固液混合火箭发动机工作过程研究
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摘要
固液混合火箭发动机具有安全性高、可靠性高、研制和生产费用低、可实现推力调节和多次启动、推进剂及其燃烧产物毒性低、燃烧产物对环境污染小等众多优点,使得它们在商用卫星发射以及载人航天领域,尤其是在先进轨姿控推进系统和快速机动发射运载器推进系统方面具有很强的竞争力和诱人的应用前景。然而,由于固液混合火箭发动机内部工作过程十分复杂,通过国内外众多科学家与工程师们多年的不懈探索,目前已取得了一些研究成果,但这些成果尚不足以帮助人们清晰理解固液混合火箭发动机的工作过程,有必要就此问题继续开展深入细致的研究。
本文采用理论分析、数值模拟和试验等手段,研究了推进剂组合、发动机几何构型以及工作状态对固液混合火箭发动机性能的影响规律,探索了提高发动机性能的途径,并为固液混合火箭发动机设计提供了理论依据和指导。
采用热力计算方法和固液混合火箭发动机理论性能计算软件,研究了推进剂组合、液固比、燃烧室压力等参数对发动机性能的影响规律。推导了固液混合火箭发动机一维内弹道方程,并分析了在发动机工作过程中燃面后退速率、固体燃料气化产物流量、液固比、压强、温度和比冲等参数随时间变化规律。
针对固液混合火箭发动机中固体燃料中氧化剂含量少、燃面后退速率低以及固体燃料中固体含量少、工艺性能和力学性能差的特点,系统地进行了推进剂配方试验研究。采用提高粘合剂网络的交联密度和填料的补强技术,提高了HTPB基固体燃料的强度。研究各类添加剂对固体燃料燃面后退速率的影响规律,提高了HTPB基固体燃料的燃面后退速率。研制出力学性能优良(抗拉强度≥1.8MPa,延伸率≥50%)、燃面后退速率较高(≥0.3mm/s)、不同金属含量的系列HTPB基固体燃料。
采用基于流-固耦合的方法,在充分考虑固液混合火箭发动机工作过程中诸多复杂物理过程的基础上,建立了一个可适用于发动机不同工作状态下燃面后退速率预示的计算模型,并开发了一个可适用于固液混合火箭发动机燃烧室内两相湍流燃烧流动数值模拟的计算程序。对发动机燃烧室内两相湍流燃烧流动过程进行了数值模拟分析,得到了燃烧室内的流场结构,系统研究了固液混合火箭发动机推进剂组合、发动机几何构型以及工作状态对发动机性能的影响规律,为合理设计发动机结构指明了方向。
设计了直径80mm的GOX、90%H_2O_2/HTPB固液混合火箭发动机燃烧实验器,开展了发动机多次点火技术以及发动机燃烧性能试验研究。设计了采用煤油作为燃料的点火器,能够可靠、可控地实现GOX/HTPB固液混合火箭发动机燃烧实验器的启动、关机和再点火。通过终止发动机燃烧试验和系统辩识的方法,研究了GOX/HTPB固液混合火箭发动机燃面后退速率的影响规律。采用烃类燃料在催化分解的90%H_2O_2中能燃烧的点火器,进行了90%H_2O_2/HTPB固液混合火箭发动机试验初步研究。
Hybrid rocket motors keep the features of low cost, high reliability, and no hazards. In recent years, there has been a resurgence of interest in the development of hybrid rocket engines for advanced launch vehicle applications. But in hybrid rocket motors, the combustion efficiency and regression rate are very slow. Although the aforementioned work has provided considerable information to the understanding of flow and hybrid combustion process in hybrid rocket motors, many fundamental issues regarding the detailed mixing and combustion process, especially the performance of regression rate, still need to be addressed.
Of particular importance in the design of hybrid rocket motors is the fuel surface regression rate and the manner in which it varies with operation conditions. In view of this, a research program involving both experimental and numerical approaches is initiated. This thesis is aimed at providing the complete details of the hybrid combustion processes and flowfields in hybrid rocket motors and seeking ways to improve the combustion efficiency in order to provide theoretical basis and direction for the design of hybrid rocket motors. The results attained are as follows.
