Showing papers on "Liquid-propellant rocket published in 1997"

AN INTRODUCTION TO PULSE DETONATION ROCKET ENGINES (PDREs)

Abstract: An overview of a Pulse Detonation Rocket Engines (PDREs) is presented. PDREs have the potential to significantly reduce the cost and complexity of today's liquid propellant rocket engines. Practical engineering issues and subsystem technologies are being addressed by ASI to ensure that this potential is realized. This paper describes the important components required for a PDRE, presents some fundamental thermodynamic, design, and operational issues associated with the engine, and discusses several past ASI test articles utilized to provide validation of the PDRE concept.

Liquid oxygen gasifying system for rocket engines

Abstract: For a liquid rocket engine including the SSME the gaseous heat exchanger that is in the liquid oxygen turbopump exit flow path that serves to gasify the liquid oxygen for pressurizing the liquid oxygen tank and the POGO is modified by changing the flow circuit to replace the liquid oxygen with high pressure hydrogen tapped off of the low pressure fuel turbopump so that hydrogen/hydrogen at a favorable pressure differential is in indirect heat relation for providing a safe environment and avoiding what may result in a catastrophic failure in the event the heat exchanger fails. An external heat exchanger in communication with the heat exchanger in the liquid oxygen turbopump is provided to gasify the liquid oxygen for the pressurizing purposes.

Cryogenic liquid rocket engine technology developments within the german national technology programme

Abstract: In a joint technology programme Daimler-Benz Aerospace (Dasa) and the German Aerospace Research Establishment (DLR) are presently investigating technologies for advanced thrust chambers of reusable, high performance cryogenic rocket engines. This Technology Programme on Cryogenic Rocket Propulsion (TEKAN) has been established in 1996 under sponsorship of the German Space Agency DARA. In this programme the extensive experience of Dasa with regeneratively cooled liquid rocket engine combustion chambers is combined with the research oriented activities of DLR in liquid rocket propulsion in order to reach the targets. The present paper describes and discusses the technology developments within TEKAN which have been initiated. Rocket engine cycle analyses for different reusable launch vehicle concepts are performed in order to define the requirements for the thrust chamber components of future launch vehicle propulsion systems. In addition, a simulation tool is developed for transient engine analysis. Basic theoretical and experimental work on combustion chamber heat transfer, cooling, lifetime, and combustion modelling is continued in this programme. Technology developments of the main thrust chamber components as injectors and combustion chamber technologies for high lifetime (thermal barrier coatings, elastic liner technologies) are investigated with experimental test programmes on subscale level and theoretical analyses. Technologies for expander cycle engines are developed and the applicability of carbon fibre reinforced composite materials in cryogenic thrust chambers is investigated.

Non-propellant fluid cooled space craft 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.

Liquid Rocket Engine.

Abstract: PURPOSE:To boost propellant to a high pressure without using a subcombustor by reheating a part of the propellant heated by the heat of a combustor by using a heat exchange means in a nozzle part. CONSTITUTION:The liquid hydrogen as fuel and liquid oxygen as oxidizer are pressurized to each high pressure by boosters 3 and 4. The total quantity of the pressurized liquid hydrogen is allowed to pass through a cooling jacket in a combustion chamber and cools a combustor 6. Therefore, propellant 1 is heated, and almost all parts of the hydrogen is combusted in the combustor 6, and jetted in gas form. While, after passing through the cooling jacket 7 in the combustion chamber, a part of the hydrogen is introduced into a high expansion nozzle cooling jacket 9 for cooling a high expansion nozzle8 and heated to obtain an increased pressure. Therefore, the driven with a sufficient output is permitted by providing the energy of the hidrogen gas at high temperature into the driving device 5 of the booster.

