吸气式高超声速飞行器冷流试验设计及验证

对于吸气式飞行器而言,地面冷流试验是检验其进气道性能及气动特性的一项重要手段.以二元混压式进气道、机体/推进系统耦合为基本特征,设计了采用超燃冲压发动机为推进系统的内外流一体化巡航飞行器,针对其高超声速特性开展了冷流风洞试验,来流速度范围Ma=5.0~7.0,攻角范围α=-4°~8°.测压试验结果表明,随着来流马赫数的增大,进气道的总压恢复系数下降;而流量系数先上升,在设计点达到最大值;在一定攻角范围内,进气道的总压恢复系数和流量系数提高,但当攻角增大至巡航攻角时,随着攻角的增大,进气道的总压恢复系数和流量系数逐渐下降.测力试验验证了数值算法的有效性,除轴向力系数以外,其余气动特性系数的发展规律及数值基本吻合,可通过修正试验值的方式外推出飞行器的气动特性数据.

Design and validation of cold-flow test for air-breathing hypersonic vehicle

Cold-flow test is an important method for air-breathing vehicle to validate the inlet capability and aerodynamic characteristics. An airframe/propulsion integrative cruise vehicle with scramjet egine was designed; two-dimensional mixed-compression inlet and airframe/propulsion system coupling was used. The cold-flow wind tunnel test of the model was finished in hypersonic wind tunnel. Experimental data was measured when free Mach number was 5.0,6.0,7.0, and angle of attack was between -4° and 8°. The pressure performance test results indicate that, with the increase of the free stream Mach number, the total pressure recovery coefficient decreases, while the mass flow ratio increases firstly, and the maximum is obtained at design point; with the range of the angle of attack to a certainty, the total pressure recovery coefficient and the flow ratio increases. But after the cruise angle of attack, with the increase of the angle of attack, the total pressure recovery coefficient and the mass flow ratio decreases. The aerodynamics force performance wind tunnel test results validate the efficiency of the numerical simulation method, which shows a consistent trend with the experimental results, except the axial force coefficient. The aerodynamics characteristics of air-breathing hypersonic cruise vehicle could be obtained by revising the experimental results.