Study on the Effect of Solid Particles on Jet Flow and Base Thermal Environment for Solid-Liquid Bundled Rocket

Compared with using liquid rocket engines,there are a lot of high-temperature solid particles in the solid-liquid bundled rocket,which make the rocket base thermal environment worse.In order to study the influence of high-temperature solid particles on the base thermal environment,firstly,the effect of particle diameter on the jet distribution and the thermal environment in a single solid motor jet was analyzed using a numerical simulation method,and the results were compared with those of a ground test.Further,the effects of high-temperature solid particles on the jet and the thermal environment of the solid-liquid bundled rocket were analyzed and compared with flight data.The results show that high-temperature solid particles can increase the jet temperature and reduce the jet velocity.The larger the particle diameter,the greater the impact on the jet core temperature.The role of high-temperature solid particles cannot be ignored in the study of the base thermal environment.

Research on the Initial Ignition of the Underwater Launching Solid Rocket Motor

The aim of this investigation is to research the initial ignition of the underwater-launching solid rocket motor.The MIXTURE multiple-phase model was set to simulate the initial ignition.The water vaporization was researched and the energy transfer was added to the energy equations.The flow field and the vaporization were calculated coupled.The initial ignition process of the underwater solid rocket motor is obtained and the vaporization influence to the underwater launching is analyzed.The ＂neck＂,＂inverted jet＂ and ＂eruption＂ phenomenon of the bubble are observed.The bubble increases more rapidly because the steam mass added to the fuel.The temperature is lower considering the vaporization because the steam enthalpy is lower than the fuel enthalpy and the flow field of the initial ignition of the underwater-launching solid rocket motor is accordant well to the reference.

Numerical simulation of multi-phase combustion flow in solid rocket motors with metalized propellant

Multi-phase flowfield simulation has been performed on solid rocket motor and effect of multi-phases on the performance prediction of the solid rocket motor(SRM) is investigation.During the combustion of aluminized propellant,the aluminum particles in the propellant melt and formliquid aluminum at the burning propellant surface.So the flow within the rocket motor is multi phase or two phase because it contains droplets and smoke particles of Al2O3.Flowsi mulations have been performed on a large scale motor,to observe the effect of the flowfield onthe chamber and nozzle as well.Uniform particles diameters and Rosin-Rammler diameter distribution method that is based on the assumption that an exponential relationship exists betweenthe droplet diameter,dand mass fraction of droplets with diameter greater thandhave been used for the si mulation of different distribution of Al2O3 droplets present in SRM.Particles sizes in the range of 1-100μm are used,as being the most common droplets.In this approachthe complete range of particle sizes is dividedinto a set of discrete size ranges,eachto be defined by single streamthat is part of the group.Roe scheme-flux differencing splitting based on approxi mate Riemann problem has been used to si mulate the effects of the multi-phase flowfeild.This is second order upwind scheme in which flux differencing splitting method is employed.To cater for the turbulence effect,Spalart-All maras model has been used.The results obtained show the great sensitivity of this diameters distribution and particles concentrations to the SRMflowdynamics,primarily at the motor chamber and nozzle exit.The results are shown with various sizes of the particles concentrations and geometrical configurations including models for SRM and nozzle.The analysis also provides effect of multi-phase on performance prediction of solid rocket motor.

Numerical study on ignition start-up process of an underwater solid rocket motor across a wide depth range

Solid rocket motors have important applications in the propulsion of trans-media vehicles and underwater launched rockets.In this paper,the ignition start-up process of an underwater solid rocket motor across a wide depth range has been numerically studied.A novel multi-domain integrated model has been developed by combining the solid propellant ignition and combustion model with the volume of fluid multiphase model.This integrated model enables the coupled simulation of the propellant combustion and gas flow inside the motor,along with the gas jet evolution in the external water environment.The detailed flow field developments in the combustion chamber,nozzle,and wake field are carefully analyzed.The variation rules of the internal ballistics and thrust performance are also obtained.The effects of environmental medium and operating depth on the ignition start-up process are systematically discussed.The results show that the influence of the operating environment on the internal ballistic characteristics is primarily reflected in the initial period after the nozzle closure opens.The development of the gas jet in water lags significantly compared with that in air.As the water depth increases,the ignition delay time of the motor is shortened,and the morphology evolution of the gas jet is significantly compressed and accelerated.Furthermore,the necking and bulging of the jet boundary near the nozzle outlet and the consequent shock oscillations are intensified,resulting in stronger fluctuations in the wake pressure field and motor thrust.

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.

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.

