基于格子玻尔兹曼方法的旋翼三维气动力学分析

提出运用格子波尔兹曼方法（Lattice Boltzmann Method, LBM）分析微型旋翼 的气动特性。以APC11x47 型螺旋翼为分析对象, 实物的结构信息通过三维激光扫描和 逆向工程的方法输入计算机, 使用基于LBM 的软件计算其不同转速下的螺旋翼产生的 升力,最后得到其升力系数。     

燃油温度对离心喷嘴雾化特性影响的试验

对某中心分级燃烧室头部的三种型号离心喷嘴副油路燃油喷入静止大气中的雾化特性进行了试验研究,获得了不同进口燃油温度(-40~80 ℃)和供油压差对燃油雾化特性的影响规律。利用相位多普勒粒子测量技术（PDPA）测量了沿流向距离离心喷嘴出口30 mm平面上的油雾特性,并利用激光粒度分析仪对试验结果进行了进一步验证。研究结果表明:①离心喷嘴的流量数随燃油温度的升高而逐渐减少,且在低温段下降幅度更大;②测量平面上沿直径各处的Sauter平均直径（SMD）在低温段随燃油温度的升高而减小,且油锥中心处的SMD下降幅度更大;③利用激光粒度分析仪测得的油雾场粒径分布在一定程度上验证了PDPA测量结果的正确性,液滴特征直径和液滴尺寸分布系数随供油压差的增大而减小。     

燃油温度对离心式喷嘴雾化性能影响

以离心式压力雾化喷嘴为研究对象,对不同压力下燃油温度对航空煤油雾化特性的影响进行了实验测试和数值模拟研究,获得了燃油在喷嘴内的流动特性及温度、压力对燃油雾化特性参数的影响规律。实验研究了燃油温度变化范围在-20 ℃至50 ℃的雾化特性,数值模拟对燃油温度在-50 ℃至50 ℃范围内喷嘴内燃油的流动特性及燃油的雾化特性进行了数值模拟。结果表明:燃油压力对雾化特性影响不大;在所研究的温度范围内,温度增加会导致雾化角增大、索太尔平均直径（SMD）减小、周向分布不均匀性增大,在-20 ℃升至50 ℃时SMD由45 μm降低到30 μm;油膜厚度会随燃油温度的降低而增厚,有利于提高燃油周向分布均匀性,但会导致雾化液滴直径增大。     

全电推进卫星小推力变轨策略星上计算方法研究

全电推进卫星星上自主变轨,是全电推进卫星重要的发展方向。为了获得运算量小、计算简单、可以星上计算且变轨时间最短的小推力变轨策略,研究了Lyapunov反馈制导律和推力矢量分段固定法两种方法。基于Lyapunov反馈制导律的变轨策略,权重系数在地面进行优化,推力指向星上实时计算,在标称任务工况下变轨时间比理论最优解加长8.18%。推力矢量分段固定法变轨策略更为简单,每10天星上对两个关键控制参数Ψ1,Ψ2进行修正,推力指向变化规律恒定,变轨时间比理论最优解加长7.43%。两种方法都具有任务适应性好和计算简单的优点,Lyapunov反馈制导律对姿态控制能力要求较高,推力矢量分段固定法姿态控制要求容易满足,后者更适合于卫星应用。     

Transient flow characteristics and performance of a solid rocket motor with a pintle valve

The pintle valve is currently the most promising technology among all thrust control methods for solid rocket motors. Pintle structure and working condition play a critical role in the successful operation of a pintle motor. Here, 2D transient simulations of a pintle motor using dynamic meshing are performed. Reynolds-averaged Navier–Stokes equations are solved with the implementation of an RNG k–ε turbulence model. In cold flow test, emphasis is placed on the effect of pintle structure, and in hot flow test, emphasis is placed on the effect of propellant pressure exponent. Validation is performed first by comparing the present results with available cold-test experimental data. This shows that the transient simulations can provide good predictions for pintle motors with a relative error of less than 2% in terms of the chamber pressure. It can be found that, when the gas supply system is different, the working principles and conditions of pintle motors are different. The feedback process in propellant combustion has a significant impact on its operation and the effect on the pintle motor performance of different pintle structures is achieved by different variations in the equivalent throat area. Finally, the pressure exponent is an important parameter in hot flow test and changes of thrust in hot flow test are not monotonic, because changes in the flow field and motor performance are asynchronous.     

