Experimental and Theoretical Research Review of Hybrid Rocket Motor Techniques and Applications

A hybrid rocket motor combines components from both solid fuel and liquid fuel rocket motors. The fuel itself is a solid grain, (often paraffin or hydroxyl-terminated polybutadiene, known as HTPB) while the oxidizing agent is liquid (often hydrogen peroxide or liquid oxygen). These components are combined in the fuel chamber which doubles as the combustion chamber for the hybrid motor. This review looks at the advances in techniques that have taken place in the development of these motors since 1995. Methods of testing the thrust from rocket motors and of measuring the rocket plume spectroscopically for combustion reaction products have been developed. These assessments allow researchers to more completely understand the effects of additives and physical changes in design, in terms of regression rates and thrust developed. Hybrid rocket motors have been used or tested in many areas of rocketry, including tactical rockets and large launch vehicles. Several additives have shown significant improvements in regression rates and thrust, including Guanidinium azotetrazolate (GAT), and various Aluminum alloys. The most recent discoveries have come from research into nano-particle additives. The nano-particles have been shown to provide enhancements to many parameters of hybrid rocket function, and research into specific areas continues in the sub-field of nano-additives for fuel grains.

Numerical analysis on combustion characteristics of hybrid rocket motor with star-tube segmented grain

A method of star-tube combined segmented grain is proposed to improve the combustion performance of hybrid rocket motor.The star-tube combined segmented grain consists of a single-port star part and a single-port tube part.A mid-chamber forms between the fore-grain and the aft-grain for better mixing effect.The single-port feature gives hybrid rocket motor several advantages,such as simple structure,high reliability,and variable combinations.This paper is mainly aimed at studying the combustion characteristics of hybrid rocket motor with star-tube segmented grain through three-dimensional steady simulations.Combustion performance of the motors with different segmented grain combinations,including fore-tube/aft-tube,fore-tube/aftstar,fore-star/aft-star and fore-star/aft-tube,is contrastively analyzed.The motor in this paper adopts polyethylene and 90%hydrogen peroxide as the propellants.Simulations reveal that segmented grain with different-type grain combinations could greatly change the flow field in the second half of the combustion chamber.Transformation of the flow field is beneficial to the mixing between the fuel and the oxidizer,and it could increase the fuel regression rate and the combustion efficiency.The turbulence effect of tube aft-grain is better than that of star aft-grain.Among the four segmented grain combinations,the combination of star fore-grain and tube aft-grain is the preferred method with optimal overall performance.This grain configuration could increase the regression rate of tube aft-grain to surpass that of star aft-grain in other combinations.Besides,hybrid rocket motor with this grain configuration achieves the highest combustion efficiency.

特征结构对复合式石蜡基药柱燃烧特性的影响研究

凹槽状特征结构是复合式石蜡基药柱相对于传统石蜡基药柱燃烧性能有效改善的重要因素,该特征结构因螺旋基底与石蜡基燃料退移速率不同而产生。研究了预设0,2mm尺度的特征结构(初始凹槽深度为0,2mm)对于燃烧室压强、退移速率、特征速度的影响规律。点火实验以气氧作为氧化剂,平均流量分别为11.6,18.5,23.1g/s,同时选择传统圆孔石蜡基药柱作为对比分析。实验结果表明:在相同氧化剂质量流量下,预设2mm尺度的复合式药柱点火到室压稳定所需的时间最短,退移速率和特征速度也最高。采用计算流体力学(CFD)分析了不同尺度的特征结构对燃气流动特性的影响规律,冷流结果表明尺度越大的特征结构引导旋流强度更高,且影响范围更大。

Numerical analysis of fuel regression rate distribution characteristics in hybrid rocket motors with different fuel types