Internal ballistic simulation of multi-burning-rate solid rocket motor based on parameterized feature CAD modelInternal ballistic simulation of multi-burning-rate solid rocket motor based on parameterized feature CAD model

Internal ballistic simulation(IBS)method of multi-burning-rate solid rocket motor(SRM)was developed based on 3-D burning regression method by parameterized feature CAD model(PFCADM)and lumped parameter,in consideration of time-dependent,erosive-burning-effect from internal ballistic numerical algorithm.By driving multi-parameter CAD model based on PFCADM,the approach is capable of conducting the geometric regression simulation of various grain combinations of complex configurations with different burning rates.Through suitably simplifying the internal ballistic numerical algorithm,the problems of coupling geometric regression simulation of sub-grains of different burning rates and high computational consumption of internal ballistic calculation were solved.One tri-burning-rate grain motor,which had been firing-tested,was used as the validation case of simulation.The results show that,with the 3-D grain regression model and sufficient accurate internal ballistic algorithm,the method realizes IBS of the case in low computationalconsumption prediction of its performance within the accuracy of 2% during 1hclock-time.The application of the method provides a practical approach to aid SRM design of multi-burning-rate grain.

脉冲触发对固体火箭发动机内弹道和瞬态流场特性的影响

为研究脉冲触发对固体火箭发动机内弹道及瞬态流场特性的影响,利用数值计算方法建立脉冲触发器内弹道模型、发动机侵蚀燃烧模型以及动态流场仿真模型,对固体火箭发动机脉冲触发过程进行瞬态流场仿真。研究结果表明:与实验结果相比,压强预示误差小于5%;脉冲触发使得燃烧室内横向气流速度增大,靠近脉冲端的推进剂发生了侵蚀燃烧,侵蚀燃烧对压强抬升的贡献高达44%;脉冲触发时,越靠近脉冲入口的推进剂受侵蚀燃烧越严重,侵蚀比最大可达7.32;改变脉冲药量研究发现,脉冲药量越大,发动机压强峰值越大,压强抬升率和脉冲结束后的衰减率越大。

可渗透壁自适应高度补偿喷管工作机理数值研究

提出了一种带有可渗透壁面的自适应高度补偿概念喷管(可渗透喷管),对该喷管的典型工作状态进行了数值模拟并对其高度补偿机理进行了解释。数值模拟结果显示,在低落压比条件下,外界大气透过喷管可渗透段进入喷管内部,抑制了喷管壁面处的回流产生,削弱了壁面处分离斜激波的强度,从而提高了喷管的低空推力性能。而在高落压比条件下,喷管内部燃气透过可渗透段壁面泄出,削弱了燃气的膨胀能力,使得喷管的高空推力性能有所损失。当可渗透喷管的低空推力性能提升幅度大于其高空推力性能损失时,喷管即具备了高度补偿能力。对喷管的推力构成分析表明,低落压比时,喷管可渗透段处外界大气流入所导致的喷管内压升高是其具有高度补偿能力的主要原因。基础扩张比的大小对可渗透喷管高度补偿性能影响显著,可渗透喷管比传统喷管的推力提升幅度最高可达34%,不同基础扩张比的可渗透喷管低空推力补偿能力相差最大25%。将不同基础扩张比喷管的高度平均比冲进行对比可发现,存在最优的喷管基础扩张比,该可渗透喷管高度平均比冲比传统喷管高度平均比冲高3%。

基于RBF代理优化的固体火箭发动机喷管型面设计

针对固定扩张比与扩张段长度的二维轴对称固体火箭发动机喷管进行扩张段型面优化,优化目标为喷管推力最大化,优化参数为贝塞尔曲线控制点的径向位置,优化方法采用径向基函数(Radial Basis Function, RBF)代理优化算法。采用纯气相与两相流两种模型分别进行优化设计,纯气相的结果表明,对于10个控制点表达的贝塞尔曲线,优化后的推力提高了1.64%。以此优化型面为初始型面,增加控制点个数至16个,二次优化后的推力又提高了0.095%。增大优化参数范围,同时引入判断拐点的约束,对于10个控制点表达的贝塞尔曲线进行单轮优化,结果同上述经过两轮优化之后的结果相近,优化后的喷管推力提高了1.78%,说明算法具有较强的稳定性。通过对不同控制参数个数的贝塞尔曲线优化过程的对比,给出了合理选择控制点个数的方法与建议。两相流的优化结果表明,由于颗粒的滞后影响造成了两相流损失,两相流喷管的推力小于纯气相喷管,但两相流喷管优化后的推力较优化前初始型面的推力提高了1.87%,略高于纯气相喷管。RBF代理优化算法适用于由任意数量控制点组成的贝塞尔曲线表达的喷管扩张段型面优化,并有较高的效率与较强的稳定性。

