Showing papers on "Liquid-propellant rocket published in 2001"

Microfabrication of a high pressure bipropellant rocket engine

Abstract: A high pressure bipropellant rocket engine has been successfully micromanufactured by fusion bonding a stack of six individually etched single crystal silicon wafers. In order to test the device, an innovative packaging technique was developed to deliver liquid coolant and gaseous propellants to the rocket chip at pressures in excess of 200 atm at temperatures above 300°C. Testing continues on the 1.2 g devices, which have been run to date at a chamber pressure of 12 atm, generating 1 N of thrust, and delivering a thrust power of 750 W.

Interaction of Acoustic Waves with a Cryogenic Nitrogen Jet at Sub- and Supercritical Pressures

Abstract: : To better understand the nature of the interaction between acoustic waves and liquid fuel jets in rocket engines, cryogenic liquid nitrogen is injected into a room temperature high-pressure chamber having optical access on its sides. A piezo-siren capable of generating sound waves with an SPL of up to 180 dB is used under three chamber pressures of 1.46, 2.48, and 4.86 MPa. The reduced pressures for these pressures are 0.43 (subcritical), 0.73 (near-critical), and 1.43 (supercritical), respectively. The assembly consisting of the acoustic driver and the high- pressure chamber form a cavity that resonates at several frequencies, the strongest being at 2700 and 4800 Hz. Three different flow rates are considered and the nature of the aforementioned interaction has been documented via a high-speed imaging system using a CCD camera. It is found that the impact of the acoustic waves on the jet structure is strongest from low to near-critical chamber pressures and at low injectant flow rates. No significant effects of the acoustic waves are detected at the supercritical chamber pressure examined. It suggests that engine operation either near the critical point or in transition passing through the critical point could be troublesome and may lead to or feed combustion instabilities in liquid rocket engines. Further work is needed to directly relate these effects to the observed instabilities.

Metal Matrix Composites for Liquid Rocket Engines

Abstract: This article presents an overview of current research and material requirements for metal-matrix composite (MMC) technologies being developed for liquid rocket engines (LRE) Developments in LRE technology for the US Air Force are being tracked and planned through the integrated high payoff rocket propulsion technologies program (IHPRPT) Current efforts and research requirements for three types of MMC systems are discussed: aluminum-, copper-, and nickel-matrix material systems Potential applications include turbopump housings, rotating machinery, and high-stiffness flanges and ductwork

Multi-mode multi-propellant liquid rocket engine

Abstract: A multi-mode multi-propellant rocket engine capable of operating in a plurality of selected modes. G f  ( K X + 1 )  C * A *  P 0 = 1 Propellant components may include liquid hydrogen, liquid hydrocarbon, liquid oxygen, liquid fluorine, and liquid air. The liquid oxygen and the liquid air are stored in separate tanks are mixed in a dedicated mixer prior to their injection into the combustion chamber.

Optimization of Geometric Parameters of Cryogenic Liquid Rocket Combustion Chamber

Abstract: A method for the optimization of rocket combustion chamber walls with respect to the life time is presented. This method can be split into four main parts: P1) Determination of the thermal field within the combustion chamber wall and the cooling channel during the hot run phase by a steady state thermo-fluid mechanical analysis. P2) Analysis of the nonlinear deformation of the combustion chamber wall under cyclic thermal and mechanical loading. P3) Estimation of the life time of the combustion chamber wall by a post processing method. P4) Application of a mathematical optimization procedure. This strategy is used to analyse the thermal load induced deformation process and life time of a typical rocket combustion chamber and to optimise selected geometry parameters of the combustion chamber wall. As one of the objectives of the presented work is a recommendation for the choice of a suitable and efficient optimization procedure for the given problem class, two different methods are compared: A standard Conjugate Gradient method and an efficient gradient free optimization procedure.

