二元混压式进气道压缩段性能快速估算及优化

Rapid Performance Estimation and Optimization for Compression Section of 2D Mixed Compression Inlet

冲压动力导弹的进气道性能决定了导弹动力系统的优劣,因此在冲压动力导弹设计初期,有必要对二元混压式进气道超声速压缩段性能进行快速估算及优化。针对该问题,提出将进气道外形进行参数化建模,建立基于特征线及边界层理论的进气道性能快速估算方法,并通过激波边界层干扰分离指标变量以及喉道流动参数二次修正,提高对进气道性能估算的精度。估算结果与CFD计算结果的对比,表明了该方法可以对设计状态下二元混压式进气道超声速压缩段阻力系数、平均总压恢复系数及流量系数进行具有较高精度的快速估算,最大误差不超过1.5%。此外该方法与遗传优化算法结合,对进气道超声速压缩段外形设计参数进行快速优化,使进气道压缩段阻力系数下降了13.8%,表明该结合方法可在冲压动力导弹设计初期阶段提高二元混压式进气道的性能。

The performance of inlet plays a key role in the propulsive system of ramjet-powered missile. Therefore, the rapid estimation and optimization of the performance of the 2D mixed compression inlet are necessary in the preliminary design of ramjet-powered missile. In order to solve this problem, a parametric model of the supersonic compression section of 2D mixed compression inlet was used. Based on the method of characteristics and boundary layer theory, a method which can rapidly estimate the drag coefficient, average total pressure recovery coefficient and mass flow coefficient of the supersonic compression section of inlet under design condition was put forward. Meantime, to improve the accuracy of the estimation method, a correction which used shock wave boundary layer interference index variable and flow parameters at throat was adopted. In contrast to Computational Fluid Dynamics, the results using rapid estimation method have high accuracy and high efficien- cy with maximum error less than 1.5%. Besides, the supersonic compression section of 2D mixed compression inlet could be optimized by using the rapid estimation method and the genetic algorithm to reduce the drag coefficient by 13.8%. It showed combination can improve the performance of the 2D mixed compression inlet in the preliminary design of ramjet-powered missile.