不同摆角超声速分离线喷管烧蚀数值研究

建立了适用于超声速分离线喷管的三维烧蚀仿真方法,与试验结果对比验证了其有效性,对不同摆角的喷管烧蚀状况进行了数值仿真。结果表明,由于喷管型面改变,分离线附近烧蚀率出现骤升骤降的复杂变化。分离线前侧壁面大部分位置烧蚀率受摆角影响较小,仅下侧点位置附近由于燃气流动受阻烧蚀率随摆角增大而增大,当摆角为5°时,为喷管整体烧蚀最严重区域,烧蚀率为79.4μm/s。分离线后侧壁面烧蚀率受摆角影响较大,且整体高于前侧。在分离线下游上侧边界线,烧蚀率会快速出现一个峰值,且出现位置随摆角增大后移,从0°的59.7 mm后移至5°的65.3 mm处。摆角为5°时,由于下侧激波切割了流场,导致了上侧边界线x=130 mm燃气流动出现分离现象,烧蚀率在此处回升到5.8μm/s。

固体发动机水下点火尾流场计算

为研究固体发动机水下点火初期喷管内燃气泡的流动和发展过程及流场结构,利用CFD方法通过fluent分析软件对固体火箭发动机水下点火过程气液两相流场进行数值模拟计算,得到了发动机尾流场燃气马赫数、压力、温度等瞬态阶段演变情况。结果表明,发动机水下点火时,燃气与水之间产生强烈地相互作用,射流通道沿轴向存在周期性胀缩,一个周期内相继出现膨胀—液体回压—回流—颈缩—继续膨胀现象,随着流场发展涨缩强度逐渐减弱,颈缩位置向下游移动;燃气射流初期,轴向压力最高点在燃气泡头部位置;点火初始时刻喷管出口出现压力峰值;全计算域内温度在一定范围内振荡,沿轴向远离出口区域有剧烈的传热和明显的气液混合。

入口流场不均匀性对固体火箭发动机大膨胀比喷管内流场的影响

在高机动飞行过程中,固体火箭发动机承受长时间的小过载会导致喷管入口处燃气方向改变,而药柱空间非均匀燃烧又会导致燃气密度分布不均匀,从而影响喷管入口处质量通量分布。二者均会造成入口燃气流动非均匀性,这是引起喷管流场不均匀性的主要原因之一,直接影响发动机性能。以膨胀比为100的固体火箭发动机喷管为研究对象,改变入口处质量通量分布规律与入射方向,结合计算流体力学软件研究入口流场不均匀性对喷管内流场的影响规律。结果表明:当喷管入口质量通量呈抛物面分布时,入口的不均匀性对喷管内流场影响较小,在喷管喉部位置不均匀效应基本消除;由俯仰过载导致燃气非垂直入射时,喷管入口位置流场周向不均匀性高达23%;随着燃气充分膨胀,不均匀效应减弱,在喷管扩张比达到92.8时,流场基本均匀,但是在有限扩张比条件下由入射角度导致的流场不均匀效应无法完全消除。

固体火箭发动机燃气射流与超声速来流相互作用数值研究

针对固体火箭发动声速来流机续航级在超声速来流中的燃气射流及外流场特点,利用数值模拟方法,分别对不同工况下固体火箭发动机尾焰及外流场进行流场结构定性分析,得到该流场结构大体可分为外场区、回流区及尾流区;通过对比不同高度轴线上参数,得到随着工作高度增加,喷管流动分离现象消失,波节减少,首个入射激波形成距离喷管出口距离增加,激波波面速度增大;通过对比回流区壁面参数,得到随着工作高度的增大,回流区壁面温度及马赫数增大,低压回流区导致喷管的实际工作压力低于环境压力30%以上;当喷管处于欠膨胀工作状态时,回流区流体马赫数呈现先减小后增大趋势,且靠近尾焰端受工作高度影响显著。

支板火箭构型对引射进气与主次流相互作用特性影响实验研究

一次火箭射流与二次空气流之间的主次流相互作用是火箭基组合循环(RBCC)发动机引射模态工作过程的主导机制,对引射模态性能具有关键性影响。基于典型支板火箭RBCC发动机特征流道模型,进行了三种不同支板火箭构型的宽工况冷喷流引射试验,获得了引射进气特性、主次流相互作用特性与规律。研究结果表明:随着一次流流量、总压比增加,二次流引射进气马赫数逐渐增大,流量逐渐增加至某一最大值,而引射比持续单调减小;无量纲分析显示,单矩形喷管的支板火箭构型相对于双圆形喷管构型具有更强的引射能力,在相同一次流流量与喷管喉道面积时能提高10%~40%的流量通量比,且支板尾缘带波瓣凹槽结构有利于提升引射能力。从主次流内流过程来看,随着一次流流量、总压比增加,内流道压力分布整体上逐渐降低,马赫数分布逐渐提高,反映出内流加速降压、引射进气流量增加的过程,同时下游的引射增压比呈增大趋势,以更大的总压损失为代价;相同一次流流量时,单矩形喷管构型相对于双圆形喷管引起的内流道压力更低、马赫数更高,实现基本相当的引射比时总压恢复系数相对更高。随着主次流总压比的增加,主次流的速度比和对流马赫数呈现减小的趋势,而压比和密度比则有所增加;双圆形喷管构型时速度比和对流马赫数明显高于单矩形喷管构型,而压比和密度比则相对更低;这些参数梯度变化在初始状态层面上决定了主次流相互作用的过程及其宏观特性。

