高马赫数超声压气机转子叶型优化设计

High Mach Number Supersonic Compressor Rotor Blade Airfoil Optimization Design

为进一步提高压气机叶尖轮缘速度和增压比,将唯一进气角原理和数值最优化技术用于叶型设计,获得两个高马赫数、高压比、低损失的"S"形超声压气机叶型。首先根据压气机流动机理,提出超声压气机叶栅的性能指标;然后通过吸力面叠加厚度的方式生成初始叶型,保证叶栅的来流马赫数和唯一进气角;最后采用基于修改量的叶型参数化方法,以给定总压比为约束条件,以总压损失系数最小为目标对初始叶型优化。设计结果表明:在设计点,叶栅1和叶栅2的总压损失系数分别为0.119和0.158;在高来流马赫数条件下,超声叶栅需采用大稠度设计才能实现多道斜激波加一道正激波增压;在叶型吸力面前端构造一个斜坡也可增加叶栅通道内的斜激波数量;平直的吸力面后段有利于削弱激波对附面层干扰,将平直吸力面后段与钝尾缘(或翘尾缘)相结合可有效抑制附面层分离,减小尾迹区。

To achieve higher blade tip speed and single stage pressure-ratio in compressor,the unique incidence principle and numerical optimization technology are used to design blade airfoil. And two ＂S＂ shape supersonic cascades with high inlet Mach number,high pressure-ratio and high efficiency were obtained. Firstly,the design target is proposed,according to compressor flow mechanism. Then,the initial airfoils,whose inlet Mach number and flow angle equal with design value,were generated through overlaying thickness to suction side curve. At last,genetic algorithm and the blade airfoil parameterization method based on modifying quantity were used to optimize the initial airfoil. The optimization object is the minimal total pressure loss coefficient. The restriction condition is that total pressure-ratio equals to design value. The results show that,the total pressure loss coefficient is 0.119 at designing point for cascade 1, and 0.158 for cascade 2. Supersonic cascade may achieve multi-shock pressurization only if solidity was high,in the case of high inlet Mach number. The number of oblique shock increases,if a slope existed in the front of suction side. The straight suction rear segment can inhibit the shock wave/boundary layer interaction,and can even weaken separation when it is combined with blunt trailing edge（or the warped trailing edge）.