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针对现有弹用固体火箭冲压发动机普遍采用的固定几何不可调节喷管,基于流量平衡的基本原理,建立了其理论设计及性能评估的数学模型。结合当前中远程空空导弹提出的Ma=2~3.5宽速度范围设计需求,运用所建立设计模型对实例设计方案开展了计算分析。结果表明,现有固定几何喷管本质上是为满足低速正常接力而折中设计出的,在高速巡航时,因扩张比偏小,不仅喷管出口气流速度和冲量小,而且导致燃烧室压强降低,还额外造成进气道结尾正激波总压损失加大,不能将进气道保有的捕获高速来流动能充分发挥出来。原设计方案在Ma=3.5高速巡航时,进气道实际总压恢复性能对比方案中的最大总压恢复性能水平,相对损失幅度高达42.67%,而且冲压发动机推力与其可能达到的最大值对比,相对损失幅度也高达31.8%。因此建议采用喷管调节技术来解决此类问题。
Aiming at the fixed geometric non-adjustable nozzle commonly used in the existing solid propellant ramjet engine, a mathematical model of its theoretical design and performance evaluation is established based on the basic principle of flow balance. Combined with the design requirements of Ma = 2 ~ 3.5 wide speed range proposed by the current medium-to-long-range air-to-air missile, the design examples are used to carry out the calculation and analysis of the design examples. The results show that the existing fixed geometry nozzle is essentially designed to meet the normal low-speed relay. In the high-speed cruising, due to the small expansion ratio, not only the nozzle exit airflow velocity and impulse are small, but also the pressure in the combustion chamber Reduce, but also caused an additional increase in total pressure loss of the positive shock wave at the end of the intake passage, and can not fully exert the flow captured by the intake passage at high speed. In the original design scheme, the maximum total pressure recovery performance in the actual total pressure recovery performance comparison of the inlet at Ma = 3.5 high-speed cruise with the relative loss rate as high as 42.67%, and the comparison between the thrust of the ramjet engine and its possible maximum, The loss rate is as high as 31.8%. Therefore, nozzle adjustment technology is suggested to solve such problems.