Transpiration cooling using gaseous hydrogen

Abstract: For future liquid rocket engines, advanced cryogenic combustion chambers are needed, both for Expandable Launch Vehicles (ELVs) or Reusable Launch Vehicles (RLVs). As a consequence of higher combustion chamber pressures extreme heat loads to the chamber walls, especially to the throat section, make alternatives to conventional regenerative cooling techniques almost mandatory. Transpiration cooling where a small amount of the fuel passes through porous walls into the combustion chamber is a very promising technique [14], [7]. This contribution summarizes both the experimental and the numerical work in the field of transpiration cooling performed at DLR Lampoldshausen. The experiments presented here have been performed using gaseous hydrogen at ambient temperature as coolant flow. Parallel to the experimental approach, some theoretical and numerical work has been carried out. Transient temperature distributions inside the porous wall are predicted modeling the heat transport problem as onedimensional in cylindrical coordinates. The hot gas side heat transfer is described applying a Bartz equation modified to account for the blowing boundary condition. Furthermore, coolant and wall are handled separately.

A standardized technique for evaluating hybrid rocket motor performance

Abstract: Numerous subscale experiments have been performed in order to examine the impact of certain features on the performance and stability of hybrid rocket motors. Proper evaluation of the recorded test data is necessary to accurately determine the performance of a given test. Under the best circumstances, only a single interpretation of the test data would be possible. A robust analytical technique is needed to ensure that the proper evaluation of performance has been obtained. Due to the nature of hybrid motors, the approaches that have been taken with solid and liquid rocket motors are not sufficient for the classification of hybrid rocket performance. New analytical techniques and a comprehensive analysis approach were developed to assess the large variability in performance that has been seen in the testing of hybrid motors. These techniques are based on industry standard thermochemistry codes and result in a complete assessment of performance for hybrid motors. If a rigorous, unbiased approach is applied, only one interpretation of the data will result. The purpose of this paper is to provide a working definition of the approach and processes that can be applied to the testing of hybrid rocket motors to properly evaluate performance without ambiguity or error.

Recovery system for the evolved expendable launch vehicle

Abstract: Development and testing of a recovery system for the Low Cost Concept Validation (LCCV) phase of the U.S. Air Force Evolved Expendable Launch Vehicle (EELV) program is discussed. This system demonstrated for the first time the recovery of a liquid fueled rocket engine , a Space Shuttle Main Engine (SSME), from altitude to an ocean touchdown with subsequent refurbishment and engine firing. New technology developed and discussed includes the largest three Ringsail parachute cluster of Apollo heritage, in-flight deployable spray shield, and Propulsion Module (PM) testing approaches to validate the integrated concept.

Acoustic igniter and ignition method for propellant liquid rocket engine

Abstract: The invention concerns an acoustic igniter for igniting a mixture of rocket fuels in a liquid propellant rocket engine combustion chamber comprising a cylindrical precombustion chamber (101) including a cylindrical wall (111) and first and second end walls (112, 113), a rocket fuel injection nozzle (103) emerging into the precombustion chamber (101) through the first end wall (112) via an orifice of diameter dn, a rocket fuel injector (104) arranged inside said nozzle (103) along the axis thereof, at least an outlet orifice (102) of minimum diameter df provided in the cylindrical wall (111), an acoustic resonator (105) defining a cavity opening into the precombustion chamber (101) opposite the nozzle (103), through the second end wall (113), via an orifice (151) of diameter dr. The acoustic resonator (105) is enclosed by a housing (106) which defines around the acoustic resonator (105) a closed auxiliary chamber (160) which communicates only with the precombustion chamber (101) by at least one conduit (107).

TRW's Ultra Low Cost LOX/LH2 Booster Liquid Rocket Engine

Abstract: TRW has designed, built and readied for test a 650,000 pound (sea level) thrust, LOX/LH2 liquid rocket engine. The engine design demonstrates a paradigm shift for large booster engines by trading a small decrease in performance for significant reductions in engine manufacturing costs. The result is an engine which can enable a viable and extremely low cost booster rocket. The pump-fed engine design incorporates a relatively low chamber pressure (700 PSIA) which allows the major components (combustion chamber, turbomachinery and feed system) to be much less expensive than those seen in higher pressure engine designs. The engine incorporates TRW's demonstrated pintle injector design, an ablative-lined combustion chamber/nozzle and low cost foil bearing turbopumps which together allow substantial reductions in engine parts count (<100), manufacturing costs, and test costs. This paper describes the engine design, manufacturing and operation while emphasizing its unique features which allow substantial reductions in manufacturing and tests costs. Finally, the paper describes the full scale test hardware, which was designed, manufactured and assembled in TRW's Redondo Beach, CA facilities in less than 12 months from program inception.

