直升机燃油系统增压转自吸供油动特性

通过某型燃油系统增压供油转自吸供油试验,研究了发动机入口处的流量历程。在试验研究和直升机燃油系统增压转自吸供油机理基础上,利用Flowmaster软件建立了某典型直升机燃油系统仿真模型,仿真结果与试验结果吻合较好,验证了模型的正确性。在此基础上,研究了直升机燃油系统增压转自吸供油动特性。结合仿真结果,分析了增压泵出口单向阀的弹簧预紧力、弹簧刚度对发动机入口燃油流量、压力的影响,为燃油系统零部件选型及参数确定提供参考。     

Numerical investigation of the impact of asymmetric fuel injection on shock train characteristics

Numerical simulations are carried out to investigate the impact of asymmetric fuel injection on shock train characteristics using the commercial-code FLUENT. The asymmetry of fuel injection is examined by changing the fuel flow rates of the upper and lower wall fuel injectors. The numerical approach solves the two-dimensional Reynolds-averaged Navier–Stokes (RANS) equations, supplemented with a k-ω model of turbulence. As a result, different ways of fuel injections will always lead to shock train transitions, with the variations of shock train structure, strength and leading edge position. For symmetric fuel injection, the flowfield of the isolator is quite asymmetric with the boundary layer of the upper wall side developing much stronger than that of the lower wall, which is due to the heterogeneity of the incoming flow. Regarding to asymmetric fuel injection with more of lower wall side, though the pressures in the combustor are nearly the same, the first shock of the shock train converts between ‘Distinct symmetric X type shock’ and ‘Obscure and weaker asymmetric shock’ and the shock train leading edge moves upstream with the increase of the asymmetry level. With regard to asymmetric fuel injection with more of upper wall side, ‘incomplete asymmetric X type shock’ occurs and the shock train structures keep nearly the same with low level of fuel injection asymmetry. Unexpected results like unstart will happen when increasing the level of fuel injection asymmetry. And the isolator will come back to normal state by decreasing the differential of upper and lower wall sides fuel injections.     

基于冷热电一体化的舱外航天服生命保障系统性能

基于质子交换膜燃料电池（PEMFC）和热驱制冷,提出一种舱外航天服生命保障系统冷热电一体化方案.在分别建立金属储氢装置、PEMFC、热驱制冷系统和辐射散热器数学模型的基础上,利用冷热电一体化的热力学分析理论和方法进行了典型案例的计算,并重点分析了核心构件热驱制冷装置的参数对舱外航天服生命保障系统性能的影响.与传统的冷热电分产舱外航天服生命保障系统比较,该系统仅消耗84.6g氢气,且不需要向空间排放工质,能源利用率高达85.29%,在工质消耗与能源利用率上有较大的优势.      

复合式共轴直升机过渡模式的操纵策略

根据复合式共轴直升机的构型特点,建立了上/下旋翼、机翼、机身、全动平尾、垂尾（含方向舵）和螺旋桨的气动力模型,并在此基础上建立了过渡过程的配平方程.针对过渡飞行过程的操纵冗余问题,分别采用线性过渡和功率最小优化过渡两种过渡飞行方案,对复合式共轴直升机的过渡飞行过程的操纵策略进行了研究.以样例复合式共轴直升机为例,进行了过渡飞行模式下两种操纵策略的配平计算,并对两种过渡操纵策略下的需用功率、操纵量和姿态进行了对比.结果表明:所述两种过渡方案均能解决操纵冗余问题,并能合理给出过渡模式的操纵量和姿态.在操纵量变化的平稳度、飞行控制的难易程度和飞行品质方面,线性过渡方案更优.      

固体燃料超燃冲压发动机燃速研究进展

对固体燃料超燃冲压发动机燃烧面退移速率（简称燃速）研究现状和进展进行了详细阐述,分别从固体燃料类型、燃烧室构型、理论预估模型、数值模拟及实验研究等方面出发,论述了固体燃料在超声速流动下燃速研究的进展和难点;从亚燃冲压发动机、热防护层、富氧环境下绝热层烧蚀3个方面提炼出可以用于超燃冲压发动机燃速研究的经验和方法:①提出了加强针对固体燃料超声速流动中受热行为、传热传质过程的研究方向;②深入探索了固体超燃燃速性质;③开发了对应的数值软件及系统地进行实验等观点,为国内该领域的研究提供参考.      

