高超声速内收缩进气道分步优化设计方法

doi:10.7527/S1000-6893.2015.0095

流体力学与飞行力学

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高超声速内收缩进气道分步优化设计方法

王骥飞¹, 蔡晋生¹, 段焰辉²

1. 西北工业大学航空学院, 西安 710072;
2. 中国空气动力研究与发展中心计算空气动力研究所, 绵阳 621000

收稿日期:2014-12-30 修回日期:2015-04-02 出版日期:2015-12-15 发布日期:2015-04-13
通讯作者: 蔡晋生,Tel.:029-88495381,E-mail:caijsh@nwpu.edu.cn E-mail:caijsh@nwpu.edu.cn
作者简介:王骥飞,男,博士研究生。主要研究方向:计算流体力学,飞行器气动布局设计,E-mail:jifeiwangnwpu@163.com;蔡晋生,男,博士,教授,博士生导师。主要研究方向:计算流体力学,飞行器气动布局设计,流动控制机理与技术。Tel:029-88495381,E-mail:caijsh@nwpu.edu.cn;段焰辉,男,博士后。主要研究方向:计算流体力学,飞行器气动布局设计,E-mail:duanyh2010@gmail.com
基金资助:
省部级项目

Multistage optimization design method of hypersonic inward turning inlet

WANG Jifei¹, CAI Jinsheng¹, DUAN Yanhui²

1. School of Aeronautics, Northwestern Polytechnical University, Xi'an 710072, China;
2. Computational Aerodynamics Research Institute, China Aerodynamics Research and Development Center, Mianyang 621000, China

Received:2014-12-30 Revised:2015-04-02 Online:2015-12-15 Published:2015-04-13
Supported by:
Provincial/Ministerial Level Project

摘要/Abstract

摘要：

提出了基准流场与唇口平面形状分步优化的高超声速内收缩进气道设计方法。基准流场以反射激波不均匀性最小和总压恢复最大进行多目标优化设计,使用结合Tayler-Maccoll方程的有旋特征线方法(MOC)进行流场计算,获得双拐点母线内收缩锥基准流场。进气道唇口形状以沿流线积分(Streamline Integral Method, SIM)获得的进气道无黏阻力最小为目标进行优化设计,获得类椭圆形唇口平面形状。针对优化设计结果进行数值模拟,与传统直母线基准流场相比,双拐点母线基准流场反射激波后流动不均匀性下降40%左右,总压损失减少35%左右,总体性能提升明显。类椭圆唇口进气道在设计点的单位质量流量无黏阻力相较于圆形唇口降低6%,具有良好的压缩特性和气动效率,能够减弱进气系统对飞行器气动性能的不利影响。研究结果表明该方法是一种高效且实用的高超声速内收缩进气道设计方法。

关键词: 高超声速, 内收缩进气道, 优化设计, 流线追踪, Navier-Stokes方程, 数值模拟

Abstract:

Hypersonic inward turning inlet is optimized by a multistage method of separating the base flow and lip shape. Multi-objective optimization is imposed on the base flow with the target of minimizing the flow non-uniformity after the reflected shock and maximizing the total pressure recovery. The flow field has been solved by the method of characteristic (MOC) integrated with the Tayler-Maccoll equations. Base flow optimization obtains a dual-infection point internal cone's element. Inlet lip shape is optimized by the target of minimizing the inviscid drag obtained by streamline integral method (SIM) and a semi-oval two-dimensional lip shape is obtained. Compared with the traditional straight element base flow, dual-infection point internal cone flow's non-uniformity is decreased by about 40% and the total pressure loss is decreased by 35%, and thus the overall performance is enhanced significantly. Semi-oval inlet's inviscid drag per unit mass flow is decreased by 6% compared with the circle lip shape inlet at the design point, and its good compress characteristics and aerodynamic efficiency can partially get over the inlet system's adverse influence on the aerodynamic performance of vehicle. Research proves the design method of hypersonic inward turning inlet proposed in this paper is efficient and functional.

Key words: hypersonic, inward turning inlet, optimization design, streamlines tracing, Navier-Stokes equation, numerical simulation

中图分类号:

V211.48

王骥飞, 蔡晋生, 段焰辉. 高超声速内收缩进气道分步优化设计方法[J]. 航空学报, 2015, 36(12): 3759-3773.

WANG Jifei, CAI Jinsheng, DUAN Yanhui. Multistage optimization design method of hypersonic inward turning inlet[J]. ACTA AERONAUTICAET ASTRONAUTICA SINICA, 2015, 36(12): 3759-3773.

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E-mail：hkxb@buaa.edu.cn

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高超声速内收缩进气道分步优化设计方法

Multistage optimization design method of hypersonic inward turning inlet

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