基于陶瓷基复合材料铠甲的涡轮导叶热防护研究

Thermal protection of turbine guide vane with ceramic matrix composite armor

关键词:  航空发动机  涡轮导向叶片  前缘冷却  陶瓷基复合材料  数值仿真

Key words:  Aero-engine  Turbine guide vane  Leading edge cooling  Ceramic matrix composites  Numerical simulation

摘要:

为了加速推进我国高推重比航空发动机研制进程，突破已有高温合金+气膜冷却架构对涡轮叶片承温能力的限制，以及解决陶瓷基复合材料（CMC）涡轮叶片制造难度过大的问题，本文提出了一种在涡轮导向叶片前缘镶嵌CMC铠甲的热防护方案，并通过多工况下的数值仿真，研究了其与典型气膜冷却相比在流动和传热上的差异。相较于后者，在相同的冷气压比1.03下，采用该CMC铠甲结构可以使得冷气质量流量显著降低58%；同时可以利用CMC铠甲与冷气的共同作用，使前缘区的叶片基体冷却效率由0.50显著提升至0.85；通过研究CMC材料导热系数对其自身温度分布影响，探讨了避免铠甲失效的极限燃气温度，相对保守的结果为2017K。有望通过对CMC材料和铠甲结构的进一步性能优化来满足第五代先进航空发动机的需求。

Abstract:

To improve the thrust-to-weight ratio of aero-engines， it is crucial to overcome the current limitations of superalloy and film cooling on the temperature capacity of turbine blades， as well as the challenges associated with the difficulty of ceramic matrix composite （CMC） turbine blade manufacturing. Therefore， novel scheme was proposed in this paper， which installed CMC armors as vane leading edge thermal protection. And numerical simulations with wide operating conditions were applied to study the differences in flow and heat transfer compared with the typical film cooling scheme. At the same pressure ratio of 1.03， the coolant mass flow rate can be greatly reduced by 58% due to the combined effect with the CMC armors. On the other hand， the cooling efficiency also increased from 0.50 to 0.85 significantly， which showed the great effect of CMC armor protection. The temperature margin of CMC armor erosion failure was studied forward by variation of its thermal conductivity. The conservative result of the allowable gas temperature was 2017K， which could meet the requirements of the 5th-generation aero-engines through further performance optimizations of CMC materials and armor schemes.