基于模态切换的航空发动机容错控制

综合了航空发动机控制和故障诊断方法,设计了基于模态切换的任务级和发动机级模式的容错控制系统.任务级模式在发动机部件故障时,通过切换控制策略和控制模式达到恢复或降低发动机性能的要求;发动机级模式在控制回路失效时,根据故障情况切换到容错控制回路,从而保证发动机继续正常工作.数字仿真结果表明:在稳态或加速过程中出现部件故障时,容错控制系统都能够100%恢复发动机的推力;在发动机中间、慢车和节流状态下,当压气机转速控制回路失效时,容错控制系统能够在3s内平稳切换到风扇转速控制回路.      

基于伴飞模式的补给星轨道设计与优化

针对采用伴飞模式的星座燃料补给问题,本文对其补给星的轨道进行了设计与优化。在给出补给星的轨道设计之后,利用遗传算法对补给星轨道机动所消耗的燃料质量以及每次转移给工作星的燃料质量进行了优化求解,使得补给星燃料利用率达到最高。仿真结果同时给出了燃料利用率与工作星轨道高度以及星座中卫星数目之间的关系,该结论对采用伴飞模式的星座燃料补给研究提供了较好的理论参考。     

Influence of fuel injection position and equivalent mixture ratio on chemical non-equilibrium effects of single expansion ramp nozzle

Chemical non-equilibrium flow was investigated for the scramjet single expansion ramp nozzle(SERN) with a strut-based liquid-kerosene-fueled combustor.Two-dimensional Reynolds-averaged Navier-Stokes(RANS) equations were solved with the species conservation equation for continuous phase and the renormalization group(RNG) k-ε turbulence model.Lagrangian discrete-phase model was analyzed for liquid-kerosene droplets behavior in the supersonic stream.Combustion was simulated by kerosene surrogate fuel's 10-species and 13-step reduced reaction kinetics mechanism with use of Arrhenius's laminar finite rate model.Parametric studies were carried out to estimate the influence of different fuel injection positions and equivalent mixture ratios on the SERN chemical non-equilibrium effects.Numerical calculation results show that the strut-based combustor enables convenient modeling of various SERN entry conditions,which is similar with many preceding investigations,by changing the injector strut position and controlling the mass flow rate of each injector.Chemical non-equilibrium effects function in the whole SERN,especially in the initial flow expansion region,leads to obviously higher SERN performance of the non-equilibrium flow than that of the frozen flow.Furthermore,the distributed fuel injection pattern plays a significant role in enhancing the combustion efficiency in combustor,but weakening the chemical non-equilibrium effects funciton in SERN.Additionally,while the equivalent mixture ratio increases,the SERN thrust coefficient and lift coefficient rise gradually,and the increment of non-equilibrium flow in relation to frozen flow becomes higher as well.To be specific,the equivalent mixture ratio is 0.6,the maximum increment of thrust coefficient and lift coefficient are 11.6% and 25% respectively.      

数值研究低热值燃料环管燃烧室燃烧流场

在任意曲线坐标系下对采用两种不同组分(天然气掺混氮气、一氧化碳)的低热值燃料环管燃烧室燃烧流场进行计算,结果表明低热值燃料的化学反应速率与其组分直接有关,燃料一氧化碳的燃烧效率高于天然气与氮气掺混物燃料,燃料组分变化对燃烧室燃烧效率与出口温度影响很大,因此选用合适的燃料十分重要.计算所得的燃烧室的燃烧效率和出口温度分布与试验数据符合较好,表明所用的数学模型与计算方法合理,计算程序可靠,可为低热值气体燃料燃烧室研制和优化设计提供有用的数据.      

固液火箭发动机工作过程三维数值仿真

根据固体燃料壁面与气相间的流固耦合得出了固体燃料燃速模型,对采用星形装药的H₂O₂/HTPB(hydroxyl-terminated polybutadiene)固液火箭发动机进行了燃烧流动三维数值仿真,得到了流场参数的分布及不同位置的固体燃料燃速,与二维轴对称仿真结果进行了对比.计算结果表明:装药截面的火焰层形状与装药星孔型面形状相似,但火焰层厚度与位置在星根与星角处存在差异;随着轴向位置的增加,氧化剂不断消耗,火焰层向通道中心移动;固体燃料燃速与氧化剂流率及不同装药位置有关,其大小随氧化剂流率的增加而增加,星根处燃速比星角处大;在相同氧化剂流率下,三维星形装药比二维轴对称装药的平均固体燃料燃速大.