HTPB推进剂脱湿与力学性能的相关性研究

针对载荷作用下影响复合推进剂力学特性的脱湿问题,采用等速拉伸和CCD显微分析的试验方法,研究了不同拉伸速率下的脱湿损伤演化过程。建立了粘弹性本构模型,利用细观力学及界面力学的理论,分析试验测得的宏观力学性能发生发展的内在细观原因。结果表明:颗粒/基体的界面脱湿是宏观应力应变曲线非线性的重要原因,也直接导致材料泊松比的下降;界面脱湿的损伤程度由应变值决定,并与应变率具有一定的相关性,泊松比也是定量表征脱湿的重要参数。     

HTPB推进剂粘聚区本构模型反演识别研究

端羟基聚丁二烯（HTPB）推进剂在拉伸过程中裂尖存在一个明显的非线性粘聚区,粘聚区本构模型的精度影响着推进剂装药裂纹起裂和扩展过程的数值仿真结果。为了准确地获得HTPB推进剂的粘聚区本构模型,建立了基于实验的反演识别方法,该方法通过实验获取粘聚区断裂能和断裂强度参数,使用有限元模型更新方法得到粘聚区本构曲线形式参数。将获得的本构模型应用于裂纹起裂和扩展过程仿真,研究结果表明所建立的粘聚区本构参数获取方法简单可行;所获得的粘聚区本构参数可以准确地模拟出HTPB推进剂裂纹起裂和扩展过程。     

利用嵌金属丝药柱调节固体火箭发动机推力的计算研究

提出一种组合应用嵌金属丝药柱和引入负热流的方法实现火箭发动机推力调节的技术方案,即通过在金属丝暴露在燃气部分加入负热流改变沿金属丝燃速来改变发动机工作压力,实现对推力的调节。可行性研究的初步计算结果表明,该方案能够在一定范围内实现对火箭发动机工作特性的调节,推力大小随负热流密度的增加而减小;从加入负热流密度到推力重新稳定有一段滞后时间,滞后时间随负热流密度的增加而上升。方案的优点是控制系统结构简单,控制箱不需工作在高温条件下。     

HTPB推进剂药柱在变温环境下的累积损伤分析

为研究变温环境对固体火箭发动机的影响,基于热粘弹性模型和累积损伤模型,对某HTPB推进剂药柱在变温环境下的应力响应和累积损伤进行了分析计算。结果表明,药柱损伤随着环境严酷度的增大而增大,温度冲击环境比温度循环对药柱产生的损伤要大。在温度冲击环境下,温度范围、推进剂材料参数对药柱的损伤有较大的影响,平衡模量和膨胀系数的影响有相同的趋势。研究结果可为固体火箭发动机的贮存和使用寿命预估提供技术支持。     

浇铸型固体发动机衬层/药柱界面脱粘低温探伤研究

基于粘弹性理论,用有限元方法分析了浇铸型固体发动机冷冻条件下衬层与药柱的脱粘间隙与脱粘长度、冷冻温度的关系.通过对仿真结果的分析,从理论上得出了固体发动机界面脱粘低温探伤的可行性.通过试验发动机冷冻前后的计算机层析成像(CT)图像的比较,得出了固体发动机界面脱粘低温探伤方法可部分检测到冷冻前检测不到的脱粘.     

海洋温度条件下的固体推进剂应力应变分析

建立以固体发动机壳体、药柱及绝热层等为主要部件的计算模型,基于三维热粘弹性积分型本构关系,考虑不同海域贮存条件的交变温度载荷情况,利用有限元方法计算了固体发动机药柱在海洋交变温度载荷下的力学响应,给出了药柱内应力应变场的分布规律及危险部位。结果表明,在未有效热保护的贮存状态下的药柱,长期的交变温度产生的交变应力载荷,使药柱微观结构受到不同程度的疲劳损伤,势必影响其结构完整性。     

某发动机药柱极端温度发射结构完整性分析

为了对某型大长径比固体火箭发动机在极端温度发射时的结构完整性进行分析,采用三维热粘弹性有限元方法,计算得到了推进剂药柱在内压、轴向过载和极端温度载荷(-40℃和+50℃)联合作用下的位移、应变和应力分布规律。研究结果表明,推进剂药柱在高温和低温发射时位移最大值分别为9.3 mm和16.4 mm,Von Mises应变最大值分别为14.6%和23.5%,依据Von Mises应变准则,安全系数分别为2.74和2.46,满足结构完整性要求。     

炮射导弹发射过程中装药衬垫材料对抗过载能力的影响计算分析

采用基于Total Lagrangian方法的3维粘弹性大变形增量本构关系,用有限元方法对某轴向高过载下的固体火箭发动机装药在发射过程中的应力进行了数值模拟计算。分析了装药底面的衬垫材料、衬垫厚度以及载荷持续时间对装药内部等效应力的影响。结果表明：随着衬垫材料泊松比的提高,装药内部等效应力有下降趋势;泊松比小于0.49时,装药内部等效应力随弹性模量增加略有增加,泊松比0.49时,装药内部等效应力随弹性模量增加先下降后上升趋势;随着衬垫厚度的增加,装药内部最大应力值呈现下降趋势;载荷达到稳定后,载荷持续时间长短对装药内部等效应力影响并不明显。     

某型炮射导弹膛内运动时期药柱的应力应变分析

炮射导弹发射过程中承受上万个g的过载,固体推进剂药柱的应力应变分析极为关键.文中采用有限元方法对药柱在发射过程中的应力应变进行数值模拟,分别对比研究轴向过载、旋转载荷和药柱-挡药板之间摩擦对药柱整体最大von Mises应力和最大应变变化的影响规律.结果表明,考虑旋转和药柱-挡药板之间摩擦的情况下,药柱整体最大yon Mises应力和最大应变都显著增加,而且药柱-挡药板之间摩擦将直接影响药柱最大von Mises应力和最大应变发生的区域.     