A rocket engine for Mars sample return using in situ propellants

Abstract: Recent studies for the planned Mars sample return mission were reviewed and modified to utilize carbon monoxide and oxygen as potential in situ propellants. Based on these studies a representative full scale engine thrust of 2225 N (500 lbf) was selected as appropriate to demonstrate performance, and the design for that engine is presented. Previous experimental results combined with parametric analyses were used to define the geometry for the engine which operates on liquid carbon monoxide and liquid oxygen. The engine was constructed using a combination of high-temperature alloys and lightweight ceramics. The materials selected were hafnium oxide, iridium, rhenium, and carbon-carbon.

Laser ignition for liquid propellant rocket engine injectors

Abstract: A laser beam is utilized to ignite the fuel and oxidizer injected into a combustion chamber. An injector has a plurality of injector elements each injecting fuel and oxidizer into a flame holding zone adjacent the injector face plate where fuel and oxidizer are mixing at a rate to sustain flammability. The laser beam is introduced parallel to the injector face plate passing through the fuel and oxidizer mixing zones igniting the propellants. The propellants are ignited by having a laser beam tuned to a frequency to excite the oxidizer such that it will chemically combine with the fuel to form an ignition kernel in the flame holding zone. When a plurality of injector elements are ignited in this manner a controlled ignition process in the combustion chamber is achieved.

Numerical Prediction of Pressure Distribution Along the Front and Back Face of a Rotating Disc With and Without Blades

Abstract: This paper presents a numerical method for developing correlations of the K-factor (fluid to solid rotational speed ratio) for rotating discs from measured test data. The correlations were developed for the front and back side of a rotating disc with and without blades. Separate correlations were developed for water and liquid hydrogen. It was found that a more accurate agreement with test data was obtained when the K-factor was assumed to have a linear relationship with radius instead of assuming a constant value. The two coefficients of the linear correlation were evaluated by an iterative method based on the Newton-Raphson scheme. An excellent agreement with the test data was obtained when these correlations were used in a numerical model built with the help of the Generalized Fluid System Simulation Program (GFSSP). These correlations are recommended for use in the design and analysis of liquid rocket engine turbopumps.

SSME Key Operations Demonstration

Abstract: A Space Shuttle Main Engine (SSME) test program was conducted between August 1995 and May 1996 using the Technology Test Bed (TTB) Engine. SSTO vehicle studies have indicated that increases in the propulsion system operating range can save significant weight and cost at the vehicle level. This test program demonstrated the ability of the SSME to accommodate a wide variation in safe operating ranges and therefore its applicability to the SSTO mission. A total of eight tests were completed with four at Marshall Space Flight Center's Advanced Engine Test Facility and four at the Stennis Space Center (SSC) A-2 attitude test stand. Key demonstration objectives were: 1) Mainstage operation at 5.4 to 6.9 mixture ratio; 2) Nominal engine start with significantly reduced engine inlet pressures of 50 psia LOX and 38 psia fuel; and 3) Low power level operation at 17%, 22%, 27%, 40%, 45%, and 50% of Rated Power Level. Use of the highly instrumented TTB engine for this test series has afforded the opportunity to study in great detail engine system operation not possible with a standard SSME and has significantly contributed to a greater understanding of the capabilities of the SSME and liquid rocket engines in general.