Investigation of an Alternative Geometry Hybrid Rocket for Small Spacecraft Orbit Transfer

Abstract: : The following testing was carried out in support of the Surrey Space Centre (SSC) Alternative Geometry Hybrid Rocket research and development program. Although VFP testing was conducted in a low cost environment, the research program collected a wealth of valuable data with regard to this all-new hybrid rocket engine. The combustion efficiency is outstanding within the VFP. The scalability test has demonstrated that the VFP indeed scales well, providing high performance over the regimes tested as well as reliable, predictable, fuel liberation based upon the engine radius. The chamber pressure mapping did not reveal any pressure gradient across the diameter of the VFP rocket engine over the regimes tested. Flight propellant testing was promising in a number of areas. The VFP has demonstrated the ability to operate smoothly for long durations (up to 45 seconds tested), and return to within 1.5% of operational values upon relight (pulsed operations). It is likely that the engine may be burned near completion without the fear of solid fuel slivers blocking the rocket nozzle. The VFP test campaign provides solid evidence that the VFP is superior to conventional hybrid design in almost every respect and holds great promise for small spacecraft applications.

Mathematical modeling of droplet heating, vaporization, and ignition including detailed chemistry

Abstract: A two-dimensional computer code for the simulation of heating, vaporization, ignition, and subsequent combustion of cold droplets injected in a hot uniform gas flow is developed. The numerical simulation of the processes is performed in an axisymmetric configuration for spherical droplets with boundary fitted grid point systems. Detailed models for the relevant processes are employed; in particular, detailed chemical reaction systems are used. Both methanol droplets in hot air as well as liquid oxygen droplets in gaseous hydrogen are studied. The first chemical system is relevant in diesel engine combustion, and typical conditions at 30 bar are investigated. The liquid oxygen/hydrogen system is studied at 10 bar, and the inlet liquid oxygen (LOX) droplet temperature is cryogenic (85 K)—this condition is found in liquid rocket propulsion. The computer code accounts for physical properties in this temperature range through addition of a database for LOX/hydrogen to the commonly used NASA polynomials in the ...

Thrust Chamber Technology Developments for Future Launch Vehicle Liquid Rocket Engines

Abstract: At Astrium, Space Infrastructure Division (SI), enabling technology developments for next generation semi-reusable and reusable liquid rocket engines are being undertaken. This paper summarises the present status and the main results of these technology developments for the following main thrust chamber components:

CFD Simulation of Liquid Rocket Engine Injectors

Abstract: 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 these investigators to be very valuable for code validation because combustion kinetics, turbulence models and atomization models based on low pressure experiments of hydrogen air combustion do not adequately verify analytical or CFD submodels which are necessary to simulate rocket engine combustion. We wish to emphasize that the simulations which we prepared for this meeting are meant to test the accuracy of the approximations used in our general purpose spray combustion models, rather than represent a definitive analysis of each of the experiments which were conducted. Our goal is to accurately predict local temperatures and mixture ratios in rocket engines; hence predicting individual experiments is used only for code validation. To replace the conventional JANNAF standard axisymmetric finite-rate (TDK) computer code 2 for performance prediction with CFD cases, such codes must posses two features. Firstly, they must be as easy to use and of comparable run times for conventional performance predictions. Secondly, they must provide more detailed predictions of the flowfields near the injector face. Specifically, they must accurately predict the convective mixing of injected liquid propellants in terms of the injector element configurations.

Liquid rocket engine test plan comparison

Abstract: Tests quantity, number of Engines fired, total cumulated firing time performed during development are key factors for reliability estimation, but above all, for Launcher customer trust. In other respect schedule and development cost are first order dependant of the Engine test plan: the difference between a high number of tests and a low rate test plan may be of about two years which is tremendous in front of the commercial satellite market trends. The paper compares the test plan of the major liquid rocket engine: Fl, J2, SSME, LE-7, LE-7A, VULCAIN 1 and 2, VINCI, RD-180, RS-68, HM7, LE-5A and LE-5B.