长尾管固体火箭发动机内部流动特性仿真分析

为分析长尾管固体火箭发动机的内部流动特性,编制微分内弹道程序进行发动机内弹道解算;同时采用自定义函数建立质量进口条件、动网格技术模拟燃面推移、欧拉-拉格朗日模型结合粒子随机游走模型,对无长尾与有长尾结构固体火箭发动机三维两相瞬态内流场展开对比研究,对比了发动机的性能,分析了长尾结构对气相流动特性与不同直径颗粒的运动分布的影响。分析结果表明:内弹道与流场仿真的发动机平衡压力基本一致。有长尾结构带来了沿程阻力损失,内部平衡压强要高于无长尾结构发动机压力,导致了装药燃烧加快,输出推力反而减小;燃气流经长尾区域流速增加明显,但粒子作用会导致轴线附近流速与温度的震荡,温度整体变化小,故此区域热防护应该重点考虑;由于惯性,粒子在长尾区域先汇集后发散,随粒径增加,粒子汇集区越靠前,散布程度越大,气流对粒子运动作用越小。

固体火箭发动机气-固两相近壁湍流特性大涡模拟研究

固体火箭发动机内流场是典型的两相湍流流动,其流动特性对推进剂的燃烧影响极大。工程实践中大多采取雷诺平均(RANS)方法计算两相流问题,精度不高且无法解析流场中的湍流结构。采用大涡模拟(LES)与离散颗粒模型(DPM)分别对纯气相和气-固两相条件下的发动机内流场流动过程进行了数值模拟;通过对时均流场与瞬态流场的分析以及对近壁面(即注入表面)的物理参数开展时空平均,重点研究了近壁湍流特性;采用时空相关性分析与降阶变分模态分解(RVMD),重点探究了近壁面的旋涡脱落。结果表明,颗粒集中在内流场的中部,由于惯性的差异,越大的颗粒越接近中心;由于颗粒相与气相之间存在速度滞后和热量交换,颗粒相的注入会使内流场的轴向速度和温度降低、压强升高;发动机内流场会出现明显的壁面涡脱落,而颗粒相的引入会增加内流场的不规则扰动,从而使得旋涡脱落、破碎的位置提前,并减小涡结构的尺寸,使观察到的旋涡更加破碎;颗粒相的引入还能够减缓近壁面附近的温度梯度,抑制不稳定燃烧现象;同时,颗粒相的存在会使近壁面的径向速度降低、压力升高,降低和升高的幅度约为1.8%。对近壁面的压力振荡分析表明,颗粒相的存在能够抑制高频压力振荡的产生,使得纯气相条件下1853 Hz的振荡模态消失,只保留1425 Hz的模态。纯气相和气-固两相流场的脱涡位置分别为x/m=0.54、x/m=0.44,说明气-固两相流场旋涡脱落的位置更靠前。

固体火箭发动机喷管扩散段破损流场仿真研究

为获取某型带长尾管的固体火箭发动机在工作过程中扩散段不同破损程度对其性能影响规律和影响机理,采用CFD方法对其进行建模仿真分析,通过对比不同破损程度下的流场分布和推力变化,揭示了该型发动机喷管扩散段破损对其流场和性能的影响规律;研究发现,气体从缺口处流出部分在轴向截面上基本呈扇形,在扇形弧线和半径相交处存在角结构;流场等值面可以划分为缺口处等值面、非缺口处等值面和两者之间相互作用等值面三部分;喷管在扩散段端口破损后会对下游流场产生影响,对上游流场影响非常小;随着扩散段端部缺口的增大,喷管推力方向与喷管轴线夹角逐渐增大。

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本文以提升固体火箭发动机研制活动的标准化程度为目标,探索强制标准嵌入设计流程的信息化管理模式及途径,开展固体火箭发动机强制标准体系的优化研究、标准件库的建立和优化、强制标准嵌入设计流程的信息化管理模式研究、嵌入强制标准的信息化系统集成开发研究以及实施效果评估研究。试点型号的应用和对实施过程的监督表明,实行强制标准嵌入流程,实现了岗位与强制标准的映射,提升了强制标准的使用率和产品化率。

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