Optimization and Sizing for Propulsion System of Liquid Rocket Using Genetic Algorithm

Flight vehicle conceptual design appears to be a promising area for application of the Genetic Algorithm （GA） as an approach to help to automate part of the design process. This computational research effort strives to develop a propulsion system design strategy for liquid rocket to optimize take-off mass, satisfying the mission range under the constraint of axial overload. The method by which this process is accomplished by using GA as optimizer is outlined in this paper. Convergence of GA is improved by introducing initial population based on Design of Experiments Technique.

An Overview of Bearing Candidates for the Next Generation of Reusable Liquid Rocket Turbopumps

There is a consensus in the aerospace field that the development of reusable liquid rockets can effectively reduce the launch expense.The pursuit of a long service life and reutilization highly depends on the bearing components.However,the rolling element bearings(REBs)used in the existing rocket turbopumps present obvious and increasing limitations due to their mechanical contacting mode.For REBs,high rotational speed and long service life are two performance indexes that mutually restrict each other.To go beyond the DN value(the product of the bearing bore and rotational speed)limit of REBs,the major space powers have conducted substantial explorations on the use of new types of bearings to replace the REB.This review discusses,first,the crucial role of bearings in rocket turbopumps and the related structural improvements of REBs.Then,with the prospect of application to the next generation of reusable liquid rocket turbopumps,the bearing candidates investigated by major space powers are summarized comprehensively.These promising alternatives to REBs include fluid-film,foil,and magnetic bearings,together with the novel superconducting compound bearings recently proposed by our team.Our more than ten years of relevant research on fluid-film and magnetic bearings are also introduced.This review is meaningful for the development of long-life and highly reliable bearings to be used in future reusable rocket turbopumps.

A Comparative Study of Genetic Algorithm Parameters for the Inverse Problem-based Fault Diagnosis of Liquid Rocket Propulsion Systems

Fault diagnosis of liquid rocket propulsion systems （LRPSs） is a very important issue in space launch activities particularly when manned space missions are accompanied, since the safety and reliability can be significantly enhanced by exploiting an efficient fault diagnosis system. Currently, inverse problem-based diagnosis has attracted a great deal of research attention in fault diagnosis domain. This methodology provides a new strategy to model-based fault diagnosis for monitoring the health of propulsion systems. To solve the inverse problems arising from the fault diagnosis of LRPSs, GAs have been adopted in recent years as the first and effective choice of available numerical optimization tools. However, the GA has many control parameters to be chosen in advance and there still lack sound theoretical tools to analyze the effects of these parameters on diagnostic performance analytically. In this paper a comparative study of the influence of GA parameters on diagnostic results is conducted by performing a series of numerical experiments. The objective of this study is to investigate the contribution of individual algorithm parameter to final diagnostic result and provide reasonable estimates for choosing GA parameters in the inverse problem-based fault diagnosis of LRPSs. Some constructive remarks are made in conclusion and will be helpful for the implementation of GA to the fault diagnosis practice of LRPSs in the future.

First Systematic Testing Platform for Pressurization Feed System Developed for Liquid Rocket Propellant in China

Beijing Aerospace System Engineering Institute of China Academy of Launch Vehicle Technology (CALT) declared recently that theinstitute has set up a laboratory whichwould operate a newly

Safety Analysis of Liquid Rocket Engine Using Bayesian Networks

Safety analysis for liquid rocket engine has a great meaning for shortening development cycle, saving development expenditure and reducing development risk. The relationship between the structure and component of liquid rocket engine is much more complex, furthermore test data are absent in development phase. Thereby, the uncertainties exist in safety analysis for liquid rocket engine. A safety analysis model integrated with FMEA(failure mode and effect analysis) based on Bayesian networks (BN) is brought forward for liquid rocket engine, which can combine qualitative analysis with quantitative decision. The method has the advantages of fusing multi-information, saving sample amount and having high veracity. An example shows that the method is efficient.