摩擦力和气体压力对双节套筒式喷管展开的影响

分析计算了作动筒内摩擦力、延伸出口锥端面密封摩擦力、大气压力及喷管中燃气压力对双节套筒式延伸喷管展开特性的影响,其分析计算方法对双节套筒式延伸喷管设计及展开时间预估有指导作用。     

双节套筒式延伸喷管运动分析

用质心运动定理对双节套筒式延伸喷管的展开过程进行分析,得到一个二阶常微分方程;然后用四阶龙格－库塔法进行数值求解,可得到延伸出口锥的轴向位移、速度和加速度及展开时间,并以单节套筒式延伸喷管为例;计算结果与试验基本吻合．     

膨胀循环液体火箭发动机推力调节阀仿真研究

根据某膨胀循环液体火箭发动机推力调节阀的结构及工作原理,通过理论分析建立了推力调节阀的数学模型,并利用AMEsim软件构建了推力调节阀的仿真计算模型,对其进行了仿真计算.计算了发动机额定工况、高工况和低工况参数下推力调节阀内部各压力及流量参数,并对推力室室压、调节阀出口压力和氢主文氏管入口压力变化引起的调节阀主阀流量变化趋势进行了计算分析,得到了调节阀内部各压力参数及流量的变化规律.     

Aerospike nozzle contour design and its performance validation

A simplified design and optimization method of aerospike nozzle contour and the results of tests and numerical simulation of aerospike nozzles are presented. The primary nozzle contour is approximated by two circular arcs and a parabola; the plug contour is approximated by a parabola and a third-order polynomial. The maximum total impulse from sea level to design altitude is adopted as objective to optimize the aerospike nozzle contour. Experimental studies were performed on a 6-cell tile-shaped aerospike nozzle, a 1-cell linear aerospike nozzle and a 3-cell aerospike nozzle with round-to-rectangle (RTR) primary nozzles designed by method proposed in present paper. Three aerospike nozzles achieved good altitude compensation capacities in the tests and still had better performance at off-design altitudes compared with that of the bell-shaped nozzle. In cold-flow tests, 6-cell tile-shaped aerospike nozzle and 1-cell linear aerospike nozzle obtained high thrust efficiency at design altitude. Employing gas H₂/gas O₂ (GH₂/GO₂) as propellants, hot-firing tests were carried out on a 3-cell aerospike nozzle engine with RTR primary nozzles. The performance was obtained under two nozzle pressure ratios (NPR) lower than design altitude. Efficiency reached 92.0–93.5% and 95.0–96.0%, respectively. Pressure distribution along plug ramp was measured and the effects of variation in the amount of base bleed on performance were also examined in the tests.     

变推力固体火箭发动机喉栓烧蚀试验研究

针对三维C/C和钨渗铜两种不同材质,开展了发动机喉栓的静态烧蚀及动态烧蚀研究,揭示了静态和互变条件下喉栓发动机的烧蚀规律.试验结果表明,高压静态条件与互变过程相比,喉栓烧蚀率有明显差别,高压静态比互变过程烧蚀更严重,互变过程引起的热环境变化没有造成烧蚀异常增大.因此,在工程中可采用高压静态烧蚀试验来考核喉栓材料,简化试验系统;在文中试验条件下,钨渗铜喉栓最大径向烧蚀率为0.085 mm/s,三维编织C/C材料最大径向烧蚀率为0.545 mm/s,钨渗铜比C/C材料更适用于喉栓;发动机非轴对称结构、粒子冲刷和沉积现象对烧蚀影响较大,采用同轴结构可改善流动的对称性,有利于进一步研究其他因素对烧蚀的影响.