Based on the thermodynamic calculation and theoretical performance analysis, effects of fuel, oxidizer/fuel ratio, combustor pressure on the performance of hybrid rocket motor and temperature of combustor are analyzed. A ballistics model has been developed for the purpose of investigating the influence of fuel grain design on overall performance of hybrid rocket motors. The model, based on steady, one-dimensional compressible flow, includes the capability to handle arbitray wagon-wheel fuel section designs. Results of the ballistics calculations are presented for liquid oxygen as fuel.
Aimed at hybrid propellant having low strength and poor combustion performance, development of HTPB propellant is presented. In order to determine the effects of solid-fuel additives on the combustion behavior of HTPB and GOX, many different types of powders are added to the solid fuel. Kinds of intensifiers with active groups are added to enhance curing network of HTPB binder system. The propellant is endowed with excellent performances of different metal content, tensile strength up to 1.8MPa, regression rate>0.3mm/s.
Based on the fluid-solid coupling technique and some comprehensive physical processes during working of hybrid rocket motor, a numerical model is developed to predict the regression rate for the solid fuel surface of hybrid rocket motor under different working conditions. The muti-dimensional Favre-averaged compressible turbulent N-S equations are used as the governing equations of the reacting flow, the improved k-εtwo-equation turbulence model is used to simulate the turbulent flow, and Eddy-Dissipation Model is used to simulate the gas combustion. Accuracy of the calculated model is verified by comparison of calculation results with experiment data. The simulation results of model motor show that combustion, flow and regression rate of solid fuel surface of hybrid rocket motor are of inhomogeneity. Information gained from CFD modeling provides future engine designers valuable information regarding the design of hybrid rocket motors combustion chambers.
The paper also makes a whole design of experimental hybrid rocket motor system, and discusses the result of motor firing. Technical project and experiment results of multiple start and shutoff of GOX/HTPB hybrid rocket motors are presented. The performance data of hybrid rocket motors are measured. Various kinds of factors, which affect hybrid rocket motors performance, are preliminarily analyzed. The experiments show that multiple-start and shutoff can be realized and its number of times and time intervals can be regulated at ramdom. An innovative igniter based on catalytic decomposition of H_2O_2 is designed and manufcatured. A H_2O_2/HTPB hybrid rocket motor using this igniter was tested. The experiment results demonstrated that this new igniter can successful start the H_2O_2/HTPB hybrid rocket motors.
本文采用理论分析、数值模拟和试验等手段,研究了推进剂组合、发动机几何构型以及工作状态对固液混合火箭发动机性能的影响规律,探索了提高发动机性能的途径,并为固液混合火箭发动机设计提供了理论依据和指导。
采用热力计算方法和固液混合火箭发动机理论性能计算软件,研究了推进剂组合、液固比、燃烧室压力等参数对发动机性能的影响规律。推导了固液混合火箭发动机一维内弹道方程,并分析了在发动机工作过程中燃面后退速率、固体燃料气化产物流量、液固比、压强、温度和比冲等参数随时间变化规律。
针对固液混合火箭发动机中固体燃料中氧化剂含量少、燃面后退速率低以及固体燃料中固体含量少、工艺性能和力学性能差的特点,系统地进行了推进剂配方试验研究。采用提高粘合剂网络的交联密度和填料的补强技术,提高了HTPB基固体燃料的强度。研究各类添加剂对固体燃料燃面后退速率的影响规律,提高了HTPB基固体燃料的燃面后退速率。研制出力学性能优良(抗拉强度≥1.8MPa,延伸率≥50%)、燃面后退速率较高(≥0.3mm/s)、不同金属含量的系列HTPB基固体燃料。
采用基于流-固耦合的方法,在充分考虑固液混合火箭发动机工作过程中诸多复杂物理过程的基础上,建立了一个可适用于发动机不同工作状态下燃面后退速率预示的计算模型,并开发了一个可适用于固液混合火箭发动机燃烧室内两相湍流燃烧流动数值模拟的计算程序。对发动机燃烧室内两相湍流燃烧流动过程进行了数值模拟分析,得到了燃烧室内的流场结构,系统研究了固液混合火箭发动机推进剂组合、发动机几何构型以及工作状态对发动机性能的影响规律,为合理设计发动机结构指明了方向。
设计了直径80mm的GOX、90%H_2O_2/HTPB固液混合火箭发动机燃烧实验器,开展了发动机多次点火技术以及发动机燃烧性能试验研究。设计了采用煤油作为燃料的点火器,能够可靠、可控地实现GOX/HTPB固液混合火箭发动机燃烧实验器的启动、关机和再点火。通过终止发动机燃烧试验和系统辩识的方法,研究了GOX/HTPB固液混合火箭发动机燃面后退速率的影响规律。采用烃类燃料在催化分解的90%H_2O_2中能燃烧的点火器,进行了90%H_2O_2/HTPB固液混合火箭发动机试验初步研究。