Simulation of spray combustion instabilities in liquid rocket engines. I - Methodology

Abstract: This paper is the first of a two-part paper describing the development and application of models for physical processes such as atomization, vaporization and combustion to predict combustion instabilities in liquid rocket engines. Previous studies have shown that unsteadiness in atomization, vaporization and heat release can cause the growth of combustion instabilities. The strong coupling between these processes provides a feed-back loop that can sustain these instabilities. The objective of this study is to develop efficient timeaccurate numerical techniques to model all the physical processes in a fully coupled manner. Novel, computationally efficient techniques have been developed to simulate atomization of coaxial as well as impinging injectors and spray combustion processes. These models are described in detail in this paper. Results for impinging jets atomization characteristics, liquid sheet and droplet interactions with acoustic fields and ignition characteristics of premixed H2-C>2 are presented. The second part of this paper presents the results of combustion instability study obtained from the integrated model.

Space round trip flying vehicle

Abstract: The present invention relates to a space round-trip flying vehicle consisting of a large flying body and a small flying body. Both flying bodies are connected by means of connecting-separating device. Each flying body is equipped with a liquid rocket engine and a propellant storage box, and its tail portion is equipped with a turbo-fan engine. Said invention possesses the advantages of good repeated usability, convenient application and maintenance, strong maneuverability, emergency return capability, wide reentry corridor, good reentry thermal environment and strong autonomous cruising power.

A simplified design methodology for coaxial swirl injector

Abstract: Gas-liquid coaxial swirl injectors are widely used in advanced hydrogen/oxygen rocket engines and hydrogen/oxygen/hydrocarbon tripropellant engines. The design quality of injector is very important to the performance, stability, reliability and life-span of the liquid rocket engine. A simplified flowrate characteristics based methodology for the design of coaxial swirl injector is presented in this paper. The design methodology of coaxial injector is based on the flowrate theory of liquid centrifugal injector and gas dynamics. Owing to the correlative experiment and theoretical analysis, the interaction between the liquid and gas phases in the injector is relatively small with common recess ratio values. The simplified approval to design the coaxial injector is to calculate the fluxes of liquid and gas phases respectively, and then revise them by experiments. The liquid flux is firstly computed by the flowrate theory of classic liquid centrifugal injector, and then calibrated through the flux coefficient ///. While the gas flux is computed by the gas dynamics of annulus passage according to its flow state. A modify coefficient //, is necessary because of interaction between the central liquid flow and the outer gas jet, and of the influence of gas viscosity. The modify coefficient should be chosen through comparison between theoretical and experimental results on flowrate characteristics. The methodology has been verified by cold experiment using air and water as analogy media, and been applied to the design process of a GO2/kerosene rocket engine. The results of the direct GO2/kerosene cold experiment and the hot tirt test of that GO2/kerosene rocket engine show the design is successful. 1. The coaxial swirl injector The schematic of the investigated injector is shown in Fig.I . It is widely used in the advanced Lox/LH2 rocket engines and the tripropellant engines for next generation space mission. It consists of a liquid central tube and a gas annulus passage. The swirling liquid flows out of the central tube and the gas jet injects through the annulus. Compared with other injectors, the coaxial swirl injector has the characteristics of finer spray, shorter atomization distance and more uniform mixing. There exists interaction between the gas jet and the liquid flow of the injector. The extent of this kind of interaction depends on the configuration parameters of the injector and the injecting working parameters such as injecting pressure drops. Fig. 1 Schematic of coaxial injector The recess ratio (RR) is an important special configuration parameter of the coaxial injector. It is defined as the ratio of the recess length of the central liquid nozzle over the entrance diameter D// of the outer gas annulus:

Application of laser-induced fluorescence (LIF) to liquid-propellant rocket engine testing

