尖侧缘机身布局的俯仰力矩特性及扰流板控制

针对尖侧缘机身布局在大迎角下存在的正俯仰力矩（抬头力矩）问题，通过风洞试验，首先研究了俯仰力矩的迎角分区特性及流动演化规律：线性增长区（迎角为0°~15°），俯仰力矩线性增加，全机从附着流到形成进气道前缘涡和机翼涡；非线性增长区（迎角为17.5°~32.5°），俯仰力矩非线性增加，机头涡出现，机头涡和进气道前缘涡逐渐增强，机翼涡增强后破裂；衰减区（迎角为35°~65°），俯仰力矩逐渐减小，机头涡增强后破裂，进气道前缘涡破裂发展，机翼涡完全破裂。其次，发现了机身前体是产生正俯仰力矩的主要来源，机头涡是导致大迎角下正俯仰力矩的主控流动。当迎角为40°时，前体各截面正俯仰力矩在进气道前缘处达到最大，主要是由于该处机头涡诱导产生了较强的法向力。最后，提出了大迎角机身扰流板控制技术，产生了较好的控制效果。当迎角为40°时，扰流板可使正俯仰力矩减少62%，其原因是扰流板降低了机头涡涡量及其诱导产生的法向力，减少了机身前体对正俯仰力矩的贡献。该控制技术的缺点是扰流板会带来一些升力损失和附加阻力。基于尖侧缘机身参考宽度的雷诺数为2.59×10⁵。

Pitching moment character and its control using forebody spoiler over a configuration with chined fuselage

Wind tunnel experiments are conducted to understand the positive pitching moment of a configuration with chined forebody at high angles of attack. Firstly, the trend of pitching moments with angles of attack can be divided into zones of linear growth (0°-15° angle of attack), nonlinear growth (17.5°-32.5° angle of attack) and decay (35°-65° angle of attack). In the zone of linear growth, the pitching moment increases linearly, and the attached flowis observedat first and the inlet leading-edge vortex and wing vortex begin to form as the angles of attack increase. In the nonlinear growth zone, the pitching moment increases nonlinearly, and the nose vortex appears and becomes incrementally strong. Both the inlet leading-edge vortex and wing vortex become more intensive and then begin to burst. In the decay zone, the pitching moment gradually decreases and the nose vortex continues to strengthen and break down at a sufficient high angle of attack. The inlet leading-edge vortex exhibits a decay trend while the wing vortex completely bursts. Secondly, it is observed that the chined forebody is the main contribution component to the positive pitching moment. And the nose vortex over the chine forebody plays a dominant role in generating the positive pitching moment. At 40° angle of attack, due to the strong suction of the nose vortex, the sectional pitching moments reach the maximum at the leading edge of the inlet. In the end, the technique using a forebody spoiler is developed to control the positive pitching moments effectively at high angles of attack. With the forebody spoiler on, the positive pitching moment can be reduced by 62% at 40° angle of attack because by deploying the forebody spoiler, the vorticity of nose vortex is decreased and results in a reduced normal force and positive pitching moment provided by the chined forebody. The downside of the control technique is that the forebody spoiler may lead to some losses in lift and additional drag. The Reynolds number in this paper is 2.59×10⁵ based on the characteristic width of the chined fuselage.