic model for the self-oscillation of coaxial swirl injector and its influence to combustion of liquid rocket engine

Abstract: During the experiment of gas/liquid coaxial swirl injector conducted with air and water under atmospheric environment, it is observed that the injector may self-oscillate The self-oscillation periodically occurs and vanishes with the increasing velocity of the gas flow Then, a theoretical model is presented based on the experiment investigation Simulation of the acoustic process has been performed and conclusions consistent with the experiment can be drawn from the theoretical model, which explains the experimental phenomena quite well At last, the comparison between phenomena of the self-oscillation and some experiments of LRE indicates that some instability phenomena in Oxygen/Hydrogen propellant rocket engine may be the result of self-oscillation in coaxial swirl injectors

Low Cost Propulsion Technology at the Marshall Space Flight Center: Fastrac Engine and the Propulsion Test Article

Abstract: The need for low cost access to space has initiated the development of low cost liquid rocket engine and propulsion system hardware at the Marshall Space Flight Center (MSFC). The engine, the 60,000 lbf, RP-1 and LOX Fastrac Engine has been designed as a robust, low cost liquid rocket engine with applications for X-34 as well as future low cost booster systems. The engine is a turbopump fed, gas generator cycle, rocket motor with an ablative nozzle. The Propulsion Test Article (PTA) is a test bed for low cost propulsion system hardware including a composite RP-1 tank, flight feedlines and pressurization system, stacked in a booster configuration. A general description of the PTA and the Fastrac engine is given, with emphasis on the technical specification of the hardware including flow rates, pressures and other operating conditions. The process which has been used for the design and integration of this hardware is described.

New experimental results on transpiration cooling for hydrogen/oxygen rocket combustion chambers

Abstract: Introduction While previous experimental activities of DLR focused on low pressure applications of combustion chamber segments cooled with either water or gaseous hydrogen wich have been performed at the DLR test facilities P6.1 and M3, this paper presents the most recent experimental results obtained at the high pressure GH^/LO^ test facility P8. For this purpose, DLR model combustor B, a 10 MPa regeneratively cooled experimental combustion chamber having modular design was applied to implement transpiration cooled segments at different downstream positions of the combustion chamber. These recent DLR experiments have been performed for combustion chamber pressures ranging between 3.5 and 10 MPa and a constant oxidizer/fuel mixture ratio of 6.5. Screening tests at lower pressures showed nickel alloys to be favorable for transpiration cooling purposes. Hence, for the high pressure tests reported here only the material porosity was changed with a variation of the typical pore size between 1 /^m and 5 /im. The mass flow rate of the coolant, ambient gaseous hydrogen, has been varied in a wide range to increase the existing data base for modeling and verification purposes as well as to gain experience and to check the limits of operability of this cooling technique especially at the low end of coolant mass flow rates. After a brief description of the governing equations and some available models as well as a short presentation of the experimental setup and the operating conditions, the paper focuses on the experimental results and their comparison with the models derived for transpiration cooling applications. Based on the results of this comparison a modification of the model of Kays and Crawford is suggested. •Member AIAA Copyright ©1998 Erhan Serbest, O.Haidn, K.Frohlke. Published by the American Institute of Aeronautics and Astronautics,Inc. with permission For future space transportation systems, advanced cryogenic combustion chambers are necessary for both Reusable Launch Vehicles (RLV) and Expendable Launch Vehicles (ELV) to meet the requirements for higher payload capabilities. Independent of either ELV or RLV applications, the major requirement for these high pressure combustion chambers is reliability. In future liquid rocket engines with combustion chamber pressures up to 25 MPa thermal loads of the chamber will lay beyond 140 MW/m? in the throat region [1]. Conventional regeneratively cooled combustion chambers made from milled slotted liners of high conductivity copper alloys, will hardly fulfill the requirements for a further increase of both chamber reliability and life. Presently, various candidate techniques such as thermal barrier coatings, improved liner materials, microchannel cooling structures, film cooling via injector trimming and transpiration cooling are discussed and tested by industry and research laboratories. All these methods seem to be means to overcome the problem of the severe thermal loads to the combustion chamber walls. Transpiration cooling, although a rather old idea [2], has not yet been applied for chamber liner cooling purposes in flight engines due to insufficient knowledge about local heat flux distribution, coolant film stability, and coolant mass flow rate controlability. Furthermore, technologies for the production of porous materials having reliable and reproducible parameters are still under development. The benefit of this cooling technique is that plastic deformation of the chamber wall is reduced and hence life time problems due to fatique may be negliable. A problematic issue might be the manufacturing process of the porous material according to the needs in a combustion chamber with varying combustion chamber wall curvature. For investigations in this field there are two facilities available at DLR Lampoldshausen. For basic reasearch there is the M3 Micro Combustor with pressures up to 2 MPa, see [3], [4], [6]. For pressures and mass flow rates the P8, a facility which allows for combustion chamber pressures above 30 MPa and mass flow rates which exceed 11 kg/s, is in use. Governing Equations and Models Reference Heat Flux Without Blowing The Bartz equation (1) in its modified form is a quite standard way for heat flux determination in rocket combustion chambers. Based on this equation a reference heat flux for the non-blowing case is defined which will be later used to evaluate the performance of the transpiration cooling. This reference heat flux is presented in form of a Stanton number StQ