Hybrid rocket motors keep the features of low cost, high reliability, and no hazards. In recent years, there has been a resurgence of interest in the development of hybrid rocket engines for advanced launch vehicle applications. But in hybrid rocket motors, the combustion efficiency and regression rate are very slow. Although the aforementioned work has provided considerable information to the understanding of flow and hybrid combustion process in hybrid rocket motors, many fundamental issues regarding the detailed mixing and combustion process, especially the performance of regression rate, still need to be addressed.
Of particular importance in the design of hybrid rocket motors is the fuel surface regression rate and the manner in which it varies with operation conditions. In view of this, a research program involving both experimental and numerical approaches is initiated. This thesis is aimed at providing the complete details of the hybrid combustion processes and flowfields in hybrid rocket motors and seeking ways to improve the combustion efficiency in order to provide theoretical basis and direction for the design of hybrid rocket motors. The results attained are as follows.
Based on the thermodynamic calculation and theoretical performance analysis, effects of fuel, oxidizer/fuel ratio, combustor pressure on the performance of hybrid rocket motor and temperature of combustor are analyzed. A ballistics model has been developed for the purpose of investigating the influence of fuel grain design on overall performance of hybrid rocket motors. The model, based on steady, one-dimensional compressible flow, includes the capability to handle arbitray wagon-wheel fuel section designs. Results of the ballistics calculations are presented for liquid oxygen as fuel.
Aimed at hybrid propellant having low strength and poor combustion performance, development of HTPB propellant is presented. In order to determine the effects of solid-fuel additives on the combustion behavior of HTPB and GOX, many different types of powders are added to the solid fuel. Kinds of intensifiers with active groups are added to enhance curing network of HTPB binder system. The propellant is endowed with excellent performances of different metal content, tensile strength up to 1.8MPa, regression rate>0.3mm/s.
Based on the fluid-solid coupling technique and some comprehensive physical processes during working of hybrid rocket motor, a numerical model is developed to predict the regression rate for the solid fuel surface of hybrid rocket motor under different working conditions. The muti-dimensional Favre-averaged compressible turbulent N-S equations are used as the governing equations of the reacting flow, the improved k-εtwo-equation turbulence model is used to simulate the turbulent flow, and Eddy-Dissipation Model is used to simulate the gas combustion. Accuracy of the calculated model is verified by comparison of calculation results with experiment data. The simulation results of model motor show that combustion, flow and regression rate of solid fuel surface of hybrid rocket motor are of inhomogeneity. Information gained from CFD modeling provides future engine designers valuable information regarding the design of hybrid rocket motors combustion chambers.
The paper also makes a whole design of experimental hybrid rocket motor system, and discusses the result of motor firing. Technical project and experiment results of multiple start and shutoff of GOX/HTPB hybrid rocket motors are presented. The performance data of hybrid rocket motors are measured. Various kinds of factors, which affect hybrid rocket motors performance, are preliminarily analyzed. The experiments show that multiple-start and shutoff can be realized and its number of times and time intervals can be regulated at ramdom. An innovative igniter based on catalytic decomposition of H_2O_2 is designed and manufcatured. A H_2O_2/HTPB hybrid rocket motor using this igniter was tested. The experiment results demonstrated that this new igniter can successful start the H_2O_2/HTPB hybrid rocket motors.
引文
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