CFD Simulation of Liquid Rocket Engine Injectors. Part 2. Simulations of the RCM-2 Experiment

Abstract: : The sub-critical combustion case, RCM-2, was simulated with both heterogeneous and homogeneous spray combustion models. The MASCOTTE test data should be better than any which have been previously used to tune the several parameters in these models. It is unreasonable to expect that spray flames, even of hydrogen and oxygen, can be accurately predicted without extensive model validation with test data representative of the conditions which exists in rocket engine combustion chambers. Even global data such as chamber pressure and thrust have not been obtained for single coaxial element combustor flows. The IWRCM data provide a good starting point, but no CFD model tuning has yet been attempted for such experiments. Direct comparisons of predictions to test data at this point will not establish which of several modeling techniques is best.

Numerical Analysis for Propellant Management in Liquid Rocket Tank

Abstract: In order to track three-dimensional behavior of liquid surface in storage tanks on orbit, a numerical scheme based on the Level Set Method was developed. For the verification of the numerical methods, experimental data were acquired through the observation on the unsteady deformations of liquid surface in cylindrical containers under low-gravity conditions in a drop tower. Main concern was placed on the relation between the value of dynamic contact angle and the behavior of the contact line on the solid wall. Based on the results, the boundary condition for surface tension was discussed and the model of wetting phenomena was adequately introduced into the computation. Compared with the experimental data, the corresponding numerical results obtained with the wetting model showed a quite good agreement. The flow fields at the draining process in the LOx tank for the LE-5B engine under low-gravity conditions were also investigated with the developed code. It was found that the buoyancy induced by a slight acceleration was efficient to prevent the dip growth, and that the serviceable propellant in the launch-vehicle tank could thereby be increased in a realistic situation. Introduction With the progress of human activities in space, the occasion to handle liquids under low gravity conditions is now growing. In weightless flights, the absence or diminution of gravity force makes it extremely difficult to position and control two-phase fluid in a desirable manner'. For the establishment of the technology for fluid * Graduate Student ** Associate Professor t Deputy Project Manager Copyright © 2001 The American Institute of Aeronautics and Astronautics Inc. All rights reserved. management in space, it is essential to accumulate technical knowledge to give appropriate assessment of designed fluid management systems for space application. However, in the atmosphere, there are not so many opportunities to realize the low-gravity state with airplanes or drop towers. The investigation methods with CFD (computational fluid dynamics) are therefore strongly desired. In the present paper, the algorithm of, what is called, CEP-LSM was developed to simulate threedimensional behavior of liquid surface driven by surface tension, wetting phenomena and gravity force. The free-surface flows under low gravity conditions were both experimentally and numerically investigated to verify the algorithm and to study the appropriate boundary conditions. Drop Tower Experiment In the present study, the unsteady deformations of liquid surface under low-gravity conditions were observed through a transparent vessel of cylindrical shape. The series of experiments aimed at acquiring the basic knowledge of free-surface flows driven by surface tension, and at obtaining the data suitable for the verification of the CFD code and the discussion on the boundary condition. Experimental Facilities The present experiments were conducted at the drop tower, shown in Fig.l, constructed in the University of Tokyo. As is shown in Fig.2, the drop box was composed of an inner box including the test vessel made of poly-acrylate resin, and an outer box of 920 mm wide, 610 mm deep, and 700 mm high in dimension. Liquid behaviors were observed through transparent wail of the test vessel with the CCD camera equipped on the observation section and recorded on a VCR. 1 American Institute of Aeronautics and Astronautics (c)2001 American Institute of Aeronautics & Astronautics or Published with Permission of Author(s) and/or Author(s)' Sponsoring Organization.

Liquid-propellant rocket engine combustion chamber mixing head

Abstract: aeronautical engineering. SUBSTANCE: mixing head of combustion chamber has housing with bottom and fitted-on bushings whose clearances form ring channels for delivery of liquid and gaseous components. Proposed mixing head has mixing members secured on outlet ends of bushings. Each mixing member consists of two concentrically interconnected rings. Two screw cuts forming a collector in between are made on one of inner surfaces of rings. Collector is connected by inclined holes with channels to deliver gaseous component. Inclined holes are made on opposite walls of gas channel being uniformly displaced relative to each other. Throttling holes are made at inlet of gaseous component channels. Such design of mixing head provides complete combustion of fuel components in chamber of liquid-propellant rocket engine at stable combustion process at all loads and stability of construction to thermal action of fuel combustion products. EFFECT: improved conditions of fuel combustion. 4 dwg