Single-component liquid-propellant rocket engine

Abstract: FIELD: single-fuel space-vehicle control systems. SUBSTANCE: engine has fuel decomposition chamber 1 with bottom 2 accommodating fuel distribution unit 3. The latter has porous catalytic material 4 and blind channel 5 made along its axis and provided with side walls 6 tight through maximum one third of its length from chamber bottom. Electric heater 7 is arranged over external surface of fuel distribution unit. Engine also has tight end surface 8, permeable catalytic stack 9, nozzle 10, fuel supply line 11. Permeable shell is placed between heater and porous catalytic material. Side walls of unit 3 are formed by electric heater turns laid in a spaced relation to each other with clearance of 0.02-0.1 mm. Vortex zone formed near bottom 2 provides for thermal damper and natural suppressor of engine instability in operation. Heated gaseous products of fuel decomposition escape through nozzle 10 and build up reactive thrust. EFFECT: improved speed in starting and improved stability and safety of running engine; simplified design. 3 cl, 1 dwg

Optical diagnostics for cryogenic liquid propellants combustion

Abstract: Detailed experimental studies of cryogenic propellant combustion are needed to improve design and optimization of high performance liquid rocket engines. A test facility called MASCOTTE has been built up by ONERA to study elementary processes (atomization, droplet vaporization. turbulent combustion...) that are involved in the combustion of liquid oxygen (LOX) and gaseous hydrogen (GH2). This article reports results from experiments carried-out on MASCOTTE under a consortium of laboratories and manufacturers associating ONERA, CNRS, CNES and SEP. on the jet flame issued from a single coaxial injector. This device fed with liquid oxygen and gaseous hydrogen is placed in a chamber equipped with quartz windows. The spray and the flame are observed with a set of optical methods : high speed cinematography. light emission from OH radicals, laser induced fluorescence of OH and O 2 , elastic scattering from the LOX jet. These techniques are used to obtain images of the spray and of the flame zone. It is then possible to deduce the flame location with respect to the liquid jet from simultaneous elastic scattering of the LOX jet and LIF of OH measurements or from average emission images treated with Abel's transform. The images obtained by exciting the fluorescence of O 2 provide complementary information on the flame shape and they may be used to estimate the local reaction rate. Quantitative temperature measurements based on Coherent Anti-Stokes Raman Scattering from H2 and LOX droplets size and velocity measurements by means of a Phase Doppler Particle Analyzer give additional clues on the spray and the combustion zone.

Design and Development of a 50K LOX/Hydrogen Upper Stage Demonstrator

Abstract: : This paper discusses design and systems integration of a 50,000 pound (222.4 kN) thrust Oxygen/Hydrogen Upper Stage Engine Demonstrator (USD) being developed by Pratt & Whitney Liquid Space Propulsion under contract for the United States Air Force Research Laboratory (AFRL) to support the Integrated High Payoff Rocket Propulsion Technology (IHPRPT) program. The objective of this program is to integrate advanced technology components into an expander cycle engine configuration and demonstrate a 1% increase in specific impulse, a 30% increase in engine thrust-to-weight, a 25% reduction in failures per 1000 uses, a 15% reduction in required support costs, and a 15% reduction in hardware costs relative to current state-of-the-art levels (RL10A-3-3A). Scheduled to be the first of the IHPRPT program engine demonstrators, it is scheduled to be test fired in late 2000 and demonstrate a chamber pressure capability of 1375 psia. This integrated 50k LOX/LH2 engine demonstrator will be used to evaluate individual component technologies as well as the system level mechanical, structural and thermodynamic interactions.

Thermal/fluid study of perforated plates for transpiration cooled rocket chambers

Abstract: Increasing liquid rocket performance will require use of active cooling techniques. Alabama Cryogenic Engineering, Inc. has developed a drawing process to fabricate metal plates with very small, uniformly distributed holes. These plates could be used for transpiration cooled combustion chambers and nozzle throats. A numerical analysis of the thermal/fluid characteristics of the porous material was developed. A thermal/fluid numerical model was developed to determine the temperature distribution through the thickness of the porous chamber wall. The different aspects that influence the temperature profile within the porous walls of a combustion chamber were studied. These aspects included thermodynamic properties for the coolant and porous material and the effects of blowing.

