Shock wave/boundary layer interaction (SBLWI) is a common coupling phenomenon that occurs in hypersonic vehicles, and it adversely affects the flow fields inside and outside the vehicle. Currently, the microjet active flow control approach has been extensively investigated experimentally. Herein, the influence of the microjet hole type on SBLWI control is investigated and the flow field structures of the microjet in the near field are obtained. Additionally, the control effects of isosceles triangular microjet holes with different vertex angles are compared with those of circular and square microjet holes. The numerical results show that the control effect of the microjet on SBLWI gradually improves with increasing triangular vertex angle, and the control effect of the triangular microjet hole with a large vertex angle is better than that of circular and square microjet holes. In all working conditions, the triangular microjet hole with a 90° vertex angle that is placed along the flow direction affords the best control effect, and it can reduce the volume of the separation zone at the corner by 74.96%. Simultaneously, the addition of the microjet causes some loss of the total pressure, but the value is small.

冲击波/边界层相互作用（SBLWI）是高超声速飞行器中常见的耦合现象，对飞行器内外流场产生不利影响。目前，微射流主动流动控制方法已在实验中得到广泛研究。本文研究了微射流孔类型对SBLWI控制的影响，得到了微射流近场流场结构。此外，比较了不同顶角的等腰三角形微喷孔与圆形和方形微喷孔的控制效果。数值结果表明，微射流对SBLWI的控制效果随着三角顶角的增大而逐渐提高，大顶角三角形微喷孔的控制效果优于圆形和方形微喷孔。在所有工况下，沿流向布置的顶角为90°的三角形微射流孔控制效果最佳，可将转角分离区体积减少74.96%。同时，微射流的加入对总压有一定的损失，但数值较小。