大攻角(大侧滑角)下超音速后置旁侧进气道试验研究

AN EXPERIMENTAL INVESTIGATION ON AFT BYPASS SUPERSONIC INLET PERFORMANCE AT HIGH ANGLE OF ATTACK AND YAW

关键词:  超音速进气道  腹部进气道  风洞试验  迎角  性能

Key words:  Supersonic inlet  Ventral inlet  wind-tunnel test  Angle of attack  Performance

摘要:

简要介绍后置旁侧进气道模型风洞吹风试验结果,特别对大攻角大侧滑角下进气道工作状态进行了详细的讨论。 本试验共设计了A、B两套模型,A模型为半锥进口,采用双下腹部后置旁侧布局;B模型为轴对称进口,采用十字型后置旁侧布局。进气道从气动上采用了单锥混合式、超额定工作设计。试验马赫数M-H为2.0,2.5;攻角为-14°,-12°,-10°,0°,10°,12°,13°;测滑角为0°,10°,12°,14°,15°。 试验结果表明A、B两模型在大攻角、大侧滑角条件下能稳定工作。在进气道拐弯突扩几股气流掺混的条件下测量总压的方案是可行的,其测出的进气道总压恢复系数是令人满意的。

Abstract:

The experimental result of aft bypass supersonic inlet model in wind tunnel test is presented in this paper. And inlet performance, specially at high angle of attack and yaw is discussed in detail.There are two models for wind tunnel test. Model A is twin half-cone inlets aft-installed under fuselage. Model B is four symmetrical inlets insta-lled around fuselage. The inlets with single cone and mixed external/internal compression were tested at flight Mach number 2.0-2.5, at which conical shock wave enters inlet’s entrance) attack angle ranging from - 14° to 13° and yaw angle ranging from 0° to 15°.Experimental results show that both models can operate stably at high angles of attack and yaw. The total pressure measurement under the mixing condition at dump chamber entrance is feasible and the pressure recovery measured is satisfactory.