Method of operation of liquid-propellant rocket engine and liquid-propellant rocket engine used for realization of this method

Abstract: liquid-propellant rocket engines. SUBSTANCE: proposed method of operation of liquid-propellant rocket engine consists in delivery of propellant components to combustion chamber of engine, gasifying one of components in combustion chamber cooling passage, delivery of it to turbine of turbo-pump unit followed by discharge to injector assembly of combustion chamber. Part of one of propellant components is directed to combustion chamber and remaining part is gasified and is directed to turbines of turbo-pump units. Gaseous component used in turbines is mixed with liquid component fed to engine under pressure exceeding pressure of saturated vapor of mixture being obtained. Liquid-propellant rocket engine has combustion with regenerative cooling passage, propellant component delivery pump and turbine. Engine is provided with pump of booster turbo-pump and mixer mounted in succession before component delivery pump of main turbo-pump unit. Outlet of pump of main turbo-pump unit is connected both with injector assembly of combustion chamber and with regenerative cooling passage of combustion chamber. Regenerative cooling passage is connected in its turn with turbines of main and booster turbo-pump units whose outlets are connected with mixer. EFFECT: enhanced reliability; improved power and mass characteristics of liquid-propellant rocket engine. 3 cl, 1 dwg

The effect of transpiration cooling with liquid oxygen on the flow field

Abstract: The study of the effects of platelet transpiration cooling with liquid oxygen on the flow field is made for the liquid rocket engine. Both the thermal analyses for the computation of the temperature distribution in the chamber wall and the combustion gases flow field in the thrust chamber are treated as two dimensional axisymmetric problem. The heat transfer from the hot gas to the wall is determined by the Battle and Leadon s correlation method. Numerical analysis of the turbulent combustion and flow process in transpiration cooled engine has been performed in this paper.

Multiple-action oxygen-hydrogen propulsion system

Abstract: FIELD: aircraft propellers; acceleration units. SUBSTANCE: pneumohydraulic circuit of oxygen-hydrogen propulsion system has receivers- gas generators 15, 14 for gasification of liquid fuel components used in liquid propellant rocket thruster 16. System has two-shaft turbopump set with turbines 2, 3 arranged in tandem. To ignite liquid fuels in chamber of cruise engine 12, hydrodynamic ignition sources are installed in receivers-gas generators and in liquid propellant rocket thruster. System provides periodical operation of reactive control system and multiple cutting in of cruise liquid propellant rocket engine. EFFECT: improved environmental protection, increased efficiency and enhanced reliability of operation. 2 cl, 1 dwgn

The use of vibration data for liquid rocket health monitoring

Abstract: An approach is presented to the construction of a liquid rocket Health Monitoring (HM) system capable of using both, vibratory and thermodynamic data. Preliminary results are obtained in the area of vibration data tracking/identification, using as a test bed a low-order linear dynamic model of the Space Shuttle Main Engine (SSME) High Pressure Fuel Turbopump (HPFTP). Limitations and future extensions are discussed.

Study of a new pasty propellant multi-purpose rocket motor

Abstract: A new method for pulse rocket motor designing is proposed in this paper, on basis of analyzing flow, heat transfer, ignition, combustion and extinction of a kind of nonNewtonian pasty propellant, in which oxidizer and fuel are mixed homogeneously. The new motor is of the advantage of solid propellant rocket motor and liquid propellant rocket motor both. This kind of pulse rocket motor is also thrust adjustable. Moreover, because propellant is ignited by combustion chamber surplus heat, no additional ignition energy is needed. The key technology of this new method is the designing of the re-ignitable igniter. A 2D flow and heat transfer model of propellant inside the reignitable igniter is established, control equations are solved numerically. Temperature distribution of propellant inside re-ignitable igniter during its flow to combustion chamber is obtained, in conjunction with a simple ignition criterion. Relationship among igniter temperature, igniter length, igniter diameter and drive pressure are also obtained. A set of prototype test apparatus is also built for multi-pulse firing. Some seven operation pulses are obtained, each lasts about five seconds, time interval among each pulse is about two second. * Associate Professor, The Fourth Academy of China Aerospace Corporation. ALAA member. + Professor and President, The Fourth Academy of China Aerospace Corporation. Copyright © 1998 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved. Nomenclature A section area of the tube k consistency coefficient L length of igniter tube P drive pressure m mass n propellant flow exponent Q propellant flow mass rate R radius of igniter tube r propellant burning rate T temperature Tig propellant ignition temperature t time tid ignition delay time tf propellant flow time u velocity x, y Catherine coordinate P propellant density /I heat conduct coefficient T shear stress T] propellant viscous coefficient u dynamic viscous coefficient

Revision of the conceptual model of combustion mechanism in a liquid-propellant rocket on the basis of the results of highly informative experiments. I. A review of experimental studies

Abstract: In seeking sources of irregular, stochastic excitation of vibrational combustion in liquid-propellant rocket of various types with the use of a number of independent highly informative experimental methods, a group of phenomena (called parametric instability) that are inconsistent with the traditional conceptual model of the combustion mechanism in given objects was found. Revision of this model is required.

Closed loop cooled rocket engine

Abstract: A space craft's rocket engines are cooled by a recirculating cooling system containing a non-propellant coolant fluid, such as water and/or ethylene glycol. With that recirculating cooling system to maintain the rocket engine combustion chamber at a lower temperature, spacecraft rocket engines may be constructed less expensively and can operate with greater safety by employing the more common metals in their construction. The cooling system also provides an easy means to warm and/or vaporize a propellant.