Combined Engine for Reusable Launch Vehicle (KLIN Cycle)

Abstract: Predicted performance and features of a combinedpropulsion concept for a small reusable launchvehicle known as KLINTM (meaning wedge in Russian) Cycle are discussed. The KLIN Cycle consists of a thermally integrated deeply cooled turbojet and a liquid rocket engine. The objective of this concept is to achieve a high-pressure ratio in a simple, lightweight turbojet engine. The proven result is an exceptional engine thrust-to-weight ratio, as well as improved speciŽ c impulse and mass fraction of the launcher. When based on the RL10 engine family, the KLIN Cycle makes a small single-stage-to-orbit and two-stage-to-orbit reusable launch vehicles feasible and very economically attractive.

Simulation of a liquid rocket engine

Abstract: A study has been carried out of the dynamic behaviour of the liquid propellant that circulates through the different engine components of a liquid.fuelled rocket. As a model we used the Vulcain engine, the main engine of the European Ariane rocket, which is fuelled by liquid hydrogen and oxygen. For this purpose, a library has been created in EcosimPro which includes the rocket components, as well as auxiliary functions to calculate the physical properties of the substances (both reagent and product) which take part in the propulsion process (combustion).

A Three Dimensional Parallel Time Accurate Turbopump Simulation Procedure Using Overset Grid Systems

Abstract: The objective of the current effort is to provide a computational framework for design and analysis of the entire fuel supply system of a liquid rocket engine, including high-fidelity unsteady turbopump flow analysis. This capability is needed to support the design of pump sub-systems for advanced space transportation vehicles that are likely to involve liquid propulsion systems. To date, computational tools for design/analysis of turbopump flows are based on relatively lower fidelity methods. An unsteady, three-dimensional viscous flow analysis tool involving stationary and rotational components for the entire turbopump assembly has not been available for real-world engineering applications. The present effort provides developers with information such as transient flow phenomena at start up, and non-uniform inflows, and will eventually impact on system vibration and structures. In the proposed paper, the progress toward the capability of complete simulation of the turbo-pump for a liquid rocket engine is reported. The Space Shuttle Main Engine (SSME) turbo-pump is used as a test case for evaluation of the hybrid MPI/Open-MP and MLP versions of the INS3D code. CAD to solution auto-scripting capability is being developed for turbopump applications. The relative motion of the grid systems for the rotor-stator interaction was obtained using overset grid techniques. Unsteady computations for the SSME turbo-pump, which contains 114 zones with 34.5 million grid points, are carried out on Origin 3000 systems at NASA Ames Research Center. Results from these time-accurate simulations with moving boundary capability will be presented along with the performance of parallel versions of the code.

Maglev for space launchers

Abstract: The Magnetic Levitation (MAGLEV) techniques applied to space launch vehicles acceleration is described in this paper. An overview of the concept plus a world-wide survey of the most significant current research in the field (and in neighbouring domains) is carried out. A preliminary assessment of cost-effectiveness and performance evaluation of the system is presented by comparing the magnetic levitation to liquid rocket engine stage.