Abstract: Combustion exhausts present a challenging problem for researchers due to the extremely harsh environment, and non-intrusive diagnostics are often sought to provide flow property information. Laser-induced fluorescence (LIF) is one technique in which a chosen flow molecule or marker is probed to yield gross flow properties, such as static temperature and flow velocities. The work presented herein describes the application of LIF to the combustion exhausts of several full-scale liquid-propellant rocket engines spanning a wide range of operational parameters. The method is based upon the use of cw ring-dye lasers which scan in frequency over either the Na D 1 or D 2 line at 5896 and 5890 A. Na is used as a basis for this approach since it occurs as a trace element in both hydrogen and amine rocket fuels. The generic apparatus is described, including a discussion of the collection and interpretation of the LIF signal to yield radial and temporal profiles of radial flow velocity, static temperature, and fuel distribution. It was found that the LIF technique provides quality data in most cases. Certain stressing situations were also found in which data on the flow properties were not obtainable. Also, computational fluid dyanamics (CFD) modeling of the plumes was used to provide baseline estimates of the exhaust flow properties. The model reasonably predicted the gross behavior of the flow as determined by the LIF technique, although some items of fine spatial structure were not reproduced very well.

Simulation of spray combustion instabilities in liquid rocket engines. II - Application

Abstract: A study is presented on simulating spray combustion in liquid rocket motors. The present goal of this work is to validate the interaction of several physical submodules which include flow solver, spray physics, and chemistry. A brief description is given for each submodule. A two dimensional experimental study is used to help verify the code. Velocity and pressure results are given for RP1/LOX as the fuel and oxidi/er. Spray combustion instability results are given for mono-disperse sprays using the same two-dimensional geometry. The conclusions from this work indicate that spray dispersion in combination with decreasing droplet size can trigger self-sustaining instabilities. INTRODUCTION Over the last four decades, the problem of spray combustion instability has been of continuing concern. Several liquid rocket engine designs, such as the F-l, J2, and J-2S engines', incurred large over runs in time and money due to the presence of spray combustion instability. During this time span, several mechanisms were identified as the potential cause of instability. These mechanisms are atomization, secondary droplet break-up, vaporization, turbulent mixing, chemical reactions, and droplet temperature thermal lag. One or a combination of these factors may be important in initiating or sustaining combustion instability. Efforts were made to experimentally analyze the mechanisms that lead to instability. Lambiris, Combs, and Levine'conducted a series of experiments to gain a better understanding of spray characteristics. Also, these researchers attempted to numerically model the steady-state processes in a two-dimensional combustor with a one-dimensional methodology. As computers have increased in speed and memory, new * Project Engineer, Member AIAA ** Senior Engineer, Member AIAA Copyright 1997 by the Authors. Published by AIAA with permission. algorithms have been used in an attempt to simulate the process of spray combustion and related instabilities. Although a large number of numerical studies have been conducted, only a few will be referenced to give a brief overview of previous efforts. Grenda et. al used a CFD procedure that was applicable to two-dimensional or axisymmetric configurations. They considered only the Euler equations for the gas phase. This approach was selected since resolving the important viscous regions around an injector are memory prohibitive. Also, the viscous boundaries at the combustor walls do not greatly affect the propagation of waves due to instabilities. The spray droplets were tracked in a Lagrangian fashion. Similarly, Litchford and Jeng studied the unsteady, Euler equations for gas and the spray was treated in Lagrangian fashion. They also focused on two-dimensional or axisymmetric geometries. A more complex algorithm was created by Liang and Ungewitter . Their model included the effects of the three-dimensional Navier-Stokes equations with turbulence taken into account. The spray model was based on the Eulerian Volume of Fluid (VOF) method. Droplet fragmentation and coalescence was permitted. Finally, Habiballah and Dubois developed the Phedre code to describe unsteady, turbulent, two-phase flow. An Eulerian-Eulerian approach was used to model the spray and gas. Only two-dimensional cases were considered in the referenced report. These two-dimensional algorithms provide insight into spray combustion instabilities, but only a three-dimensional methodology can treat all of the unstable modes in a liquid rocket engine. The goal of the present work is to develop a fast, general, three-dimensional methodology to study spray combustion instabilities. Due to the nature of combustion chamber geometries and the modes of instability, the solution algorithm should be threedimensional in order to capture the major modes such as first tangential, second tangential, and first radial. The algorithm must also be transient in nature. To simulate American Institute of Aeronautics and Astronautics realistic rocket motors, the code must be able to model 10's to 100's of spray injectors and track the released droplets. These requirements are not difficult to include in a numerical model. However, the code can become memory and CPU intensive. The focus of this paper is two fold. First, the spray combustion model is verified against experimental data. The spray model consists of various submodules that have been developed and tested'''. These submodules include flow solver, injector atomization, spray physics, and chemistry. To validate the interaction of the various submodules, a comparison is made to the experimental results of Lambiris, Combs, and Levine''. Finally, the transverse instabilities of a twodimensional combustor are studied. SOLUTION METHODOLOGY The present numerical method allows for the interaction between the surrounding gas, spray droplets, and chemical reactions. The model uses an EulerianLagrangian approach to track the gas and spray, respectively. The gas, spray, and chemistry submodules interact through source terms in the governing equations and are also linked by a common database. The final results of this model are based on a fractional step technique with operator splitting. Flow Model The main submodule is the flow algorithm which solves an extended set of Navier-Stokes equations. These equations consist of the conservation of momentum in 3D, conservation of total energy, and the conservation of mass for each chemical species. Appropriate diffusion and source terms are included to account for vaporization of spray droplets and subsequent reactions. The gas phase equations can be written in the general conservation form of computed separately.