This paper presents three-dimensional numerical simulations of the hybrid rocket motor with hydrogen peroxide (HP) and hy-droxyl terminated polybutadiene (HTPB) propellant combination and investigates the fuel regression rate distribution charac-teristics of different fuel types. The numerical models are established to couple the Navier-Stokes equations with turbulence,chemical reactions, solid fuel pyrolysis and solid-gas interfacial boundary conditions. Simulation results including the temper-ature contours and fuel regression rate distributions are presented for the tube, star and wagon wheel grains. The results demonstrate that the changing trends of the regression rate along the axis are similar for all kinds of fuel types, which decrease sharply near the leading edges of the fuels and then gradually increase with increasing axial locations. The regression rates of the star and wagon wheel grains show apparent three-dimensional characteristics, and they are higher in the regions of fuel surfaces near the central core oxidizer flow. The average regression rates increase as the oxidizer mass fluxes rise for all of the fuel types. However, under same oxidizer mass flux, the average regression rates of the star and wagon wheel grains are much larger than that of the tube grain due to their lower hydraulic diameters.

Uncertainty analysis and design optimization of hybrid rocket motor powered vehicle for suborbital flight

Abstract In this paper, we propose an uncertainty analysis and design optimization method and its applications on a hybrid rocket motor （HRM） powered vehicle. The multidisciplinary design model of the rocket system is established and the design uncertainties are quantified. The sensitivity anal- ysis of the uncertainties shows that the uncertainty generated from the error of fuel regression rate model has the most significant effect on the system performances. Then the differences between deterministic design optimization （DDO） and uncertainty-based design optimization （UDO） are discussed. Two newly formed uncertainty analysis methods, including the Kriging-based Monte Carlo simulation （KMCS） and Kriging-based Taylor series approximation （KTSA）, are carried out using a global approximation Kriging modeling method. Based on the system design model and the results of design uncertainty analysis, the design optimization of an HRM powered vehicle for suborbital flight is implemented using three design optimization methods： DDO, KMCS and KTSA. The comparisons indicate that the two UDO methods can enhance the design reliability and robustness. The researches and methods proposed in this paper can provide a better way for the general design of HRM powered vehicles.

Multi-disciplinary design optimization with fuzzy uncertainties and its application in hybrid rocket motor powered launch vehicle

In this paper,an Uncertainty-based Multi-disciplinary Design Optimization (UMDO)method combining with fuzzy theory and Multi-Discipline Feasible (MDF) method is developed for the conceptual design of a Hybrid Rocket Motor (HRM) powered Launch Vehicle (LV).In the method proposed,membership functions are used to represent the uncertain factors,the fuzzy statistical experiment is introduced to analyze the propagation of uncertainties,and means,standard deviations and credibility measures are used to delineate uncertain responses.A geometric programming problem is solved to verify the feasibility of the Fuzzy-based Multi-Discipline Feasible(F-MDF) method.A multi-disciplinary analysis of a three-stage HRM powered LV involving the disciplines of propulsion,structure,aerodynamics and trajectory is implemented,and the mathematical models corresponding to the F-MDF method and the MDF method are established.A two-phase optimization method is proposed for multi-disciplinary design optimization of the LV,including the orbital capacity optimization phase based on the Ziolkowski formula,and the scheme trajectory verification phase based on the 3-degree-of-freedom point trajectory simulation.The correlation coefficients and the quadratic Response Surface Method (RSM) based on Latin Hypercube Sampling (LHS) are adopted for sensitive analysis of uncertain factors,and the Multi-Island Genetic Algorithm (MIGA) is adopted as the optimization algorithm.The results show that the F-MDF method is applicable in LV conceptual design,and the design with the F-MDF method is more reliable and robust than that with the MDF method.

Hybrid uncertainty-based design optimization and its application to hybrid rocket motors for manned lunar landing

Design reliability and robustness are getting increasingly important for the general design of aerospace systems with many inherently uncertain design parameters.This paper presents a hybrid uncertainty-based design optimization（UDO） method developed from probability theory and interval theory.Most of the uncertain design parameters which have sufficient information or experimental data are classified as random variables using probability theory,while the others are defined as interval variables with interval theory.Then a hybrid uncertainty analysis method based on Monte Carlo simulation and Taylor series interval analysis is developed to obtain the uncertainty propagation from the design parameters to system responses.Three design optimization strategies,including deterministic design optimization（DDO）,probabilistic UDO and hybrid UDO,are applied to the conceptual design of a hybrid rocket motor（HRM） used as the ascent propulsion system in Apollo lunar module.By comparison,the hybrid UDO is a feasible method and can be effectively applied to the general design of aerospace systems.