Performance of a Titan Rocket Engine Using Laser-Induced Fluorescence of OH

Abstract: Rocket engine performance can be modeled by considering separately the propellant delivery system distribution, combustion efe ciency including propellant vaporization and gas-phase mixing, and nozzle expansion efe ciency. Although these quantities can be modeled, experimental verie cation is extremely helpful for separately understanding these processes and for design improvements. Laser-induced e uorescence of OH, excited by a KrF excimer laser operating at 248 nm, is used to measure the concentration of the OH radical across the exit plane of a e ring Titan IV, stage I, liquid rocket engine, from which the combusting mixture ratio proe le could be inferred. Thesemeasurementsallowassessmentofthedegreeofmixing and potential e owstratie cationbetween theinjector core, combustion bafe es, and combustion chamber fuel-e lm cooling and can help to provide the basis for future performance optimization. Nomenclature A = Einstein A coefecient Aa = area of the laser beam (height times thickness ) B12 = Einstein second coefe cient for stimulated absorption c = speed of light E = laser energy per pulse fB.T/ = temperature-dependent Boltzmann fraction of the absorbing state g.o/ = spectral overlap function ho = energy of a scattered photon K = non-noise-free gain factor M = magnie cation of the imaging system Nc = number of counts recorded by the camera per pixel Np = number of laser pulses integrated NT = total number density of the gas Pc = rocket engine combustion chamber pressure Qpre = predissociation rate S = e uorescence signal T = gas temperature uradial = radial component of the velocity V = collection volume ´ = collection efe ciency ÂOH = mole fraction of OH A = collection solid angle per pixel ! = wave number of the laser Subscript pp = per pixel

Adjustable liquid rocket engine

Abstract: The invention relates to rocket engineering, more specifically to rocket engines using liquid fuel components The inventive liquid rocket engine comprises a chamber (1), a gas generator (2), a pump-turbine unit provided with centrifugal pumps (6,7,8), a turbine (5) and an automatic system provided with valves (9,10,11) and flow control valves (19,20) Said engine also comprises high-pressure fuel lines and low-pressure fuel lines At least two high-pressure fuel lines comprise electric control valves actuated by corresponding electrical commands The inventive engine is characterised in that it is provided with the electric control valves and units for command formation (70) including the commands sent to the electric control valves At least two high-pressure fuel lines are embodied in such a way that they are bifurcated into parallel branches which have a common input and output for each line Each flow control valve is embodied in the form of a hydraulic valve (39,40,41,42,43,44,45,46) arranged in the corresponding branches and actuated by the electric control valves (55) etc

Rocket propulsion drive method uses injection of water into gas plasma for providing water vapor discharged via rocket engine jets

Abstract: The rocket propulsion drive method uses water injected into a rare gas plasma, e.g. an argon or helium plasma, with the resulting water vapor discharged via the jets of the rocket engine. The ratio between the heat content of the plasma and the water injection can be regulated for maintaining an exit temperature of between 1500 and 3000 at the rocket engine jets. An Independent claim for a rocket engine is also included.

Method of organization of working process of liquid-propellant rocket engine and liquid- propellant rocket engine for realization of this method

Abstract: liquid-propellant rocket engines. SUBSTANCE: proposed method includes preheating effected due to supply of thermal power of thermocatalytic decomposition chamber, delivery of liquid propellant to end section of chamber through propellant delivery unit, decomposition of rocket propellant in thermocatalytic pack and efflux of decomposition products through gas-dynamic nozzle. After delivery of propellant to chamber, it is preliminarily evaporated and propellant vapor is fed to thermocatalytic pack. Liquid-propellant engine has decomposition chamber with thermocatalytic pack, propellant delivery unit adjoining the chamber bottom and gas-dynamic nozzle. Mounted in chamber between propellant delivery unit and thermocatalytic pack is evaporator made from catalytically inert permeable conducting material with current leads. EFFECT: improved specific characteristics at low thermal losses; increased service life; enhanced reliability and operational stability. 7 cl, 3 dwg

Leak fault identification of rocket engine using self organizing feature map network

Abstract: Some kinds of leak fault were analyzed in liquid rocket engine. The method of principle component analysis was used to reduce the dimension of the original samples, which can represent the leak. And then input the low dimension samples, the self organizing feature map network can identify the leak fault. The simulating results show that the approach is feasible and effective.

Approach to dynamic data processing for liquid rocket engine based on wavelet analysis

Abstract: In order to process dynamic data of liquid propellant rocket engine , an approach using wavelet analysis was studied. Based on the singularity detection and error filtering theories of wavelet analysis, the process of dynamic data processing was analyzed and validated by the dynamic data coming from one ground fire-test of a liquid propellant rocket engine. The results demonstrate that the wavelet analysis method is very fit for dynamic process of engine. It does not require the prior knowledge of process error and also can fulfill the tasks of singular data detection and corresponding error filtering simultaneously.

