Superiority analysis of mass injection pre-compressor cooling technology based on aircraft-engine integration model
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摘要:
为探索射流预冷技术在未来先进航空动力上的应用前景，基于涡喷发动机建立飞发一体化性能仿真计算模型，并选取水作为预冷剂，分析不同飞行状态与预冷方案对飞机作战性能、发动机热力性能及热端部件温度的影响规律。结果表明：启用射流预冷技术可有效增加发动机推力，提升飞机的爬升性能与加速性能，帮助飞机在规定任务内减少机动时间与载荷消耗，同时预冷技术可提升压气机末级引气冷却品质，从而降低涡轮叶片表面温度，增强发动机可靠性。在飞发推力匹配条件与涡轮叶片表面温度约束下，射流预冷可有效提升飞机的极限飞行能力，当预冷剂流量为1 kg/s时，飞机理论升限与最大马赫数分别提升11.67%和10.51%。
Abstract:In order to explore the application prospect of mass injection pre-compressor cooling technology in advanced aero-engines in the future, a simulation model of aircraft-engine integration was built based on turbojet engine, water was selected as the coolant, and the effects of different flight conditions and precooling schemes on aircraft combat capability, engine performance and temperature of hot end component were analyzed. The results showed that the thrust of aero-engine can be increased by mass injection pre-compressor cooling technology, which can improve the climbing and acceleration performance of aircraft and reduce the mission time and load consumption within the assigned task. The surface temperature of turbine blade can be also decreased with the reduction of the temperature of bleed air due to mass injection pre-compressor cooling technology. Given the matching principle of aircraft-engine thrust and the constraint of the surface temperature of turbine blade, the ultimate flight performance of aircraft can be improved effectively by mass injection pre-compressor cooling technology. If the flow rate of coolant was 1 kg/s, the theoretical ceiling and maximum Mach number of aircraft can be increased by 11.67% and 10.51%, respectively.
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图 6 模型与GasTurb软件的性能计算结果对比
Figure 6. Comparison of caculation results between model and GasTurb
图 7 不同高度与预冷剂流量下爬升率与推力变化
Figure 7. Variation of climb rate and thrust with different heights and flow rates of coolant
图 8 不同高度与预冷剂流量下进气流量与耗油率变化
Figure 8. Variation of flow rate of engine and specific fuel consumption with different heights and flow rates of coolant
图 9 不同高度与预冷剂流量下机动时间与耗水量变化
Figure 9. Variation of mission time and water consumption with different heights and flow rates of coolant
图 10 不同高度与预冷剂流量下耗油量与水油消耗量变化
Figure 10. Variation of fuel consumption and payload consumption with different heights and flow rates of coolant
图 11 不同马赫数与预冷剂流量下加速度与推力变化
Figure 11. Variation of accelerated speed and thrust with different Mach numbers and flow rates of coolant
图 12 不同马赫数与预冷剂流量下机动时间与水油消耗量变化
Figure 12. Variation of mission time and payload consumption with different Mach numbers and flow rates of coolant
图 13 发动机内部温度随预冷剂流量的变化
Figure 13. Temperature of aero-engine interior with different flow rates of coolant
图 14 理论升限与马赫数随预冷剂流量的变化
Figure 14. Variation of theoretical ceiling and Mach number with different flow rates of coolant
图 15 不同飞行状态下供给推力比与叶片表面温度的变化
Figure 15. Variation of supply ratio of thrust and blade surface temperature under different flight conditions
表 1 设计点循环参数取值
Table 1. Cycle parameters of design point
循环参数 H=0 km, Ma=0 H=11 km, Ma=0.9 进口流量/（kg/s） 125 35 压气机压比 12 6.85 涡轮前温度/K 1 850 1 200 最高涡轮前温度/K 2 000 2 000 燃烧效率 0.999 0.975 压气机绝热效率 0.8 0.79 涡轮绝热效率 0.9 0.88 涡轮导叶引气比例/% 9 9 涡轮动叶引气比例/% 6 6 飞机引气比例/% 1 1 
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表 2 设计点模型验证
Table 2. Validation of design points
设计点 性能参数 模型 GasTurb 误差/% 1 推力/kN 126.69 126.96 −0.21 耗油率/（kg/（N·h）） 0.1117 0.1096 1.92 2 推力/kN 18.05 18.11 −0.33 耗油率/（kg/（N·h）） 0.1196 0.1162 2.93
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表 3 发动机非设计点
Table 3. Off-design points of the aero-engine
非设计点 飞机状态 H/km Ma 发动机工作状态 1 起飞滑跑 0 0.1 加力最大状态 2 低空爬升 5 0.5 最大状态 3 亚声速巡航 11 0.9 巡航状态 4 超声速巡航 11 1.5 巡航状态 5 近实用升限 20 2.5 加力最大状态
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