Analysis of combustion instability via constant volume combustion in a LOX/RP-1 bipropellant liquid rocket engine

Turbulent two-phase reacting flow in the chamber of LOX/RP-1 bipropellant liquid rocket engine is numerically investigated in this paper. The predicted pressure and mean axial velocity are qualitatively consistent with the experimental measurements. The self-excited pressure oscillations are obtained without any disturbance introduced through the initial and boundary conditions. It is found that amount of abrupt pressure peaks appear frequently and stochastically in the head regions of the chamber, which are the important sources to drive and strengthen combustion instability. Such abrupt pressures are induced by local constant volume combustion, because local combustible gas mixtures with high temperature are formed and burnt out suddenly due to some fuel droplets reaching their critical state in a rich oxygen surrounding. A third Damkhler number is defined as the ratio of the characteristic time of a chemical reaction to the characteristic time of a pressure wave expansion to measure the relative intensity of acoustic propagation and combustion process in thrusters. The analysis of the third Damkhler number distributions in the whole thrust chamber shows that local constant volume combustion happens in the head regions, while constant pressure combustion presents in the downstream regions. It is found that the combustion instability occurs in the head regions within about 30 mm from the thruster head.

Coupled Lagrangian impingement spray model for doublet impinging injectors under liquid rocket engine operating conditions

To predict the effect of the liquid rocket engine combustion chamber conditions on the impingement spray, the conventional uncoupled spray model for impinging injectors is extended by considering the coupling of the jet impingement process and the ambient gas field. The new coupled model consists of the plain-orifice sub-model, the jet-jet impingement sub-model and the droplet collision sub-model. The parameters of the child droplet are determined with the jet-jet impingement sub-model using correlations about the liquid jet parameters and the chamber conditions.The overall model is benchmarked under various impingement angles, jet momentum and offcenter ratios. Agreement with the published experimental data validates the ability of the model to predict the key spray characteristics, such as the mass flux and mixture ratio distributions in quiescent air. Besides, impinging sprays under changing ambient pressure and non-uniform gas flow are investigated to explore the effect of liquid rocket engine chamber conditions. First, a transient impingement spray during engine start-up phase is simulated with prescribed pressure profile. The minimum average droplet diameter is achieved when the orifices work in cavitation state, and is about 30% smaller than the steady single phase state. Second, the effect of non-uniform gas flow produces off-center impingement and the rotated spray fan by 38°. The proposed model suggests more reasonable impingement spray characteristics than the uncoupled one and can be used as the first step in the complex simulation of coupling impingement spray and combustion in liquid rocket engines.

Study of pressure surge during priming phase of start transient in an initially unprimed pump-fed liquid rocket engine

In this paper, transient phenomenon during start up process of a pump fed liquidrocket engine is investigated through numerical simulation. The engine studied in this workis designed such that engine systems are not wetted with propellant until the engine is com-manded to start. This is achieved by positioning the valves for propellant admission at the inter-face of test stand/flight stage and the engine. To evaluate engine performance during starttransient for such systems, unsteady flow simulation was conducted using Method of Charac-teristics and equations for priming. The same has been reported in this work. The results indi-cated a brief period of abrupt pressure rise at pump upstream after opening of the propellantadmission valves, during the process of priming of engine systems at valve downstream.The peak pressure obtained was significantly higher than the propellant tank pressure as wellas the steady state pump suction pressure. The transitory pressure rise was found to occurdue to flow resistance at impeller inlet caused by formation of a forced vortex for orientingthe flow through impeller blades during off design transient regime. The maximum pressureat pump upstream, as computed from start transient simulation, was used as a design inputfor pump inlet feed lines. The engine was realized and subsequently qualified in a ground test facility. Hot test data obtained for pressure and flow rate during transient regime were found tobe in good agreement with the simulation results.