Three-dimensional numerical analysis on combustion performance and flow of hybrid rocket motor with multi-segmented grain

This paper presents the combustion characteristics in hybrid rocket motors with multisegmented grain through three-dimensional numerical simulations.Multi-segmented grain is composed of several thin grains with two or more ports.The numerical model consists of Navier-Stokes equations with turbulence,solid fuel pyrolysis,chemical reactions,a fluid–solid coupling model and a regression rate model.The simulations adopt 90%Hydrogen Peroxide(HP)and PolyEthylene(PE)as the propellant combination.The effects of the rotation,port number,fuel grain segment number and mid-chamber length on the flow field and combustion performances are analyzed.The results indicate that the multi-segmented grain configuration can strengthen the flow field,and the regression rate and combustion efficiency are enhanced.Take the cases with two grain segments and three ports for example,the regression rate is increased by 32.4%-45.1%and the combustion efficiency increases by 6%-8.6%in different rotation angles.

Design and optimization of variable thrust hybrid rocket motors for sounding rockets

Besides safety and low-cost,the start/shutdown/restarting and throttling ability are the other two significant advantages of hy-brid rocket motors(HRMs) compared with liquid and solid ones.In this study,a two-stage variable thrust and non-toxic 98%HP/HTPB hybrid rocket motor(VTHRM) is designed and applied in a sounding rocket,and the design parameters of the motor are analyzed and optimized.A computational program is developed to design the motor system structure,to predict the interior ballistics and the ballistic trajectory.A star grain and a wheel grain are compared.The design of experiment(DOE),variance analysis and the main effect analysis are employed to investigate the influence of the main design parameters on mo-tor performance.The multidiscipline feasible(MDF) approach is applied to establish the optimization procedure after analyz-ing the system design structure matrix.A modified differential evolution algorithm is employed to maximize the load mass.The results indicate that the wheel grain could obtain a larger load mass and a lower length to diameter ratio,and that throttling markedly meliorates the motor and rocket performance.The conclusions drawn from the analysis and optimization could pro-vide instructive guide and theoretical basis for engineering designs.

Numerical and experimental investigation on the effects of aft mixing chamber diaphragm in hybrid rocket motor

This paper focuses on the investigation of an aft mixing chamber diaphragm in a hybrid rocket motor.Both numerical and experimental researches are carried out to study its effects on the motor performances.The hybrid rocket motor with star fuel grain is utilized.The 90%hydrogen peroxide(HP)oxidizer and hydroxyl terminated polybutadiene(HTPB)based fuel are adopted as propellants.The diaphragm configuration settled in the aft mixing chamber includes four circular-holes with a uniform circumferential distribution.For both motors with and without the diaphragm,three-dimensional numerical simulations with gaseous combustions considered are carried out to study the flow field characteristics and motor performances.The comparison results show that the diaphragm can improve the mixing of the oxidizer and fuel.It has evident effect on increasing the motor efficiencies.Two firing experiments are conducted with full scale motors to investigate the effects of the diaphragm.The diaphragm used in the test is composed of a central steel framework and a closed thermal insulation layer.With the diaphragm employed,the combustion efficiency increases from 93.9%to 97.34%and the specific impulse efficiency increases from 80.77%to 87.28%,which verifies the positive effect of the diaphragm.Both numerical and experimental studies indicate that the scheme of the aft mixing chamber diaphragm proposed in the paper can improve the efficiencies of the hybrid rocket motor obviously.