Nonlinear dynamic neural network model for rocket propulsion systems

Abstract: It is not only very essential for control system design but also for failure detection and diagnosis to set up a real time, precise and reliable dynamic model of liquid propellant rocket's propulsion system. The feed forward neural network if successfully trained, can map the inputs to the desired outputs, so recent years have seen an extensive amount of research to explore its approximation properties. On the basis of studying RBF (Radial Basis Function) neural networks' theory and system mechanism, a nonlinear dynamic neural networks' model with multi inputs and multi outputs for liquid propellant rocket's propulsion system was built. During the modeling, necessary dynamic information was included and parameters of model were also well chosen. The contrastive results of outputs of the model and measuring data of one real test firing demonstrates that the model is of many advantages, such as short computational time, better real time property and good precision. The model is very well fit for the applications of real time condition monitoring, fault diagnosis and control system design of liquid propellant rocket's propulsion system.

Numerical simulation of a liquid propellant rocket motor

Abstract: This work presents a numerical simulation of the flow field in a liquid propellant rocket engine chamber and exit nozzle using techniques to allow the results to be taken as starting points for designing those propulsive systems. This was done using a Finite Volume method simulating the different flow regimes which usually take place in those systems. As the flow field has regions ranging from the low subsonic to the supersonic regimes, the numerical code used, initially developed for compressible flows only, was modified to work proficiently in the whole velocity range. It is well known that codes have been developed in CFD, for either compressible or incompressible flows, the joint treatment of both together being complex even today, given the small number of references available in this area. Here an existing code for compressible flow was used and primitive variables, the pressure, the Cartesian components of the velocity and the temperature instead of the conserved variables were introduced in the Euler and Navier-Stokes equations. This was done to permit the treatment at any Mach number. Unstructured meshes with adaptive refinements were employed here. The convective terms were treated with upwind first and second order methods. The numerical stability was kept with artificial dissipation and in the spatial coverage one used a five stage Runge-Kutta scheme for the Fluid Mechanics and the VODE (Value of Ordinary Differential Equations) scheme along with the Chemkin II in the chemical reacting solution. During the development of this code simulating the flow in a rocket engine, comparison tests were made with several different types of internal and external flows, at different velocities, seeking to establish the confidence level of the techniques being used. These comparisons were done with existing theoretical results and with other codes already validated and well accepted by the CFD community.

Liquid Rocket Combustion Instability Analysis Methodology by CFD

Abstract: The high frequency combustion instability in liquid rocket engine LRE is analytically studied by using CFD methods.The comprehensive model and governing equations for numerical simulation of high frequency combustion instability are presented.The PISO and MacCormack algorithms are compared in this study.Both methods can successfully handle combustion instability analysis.The ACLRECI program is developed.The combustion instability phenomena of YF 860 LH2/LOX rocket and NAL LOX/Methane rocket are numerically simulated with the ACLRECI program.The combustion stability maps of these rockets are obtained.Numerical simulation results are in good agreement with test data.

Method of operation of liquid-propellant rocket engine with turbopump delivery of oxygen-methane propellant

Abstract: FIELD: rocket engines. SUBSTANCE: invention relates to operation of liquid-propellant rocket engine with turbopump delivery of oxygen-methane propellant with part of expendable methane propellant used as cooling agent for direct-flow cooling of chamber. Said cooling agent, after using, is mixed with remaining mass of propellant, pressure of mixture is raised and mixture is fed into gas generator for combustion with part of oxygen oxidizer for getting recovery gas used on turbine and then waste gas is afterburned in chamber with remaining mass of oxidizer. Cooling agent, before mixing with remaining mass of propellant, is cooled by heat exchange with oxygen oxidizer, and part of propellant is used to create curtain cooling of chamber by delivering cooling agent to inner wall of chamber through provided belt of holes. EFFECT: increased specific thrust impulse of rocket engine. 2 cl, 2 dwg