High Pressure Combustion Studies Under Combustion-Driven Oscillatory Flow Conditions

Abstract: : Rocket engines fueled by a dense propellant such as kerosene provide a number of advantages over hydrogen-fueled engines for primary stages. A major problem in the development of liquid fueled rocket engines has been the occurrence of combustion instability. The lack of a detailed understanding of how combustion instability occurs in liquid-fueled rocket engines has resulted in costly engine development programs that must be avoided in the future. The present research program examined the specific effects of atomization in combustion instability. The effects of mean drop size, drop size distribution, and atomization periodicity were examined explicitly with a combustion response model, the results from which indicated that all of these effects were important. It was shown that periodic atomization, in particular, results in large variations in the magnitude of the response when the atomization frequency is on the same order as the acoustic oscillation frequency. Experimental results from a sub-scale rocket combustor that used electro-mechanically forced atomization to accentuate the natural frequency of periodic atomization associated with impinging jet injectors were also undertaken. The presence of forced longitudinal modes, corresponding to the forced atomization frequencies, substantiate the importance of periodic atomization. A conceptual model of this potentially dominant mechanism of combustion instability was also developed as part of the study.

Aerodynamic performance test results of single stage oxidizer turbine with volute manifolds

Abstract: Test results from a rocket turbine test model, called the Oxidizer Technology Turbine Rig (OTTR), are discussed in this paper. The turbine was designed to support the development of advanced turbines for future liquid rocket engines. It is a highly loaded single stage liquid oxygen pump drive turbine which uses inlet and exit volutes to provide optimum performance in a compact configuration. The system design creates high pressure and temperature gradients as well as high Mach number flow. These factors make it especially difficult to accurately measure the flowfield. Test issues such as probe calibration, probe interference, rake blockage, and averaging techniques were discussed in a previous paper. Test results including volute and diffuser static pressure distributions, stator airfoil static pressure distributions, total and static pressure drops through the system, and overall performance parameters at the turbine aerodynamic design point are presented here.

Propellant emergency discharge device for liquid propellant rocket

Abstract: PROBLEM TO BE SOLVED: To safely discharge a propellant in a short time by discharging the propellant from an airframe tank by ejector effect so as to be sucked out when the original valve of the airframe tank is opened simultaneously with the blowing of an inert gas into an inert gas blowing part provided near the end part of a discharge pipe. SOLUTION: An end part 11a of a propellant emergency discharge device 11 having a diameter larger than the diameter of a filling port 3a is pressed onto the filling port 3a. When an inert gas is blow into an inert gas blowing pipe 11b toward discharging direction by an inert gas supplying means in this state, and the valve 3c of the filling port 3a is opened by remote control, the propellant within a tank is discharged is discharged so as to be evacuated by ejector effect. Even when a slight clearance is present between the filling port 3a and the end part 11a of the propellant emergency discharge device 11, the propellant is safely discharged by the ejector effect without being leaked to the outside.