Steady-state coupled analysis of flowfields and thermochemical erosion of C/C nozzles in hybrid rocket motors

A hybrid rocket can be used in various applications and is an attractive propulsion system.However,serious erosion of nozzles is common in motor firing operations,which could restrict the application of hybrid rocket motors.Usually,the serious erosion is attributed to the high concentration of oxidizing species in hybrid motors,while the details of flowfields in the motors are not paid special attention to.In this paper,first the thermochemical erosion of C/C nozzle is simulated coupled with the flowfields in a 98%H2O2/hydroxyl-terminated polybutadiene(HTPB)hybrid rocket motor.The simulation is made on a typical axisymmetric motor,including a pre-combustion chamber,an aft-combustion chamber and nozzle structures.Thermochemical reactions of H2O,CO2,OH,O and O2with C are taken into account.Second,the change of flowfields due to fuel regression during motor firing operations is considered.Nozzle erosion in different flowfields is evaluated.Third,the results of nozzle erosion in the coupled simulation are compared with those under uniform and chemical equilibrium flow and motor firing test results.The results of simulation and firing tests indicate that the thermochemical erosion of nozzles in hybrid motors should be calculated coupled with flowfields in the motor.In uniform and chemical equilibrium flowfields,the erosion rate is overestimated.The diffusion flame in hybrid motors protects the nozzle surface from the injected oxidizer and high temperature products in flowfields,leading to a relatively fuel-rich environment above the nozzle.The influence of OH and the geometry of motor should also be considered in the evaluation of nozzle erosion in hybrid motors.

Transient simulation of regression rate on thrust regulation process in hybrid rocket motor

The main goal of this paper is to study the characteristics of regression rate of solid grain during thrust regulation process. For this purpose, an unsteady numerical model of regression rate is established. Gas–solid coupling is considered between the solid grain surface and combustion gas.Dynamic mesh is used to simulate the regression process of the solid fuel surface. Based on this model, numerical simulations on a H2O2/HTPB（hydroxyl-terminated polybutadiene） hybrid motor have been performed in the flow control process. The simulation results show that under the step change of the oxidizer mass flow rate condition, the regression rate cannot reach a stable value instantly because the flow field requires a short time period to adjust. The regression rate increases with the linear gain of oxidizer mass flow rate, and has a higher slope than the relative inlet function of oxidizer flow rate. A shorter regulation time can cause a higher regression rate during regulation process. The results also show that transient calculation can better simulate the instantaneous regression rate in the operation process.

Uncertainty analysis and probabilistic design optimization of hybrid rocket motors for manned lunar landing

To obtain a conceptual design for a hybrid rocket motor(HRM)to be used as the Ascent Propulsion System in the Apollo lunar module,the deterministic design optimization(DDO)method is applied to the HRM design.Based on the results of an uncertainty analysis of HRMs,an uncertainty-based design optimization(UDO)method is also adopted to improve the design reliability.The HRM design process,which is a multidisciplinary system,is analyzed,and a mathematical model for the system design is established to compute the motor performance from the input parameters,including the input variables and model parameters.The input parameter uncertainties are quantified,and a sensitivity analysis of the model parameter uncertainties is conducted to identify the most important model parameter uncertainties for HRMs.The DDO and probabilistic UDO methods are applied to conceptual designs for an HRM to be used as a substitute for the liquid rocket motor(LRM)of the Ascent Propulsion System.The conceptual design results show that HRMs have several advantages as an alternative to the LRM of the Ascent Propulsion System,including nontoxic propellant combination,small motor volume,and comparable functions,such as restarting and throating.Comparisons of the DDO and UDO results indicate that the UDO method achieves more robust and reliable optimal designs than the DDO method.The probabilistic UDO method can be used to develop better conceptual designs for HRMs.

