超声速颌下乘波进气道一体化设计

Integration Design of Supersonic Waverider Chin Inlet

关键词:  冲压发动机  超声速进气道  计算流体动力学  风洞试验

Key words:  Ramjet  Supersonic inlet  Computation fluid dynamics  Wind tunnel test

摘要:

为实现冲压发动机进气道/飞行器前体融合化设计，采用轴对称弯曲激波压缩基准流场和等熵压缩基准流场的锥导乘波设计方法，设计出工作范围Ma=2.5～4.5的超声速颌下乘波进气道方案，利用三维流场数值模拟获得了进气道基本性能，并对等熵压缩颌下乘波进气道进行了风洞吹风试验，验证了进气道的性能特性。研究结果表明:(1)设计的颌下进气道可以在Ma=2.5～4.5工作，在设计点Ma=4.0实现前体乘波，并具有大捕获流量、高压缩特性及高升阻比的优点；(2)进气道总压恢复系数和流量系数随着攻角增大而提高，在Ma=4.0，6°攻角状态，该进气道总压恢复系数可以达到0.47，流量系数达到1.20；(3)在更大攻角下由于捕获流量大幅增大，颌下进气道会出现不起动现象，但流场结构和性能稳定。

Abstract:

For ramjet inlet-forebody integration， on the basis flow field of axisymmetric curved shock compression and axisymmetric isentropic compression， an integration waverider design of supersonic chin inlet was presented.The chin inlet has a working range from Mach number 2.5 to 4.5. Inlet performance has gained by 3D CFD simulation， and wind tunnel test has developed on the isentropic-compression-waverider chin inlet. The results showed that:(1)The chin inlet can be worked from Mach 2.5 to 4.5， the forebody shock realized waverider on design Mach number 4.0.The principle advantages of waverider chin inlet are the higher compression efficiency and higher capture massflow ratio and higher lift-to-drag ratio. (2)The total pressure recovery and mass flow ratio of the chin inlet were rised with the increasing of incidence angle. The total pressure recovery reached 0.47 and massflow ratio reached 1.20 on condition of Mach number 4.0 at incidence angle 6°. (3)The chin inlet has unstarted at higher incidence angle because of increasing capture massflow， the flow field structure was still steady and performance was stable.