Study on Instable Combustion of Solid Rocket Motor with Finocyl Grain

The instable combustion or oscillation combustion which occurs in three high capacity solid rocket motors using high energy composite propellant with finocyl grain is studied. The reasons of the acoustic combustion instability are also discussed. Three engineering methods that can eliminate combustion instability are proposed and discussed. The study shows that the combustion instability mainly depends on the propellant grain shape and nozzle structure. Some measures to reduce the acoustic energy and mass generation rate of combustion gas can be adopted. The test results indicate that the modified rocket motors can significantly eliminate the instable combustion and improve the motor internal ballistic performance.     

Numerical Simulation of Particle/Matrix Interface Failure in Composite Propellant

Interface debonding between particle and matrix in composite propellant influences its macroscopic mechanical properties greatly. For this, the laws of interface cohesive damage and failure were analyzed. Then, its microscopic computational model was established. The interface mechanical response was modeled by the bilinear cohesive zone model. The effects of interface properties and particle sizes on the macroscopic mechanical behavior were investigated. Numerical simulation of debonding damage evolution of composite propellant under finite deformation was carried out. The debonding damage nucleation, propagation mechanism and non-uniform distribution of microscopic stress-strain fields were discussed. The results show that the finite element simulation method based on microstructure model can effectively predict the trend of macroscopic mechanical behavior and particle/matrix debonding evolution process. It can be used for damage simulation and failure assessment of composite propellants.     

Simulation of Aviation Kerosene Combustion in Dual-mode Scramjet Combustor

Supersonic combustion of aviation kerosene is investigated under the flight conditions of Mach number 5 and fuel-air equivalence ratio 0. 551. The trajectories of the fuel droplets and the heat/mass transfer between them are simulated by means of discrete phase model （DPM）. The k-ω model is chosen for turbulence closure and the non-premixed probability density function （PDF） approach is used to calculate the turbulence-chemistry interaction. The calculated wall static pressure and the total pressure loss coefficient are very close to the experiment results. The strut and cavity devices significantly increase the combustion efficiency.     

空气旋转进气对含硼固体冲压发动机二次燃烧性能影响的研究

采用Realizable k-ε湍流模型、单步涡团耗散燃烧模型,应用Fluent软件UDF功能,编写考虑硼颗粒在高速气流中气动剥离效应下的KING点火燃烧计算程序,对典型的双下侧90°含硼固体冲压发动机补燃室进行不同旋转进气下三维两相流动与燃烧数值计算。计算结果表明,当进气道两侧空气同向与反向旋转进入补燃室时,气流产生的旋转均使燃气与空气的混合更充分,燃烧效率更高,并且随着旋流数的增加而增加;对于颗粒燃烧效率与总的燃烧效率,当旋流数小于0.179时,同旋条件高于反旋条件,当旋流数大于0.385时,反旋条件高于同旋条件,当旋流数约为0.2时,同旋与反旋效果相当;对于硼颗粒点火时间,旋流进气减小了点火时间,在旋流数为0.385时最小。     

航空整体结构件加工变形校正技术研究

针对航空结构件的变形问题,提出了包含有加工残余应力的变形整体结构件三维质量模型的重构技术,建立加工变形—校正过程的协同仿真环境,将2个过程有机地结合起来;利用改进的二分法进行校正参数的计算,分析了校正支点位置的确定及不同校正载荷作用方式对于校正效果影响,揭示出加工后工件残余应力与校正残余应力的耦合作用规律;通过大量有限元仿真计算,获得了典型框类工件变形校正的规律性曲线,建立了适用于反弯校正的有限元数学模型,设计了变形校正试验,验证了有限元仿真结果的可行性和有效性。     

Influence of Endwall Contouring on the Secondary Flow in Turbine Nozzle Guide Vane

Numerical simulations are carried out to investigate the effect of the endwall contouring on the secondary flow in turbine nozzle guide vane. The three contoured cascades with the same contouring profile and the different positions where the contoured profile locates at are researched. The results show that the contouring configuration can reduce the aerody- namic losses of the cascade. The flat side takes advantage of a stronger decrease of the losses, compared to the contoured side. The contouring configuration can also inhibit the secondary flow. The contoured cascade in which the contouring profile starts upstream of the airfoil, ends at the middle of the airfoil has the best effect of improving secondary flow.     

Experimental Investigation on Active Cooling for Ceramic Matrix Composite

Compared with conventional materials, the active cooling ceramic matrix composite used in ramjet or scramjet makes their structures lighter in mass and better in performance. In this paper, an active and a passive cooling refractory composite specimens are designed and tested with an experimental facility composed of muhilayer smale scale cooling penel which consists of a water cooling system and a ceramic matrix composite specimen, and a gas generator used for providing lower and higher transfer rate gases to simulate the temperatures in combustion chamber of ramjst. The active cooling specimen can continuously suffer high surface temperature of 2 000 K for 30 s and that of 3 000 K for 9.3 s, respectively. The experiment results show that the active cooling composite structure is available for high-temperature condition in ramjet.