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首先利用高速风洞对一种与机身保形的双S进气道原始模型进行了研究,结果表明进气道出口截面总压周向畸变指数较大。继而,在低速风洞试验的基础上选择了一种基于涡流发生器的流场控制方案,并在高速风洞中开展了对该进气道高速风洞流场控制试验研究,分别获得了流量特性、速度特性、攻角特性和侧滑角特性规律。研究结果表明:(1)原型方案的高速风洞试验结果说明双S弯进气道第二S弯上壁面产生了气流分离,在横截面二次流的共同作用下,导致该方案出口截面的上方存在一较大的低压区,当Ma0=0.8,α=0°,β=0°时匹配点处总压恢复系数σ为0.958,周向总压畸变指数Δσ0达到11.7%,超过了一般航空发动机的忍受范围。(2)与原型方案的风洞试验结果相比,涡流发生器控制技术能够有效抑制双S弯进气道第二S弯上壁面的气流分离,大幅度降低了该进气道的流场畸变。设计状态下(Ma0=0.8,α=0°,β=0°)总压恢复系数σ为0.953,周向总压畸变指数Δσ0仅有2.3%,综合畸变指数W为4.1%,满足了发动机的使用条件。(3)研究范围内,较低的飞行马赫数使得流场控制方案出口截面的总压恢复系数略有升高,但对周向畸变指数有着不利影响。此外,随着攻角从-4°增加到8°,出口总压恢复系数和周向畸变指数均逐渐降低。而当侧滑角从0°变化到6°时总压恢复系数几乎不变,但大侧滑角给周向畸变指数带来的不利影响较为显著。(4)在飞行马赫数Ma0=0.6～0.85,攻角α=-4°～8°,β=0°～6°的范围内,匹配点处进气道的总压恢复系数在0.936～0.961之间,周向畸变指数在1.4%～5.4%之间,综合畸变指数在3.8%～7.0%之间,表明采用流场控制后的进气道方案已达到实用水平。
Firstly, a high-speed wind tunnel was used to study a twin-S inlet model with shape-preserving fuselage. The results show that the total circumferential pressure at the outlet of the inlet section is larger. Then, based on the low-speed wind tunnel test, a flow field control scheme based on the vortex generator is selected, and the experimental study on the flow field control of the high-speed wind tunnel in the high-speed wind tunnel is carried out. The flow rate Characteristics, speed characteristics, angle of attack characteristics and characteristics of the law of the side slip angle. The results show that: (1) The high-speed wind tunnel test results of the prototype scheme show that the airflow separation occurs on the upper wall of the second S-bend of the dual-S curved inlet duct. Under the combined effect of the secondary flow in the cross-section, There is a large low-pressure zone above it. When Ma0 = 0.8, α = 0 ° and β = 0 °, the total pressure recovery coefficient σ at the matching point is 0.958 and the circumferential total pressure distortion index Δσ0 reaches 11.7% Endurance range of the engine. (2) Compared with the wind tunnel test results of the prototype scheme, the vortex generator control technology can effectively restrain the airflow separation on the upper wall of the second S-bend of the double-S curved inlet and greatly reduce the flow field distortion of the inlet . Under the condition of design (Ma0 = 0.8, α = 0 °, β = 0 °), the total pressure recovery coefficient σ is 0.953, the circumferential total pressure distortion index Δσ0 is only 2.3%, and the comprehensive distortion index W is 4.1% Conditions of Use. (3) In the study area, the lower flight Mach number makes the total pressure recovery coefficient of exit section of flow control scheme slightly increase, but has an adverse effect on the circumferential distortion index. In addition, as the angle of attack increases from -4 ° to 8 °, the total outlet pressure recovery coefficient and the circumferential distortion index gradually decrease. When the side slip angle changes from 0 ° to 6 °, the total pressure recovery coefficient does not change, but the large side slip angle to the circumferential distortion index adverse effects more significant. (4) The total pressure recovery coefficient of intake air at the matching point is in the range of 0.936 ~ 0.961 for the flight Mach number Ma0 = 0.6 ~ 0.85, the angle of attack α = -4 ° ~ 8 °, β = 0 ° ~ 6 °, , The circumferential distortion index is between 1.4% and 5.4%, and the comprehensive distortion index is between 3.8% and 7.0%, which indicates that the inlet scheme with flow field control has reached the practical level.