Geometric effects of fuel regression rate in hybrid rocket motors

The geometric configuration of the solid fuel is a key parameter affecting the fuel regression rate in hybrid rocket motors. In this paper, a semi-empirical regression rate model is developed to investigate the geometric effect on the fuel regression rate by incorporating the hydraulic diameter into the classical model. The semi-empirical model indicates that the fuel regression rate decreases with increasing hydraulic diameter and is proportional to dh?0.2 when convective heat transfer is dominant. Then a numerical model considering turbulence, combustion, solid fuel pyrolysis, and a solid–gas coupling model is established to further investigate the geometric effect. Eight motors with different solid fuel grains are simulated, and four methods of scaling the regression rate between different solid fuel grains are compared. The results indicate that the solid fuel regression rates are approximate the same when the hydraulic diameters are equal. The numerical results verify the accuracy of the semi-empirical model.

固液火箭发动机燃烧稳定性试验(英文)

针对固液火箭发动机低频非声振荡燃烧问题进行了发动机试验研究。介绍了挤压式过氧化氢(H2O2)-低密度聚乙烯(LDPE)固液火箭发动机试验系统及试验发动机结构特点,并分析了固液火箭发动机不稳定燃烧特点。试验结果表明,通过提高氧化剂供应管路的刚性、改变喷嘴雾化效果,可解决低频振荡燃烧。

固液混合火箭发动机中的关键技术及其发展

根据固液混合火箭发动机的结构特点及工作过程,分析了它所具有的优、缺点。纵观国内、外的研究情况,总结出固液混合火箭发动机中的关键技术和难点,主要包含:(1)发展新的高能固体燃料和氧化剂,提高固体燃料的退移速率;(2)测量工作过程中固体燃料退移速率;(3)研究不同类型固体燃料热解机理及规律;(4)建立固体燃料退移速率的公式;(5)抑制不稳定燃烧;(6)建立缩比准则;(7)发展混合火箭发动机工作过程模拟的数值模型等。分析了国内、外在研究中采用的技术途径、取得的成果,以及今后的发展方向。

固液混合火箭发动机燃烧室和喷管流动数值模拟

固液混合火箭发动机是采用液体作为氧化剂,固体作为燃料的一种典型的混合火箭发动机。固液混合火箭发动机中的燃烧和流动问题是固液混合火箭发动机设计中的关键问题,对固液混合火箭发动机的燃烧室和喷管进行一体化计算很有必要。利用二维轴对称N-S方程和组分方程对选用液氧/端羟基聚丁二烯推进剂的固液混合火箭发动机的燃烧室和喷管进行了一体化计算。计算采用LU时间隐式格式、MUSCL空间离散和Van Leer矢通量分裂方法,采用有限速率化学反应模型,对化学源相进行了点隐式处理。计算中分别采用了一步化学反应模型和两步化学反应模型方案,计算了多个氧化剂流速和燃烧室压强下的燃烧室和喷管流场分布,对化学模型进行了选择,为固液混合火箭发动机的设计提供了依据。

国外固液混合火箭发动机关键技术研究进展

固液混合火箭发动机兼具固体火箭发动机简单和液体火箭发动机可控的优点,必将成为未来运载火箭、亚轨道飞行器、轨道转移系统上有效的推进系统之一。简要介绍了固液混合火箭发动机的发展历史和原理模型,重点介绍了2008年以来国外固液混合火箭发动机亟待解决的不稳定燃烧的理论研究方法、提高固体装药的退移速率等关键技术及实验研究进展情况,最后对HRM的工程化方法进行了简要介绍。

低密度聚乙烯的热解试验研究

通过热重法(TG)和差示扫描量热法(DSC)研究了低密度聚乙烯(LDPE)燃料的热解反应,采用裂解气相色质联用(PyGC-MS)实验装置研究了热解产物分布随温度的变化趋势,得到了LDPE燃料的热解特性。结果表明,LDPE的热解反应为动力学一级反应;热解产物成分浓度随热解温度的升高而变化。

PE燃料热解过程对H_2O_2-PE固液发动机点火的影响

通过85%H2O2-PE固液火箭发动机中含有液体层的燃料热解模型,计算了PE燃料的热解过程,分析了燃料热解过程对点火时间的影响,计算结果和试验值一致。H2O2的流量对点火过程起决定性作用,当氧化剂流量低于最小点火流量时,发动机将无法点火。