提高超声速压气机级喘振裕度方法研究

Study of Improving Surge Margin for aSupersonic Compressor Stage

关键词:  压气机  流场计算  喘振裕度  叶片前掠  叶尖稠度  几何进口角

Key words:  Compressor  Flow field calculation  Surge margin  Forward sweep  Blade tip solidity  Geometric inlet angle

摘要:

在超声速压气机气动设计时，为实现设计点高性能和宽喘振裕度，提出采用优化方法以设计点性能为目标进行叶片设计，通过转/静子叶片几何手动修改提高压气机喘振裕度。以NASA Rotor 37为原型，应用此方法进行更高性能超声速压气机转子气动设计，并匹配静子，构成压气机级。结果表明：超声速压气机转子通道激波推出和静子大攻角分离是失速发生的主要原因，因此分别进行转子叶片前掠设计、改变叶尖稠度，以控制激波位置，单转子喘振裕度可从约7%提高到18%以上；静子上采用前掠、切向弯、修改叶片数及几何进口角等措施，最终将此压气机级的喘振裕度由约18%提高到30%以上。

Abstract:

In order to achieve high performance at design point and wide surge margin in the aerodynamic design of a supersonic compressor, optimization method was adopted for the blade design at design point, and then the geometry of rotor/stator blade was modified manually to improve the compressor surge margin. This method was applied to the aerodynamic design of the higher performance supersonic compressor rotor and the matching stator to form the compressor stage based on NASA Rotor 37 design parameters. The results show that the supersonic compressor stall occurs because rotor shock is pushed out passage at the tip, or stator boundary layer separates as a result of large incidence. The surge margin of the single rotor improved from 7% to above 18% if measures， such as forward sweep and increasing blade tip solidity， are taken to control the shock position in the passage. Beneficial methods such as forward sweep tangential lean, modification of blade number and geometric inlet angle are also adopted for the stator. The predicted surge margin of the compressor stage finally improved from 18% to above 30%.