摘要
对自燃推进剂 ( MMH/ NTO)初始温度对火箭发动机燃烧稳定性的影响进行了研究 ,从推进剂初始温度对蒸发速率的影响规律出发 ,发展了初始温度大小影响蒸发速率的物理模型。燃烧流动过程应用圆柱坐标系下的两相湍流化学反应 Navier- Stokes方程来描述 ,控制方程是用有限体积法在任意曲线坐标系下进行离散 ,计算采用 TTM方法生成的正交网格。完善了一种压力隐式算子分裂算法 ,使之应用到燃烧过程和不稳定燃烧中 ,提高了计算的精度和稳定性 ;以蒸发作为燃烧速率控制过程 ,由 MMH的分解蒸发速率来控制。用蒸发和分解的时滞来分析燃烧不稳定性。应用CFD技术发展了评定燃烧稳定性的脉冲枪模型 ,之后对推进剂初始温度对燃烧稳定性的影响进行了数值研究 ,得到其对振荡的敏感分析 ,并给出燃烧稳定性的极限图 ,说明了该物理模型和算法的可靠性。
The effect of initial propellant temperature on rocket engine combustion stability for hypergolic propellant (MMH/NTO) is studied. A physics model for initial propellant temperature effecting evaporation rate has been developed in this paper. The Navier Stokes equations in the cylindrical coordinates are employed to describe the turbulent combustion and flow process in the chamber of rocket engine, and the N S governing equations are transfered into arbitrary curvilined coordinate and discritized by finite volume method. Orthogonal body fitted grid generated by TTM method is used. The Pressure Implicit Splitting Operator algorithm is extended, and the accuracy of computation is improved and numerical stability is enhanced. Fuel vaporization is identified as the combustion rate limiting process, and determines the decomposition rates of monopropellant gas phase of MMH. And the time lag of evaporation and decomposition has been utilized to analyze the combustion instability. Computational Fluid Dynamics(CFD) technique is used to stability testing ,and pressure pulse gun methodology has been modeled numerically. Initial propellant temperature as driving mechanism for combustion instability has been studied .The sensitivity to fluctuations is determined, and a combustion stability limiting map is given. The results show that the model and the algorithm are suitable.
出处
《导弹与航天运载技术》
2000年第4期32-37,共6页
Missiles and Space Vehicles
