作者在试验中发现,超音速飞机轴对称进气道在迎角下,中心锥体上背风侧的斜激波倾角增大,斜激波离开唇部,而迎风侧的斜激波倾角减小,它进入外罩唇口的内侧。计算结果与试验观察到的完全吻合。这同苏联教科书中和国内沿用的观点恰好相反。 The authors found in experiments that the tilt angle of the lean shock on the leeward side of the central cone increases at the angle of attack of the axisymmetric inlet of the supersonic aircraft, the oblique shock leaves the lip, and the dip of the oblique shock on the windward side decreases , It enters the inside of the cover lip. The calculated results are in good agreement with the experimental ones. This is exactly the opposite of the Soviet textbook and domestic use.