钛合金叶片燃烧后冷却过程的三维热流耦合数值模拟

Cooling Process of Titanium Alloy Blades After Combustion Using Three Dimensional Heat Flow Coupling Numerical Simulation

为了理解航空发动机钛火发生后叶片的冷却过程,采用有限元方法结合ROTOR37转子模型,分别对550℃阻燃钛合金(TF550钛合金)和600℃高温钛合金(TA29钛合金)燃烧后压气机通道内温度场、流场进行数值模拟研究。结果表明:相对马赫数对于压气机通道内叶片散热有一定的影响,其中叶尖燃烧区域在0.7~1的低马赫数区域散热能力最佳;相对于前缘燃烧区域,叶尖燃烧区域的冷却过程更为复杂,且冷却速率比前缘燃烧区域低一个数量级;在叶尖燃烧区域内,TF550钛合金和TA29钛合金的冷却温度差异比较显著,在1000~2500K温度区间内的差别最大,前者比后者低100K以上,在300~500K温度区间内前者比后者低30K以内;叶尖燃烧区域流场的温度畸变会增加喘振的剧烈程度,设计叶片时应充分考虑燃烧对喘振裕度的影响。

In order to understand the cooling procedure of aviation engine blade after titanium combustion,the finite element method was used to simulate the temperature and fluid field of ROTOR37 model after combustion occurred with 550℃fire proof titanium alloy(TF550 titanium alloy)and 600℃high temperature titanium alloy(TA29 titanium alloy),respectively.The results show that the relative mach number influences the cooling procedure of blade,the cooling performance at the area of mach number about 0.7-1 is much higher than other area;compared with the leading edge,the cooling process of the tip is more complex,and the cooling rate is an order of magnitude lower than that of the leading edge.The difference of cooling temperature between TF550 titanium alloy and TA29 titanium alloy at tip combustion area is quite observable;and the maximum value occurs within the scope of 1000-2500K;the former is more than 100K lower than the latter,the value is reduced into 30K within the scope of 300-500K.The temperature distortion of the flow field would increase the intensity of the surge,the effect of combustion on the surge margin should be fully considered during the design of the blade.