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

During the experiment of gas/liquid coaxial swirl injector conducted with air and water under atmosphere environment, it is observed that the injector may selfoscillate. The self oscillation periodically occurs and vanishes with the increasing velocity of the gas flow.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 exPeriment 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 related to self oscillation in coaxial injectors

Verification on Spray Simulation of a Pintle Injector for Liquid Rocket Engine

The pintle injector used for a liquid rocket engine is a newly re-attracted injection system famous for its wide throttle ability with high efficiency. The pintle injector has many variations with complex inner structures due to its moving parts. In order to study the rotating flow near the injector tip, which was observed from the cold flow experiment using water and air, a numerical simulation was adopted and a verification of the numerical model was later conducted. For the verification process, three types of experimental data including velocity distributions of gas flows, spray angles and liquid distribution were all compared using simulated results. The numerical simulation was performed using a commercial simulation program with the Eulerian multiphase model and axisymmetric two dimensional grids. The maximum and minimum velocities of gas were within the acceptable range of agreement, however, the spray angles experienced up to 25% error when the momentum ratios were increased. The spray density distributions were quantitatively measured and had good agreement. As a result of this study, it was concluded that the simulation method was properly constructed to study specific flow characteristics of the pintle injector despite having the limitations of two dimensional and coarse grids.

Genetic Algorithm to Optimize the Design of Main Combustor and Gas Generator in Liquid Rocket Engines

A genetic algorithm was used to develop optimal design methods for the regenerative cooled combustor and fuel-rich gas generator of a liquid rocket engine. For the combustor design, a chemical equilibrium analysis was applied, and the profile was calculated using Rao's method. One-dimensional heat transfer was assumed along the profile, and cooling channels were designed. For the gas-generator design, non-equilibrium properties were derived from a counterflow analysis, and a vaporization model for the fuel droplet was adopted to calculate residence time. Finally, a genetic algorithm was adopted to optimize the designs. The combustor and gas generator were optimally designed for 30-tonf, 75-tonf, and 150-tonf engines. The optimized combustors demonstrated superior design characteristics when compared with previous non-optimized results. Wall temperatures at the nozzle throat were optimized to satisfy the requirement of 800 K, and specific impulses were maximized. In addition, the target turbine power and a burned-gas temperature of 1000 K were obtained from the optimized gas-generator design.

Development of Preliminary Design Program for Combustor of Regenerative Cooled Liquid Rocket Engine

An integrated program was established to design a combustor for a liquid rocket engine and to analyze regenerative cooling results on a preliminary design level.Properties of burnt gas from a kerosene-LOx mixture in the combustor and rocket performance were calculated from CEA which is the code for the calculation of chemical equilibrium.The heat transfer of regenerative cooling was analyzed by using SUPERTRAPP code for coolant properties and by one-dimensional correlations of the heat transfer coefficient from the combustor liner to the coolant.Profiles of the combustors of F-1 and RS-27A engines were designed from similar input data and the present results were compared to actual data for validation.Finally,the combustors of 30 tonf class,75 tonf class and 150 tonf class were designed from the required thrust,combustion chamber,exit pressure and mixture ratio of propellants.The wall temperature,heat flux and pressure drop were calculated for heat transfer analysis of regenerative cooling using the profiles.

Conceptual Design for a Kerosene Fuel-rich Gas-generator of a Turbopump-fed Liquid Rocket Engine

A design method for a kerosene fuel-rich gas-generator of a liquid rocket engine using turbopumps to supply propellant was performed at a conceptual level. The gas-generator creates hot gases, enabling the turbine to operate the turbopumps. A chemical non-equilibrium analysis and a droplet vaporization model were used for the estimation of the burnt gas properties and characteristic chamber length. A premixed counter-flow flame analysis was performed for the prediction of the burnt gas properties, namely the temperature, the specific heat ratio and heat capacity, and the chemical reaction time. To predict the vaporization time, the Spalding model, using a single droplet in convective condition, was used. The minimum residence time in the chamber and the characteristic length were calculated by adding the reaction time and the vaporization time. Using the characteristic length, the design methods for the fuel-rich gas-generator were established. Finally, a parametric study was achieved for the effects of the O/F ratio, mass flow rate, chamber pressure, initial droplet temperature, initial droplet diameter and initial droplet velocity.

Thermal state calculation of chamber in small thrust liquid rocket engine for steady state pulsed mode

This paper presents a method of thermal state calculation of combustion chamber in small thrust liquid rocket engine. The goal is to predict the thermal state of chamber wall by using basic parameters of engine: thrust level, propellants, chamber pressure, injection pattern, film cooling parameters, material of wall and their coating, etc. The difficulties in modeling the startup and shutdown processes of thrusters lie in the fact that there are the conjugated physical processes occurring at various parameters for non-design conditions. A mathematical model to predict the thermal state of the combustion chamber for different engine operation modes is developed. To simulate the startup and shutdown processes, a quasi-steady approach is applied by replacing the transient process with time-variant operating parameters of steady-state processes. The mathematical model is based on several principles and data commonly used for heat transfer modeling: geometry of flow part, gas dynamics of flow, thermodynamics of propellants and combustion spices, convective and radiation heat flows, conjugated heat transfer between hot gas and wall, and transient approach for calculation of thermal state of construction. Calculations of the thermal state of the combustion chamber in single-turn-on mode show good convergence with the experimental results. The results of pulsed modes indicate a large temperature gradient on the internal wall surface of the chamber between pulses and the thermal state of the wall strongly depends on the pulse duration and the interval.

Regenerative Cooling for Liquid Rocket Engines

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Nonlinear theory of combustion stability in liquid rocket engine based on chemistry dynamics

Detailed models of combustion instability based on chemistry dynamics are developed. The results show that large activation energy goes against the combustion stability. The heat transfer coefficient between the wall and the combust gas is an important bifurcation parameter for the combustion instability. The acoustics modes of the chamber are in competition and cooperation with each other for limited vibration energy. Thermodynamics criterion of combustion stability can be deduced from the nonlinear thermodynamics. Correlations of the theoretical results and historical experiments indicate that chemical kinetics play a critical role in the combustion instability.

Numerical simulation of combustion stability of liquid rocket engine based on chemistry dynamics

Combustion instability of O2/kerosene, O2/kerosene/hydrogen, and O2/kerosene/hydro- gen spray flame is numerically studied. The numerical results of combustion self-oscillation are consistent with the historical experiments. Hydrogen is helpful to stabilizing oxygen/hydrocarbon combustion. High gas injecting velocity of the coaxial injector would increase the combustion stability. Contrary to the former expectation, the most sensitive region for combustion instability is not where the heat releases most intensely but is the low-temperature premixed region near the injectors. According to the simulation, the technology steps, such as adding catalyzer to decrease the reaction activity energy, or improving the injector design to reduce the premixed low temperature region, would improve the combustion stability.

重复使用运载火箭液体动力技术发展

重复使用是未来运载火箭更新换代的技术发展趋势,是降低航天发射成本、实现规模化航天发射的有效途径。重点概述了国内外垂直起降重复使用运载火箭动力技术的发展现状,分析了垂直起降重复使用运载火箭发射和回收全任务剖面,总结了垂直起降重复使用运载火箭动力技术的特点,包括宽范围入口压力多次启动技术、大范围快速高精度推力调节技术、故障诊断及健康管理技术、状态检测与维修维护技术等。

基于Hessian局部线性嵌入和MLP-Mixer的液体火箭发动机涡轮泵轻量化故障诊断框架

作为液体火箭发动机推进剂输送系统的关键部件,涡轮泵的运行状态直接影响着整个运载系统的性能,然而,现有的故障诊断方法往往面临特性参数选择片面及计算复杂度高等问题。针对上述局限,提出了面向涡轮泵的轻量化故障诊断框架。所提方法利用Hessian局部线性嵌入算法对信号时域、频域及时频特征进行降维,并引入一种轻量化的深度学习模型MLP-Mixer作为分类器,进而实现不同故障状态的辨识。采用某型号涡轮泵试车数据验证了所提方法的有效性,结果表明,该方法能够在保障诊断精度的同时有效降低计算复杂度,提高诊断效率。