For air-breathing high-speed vehicles, inner and outer flows are highly coupled and there are strong interactions between shockwaves and boundary layers. Boundary layers are sensitive to the variation of the wall temperature, so do the aerodynamic characteristics. Consequently, how to predict the wall-temperature effect exactly and improve the prediction accuracy of aerodynamic characteristics are important for the design of high-speed vehicles. In this article, the effect of the inner flow wall temperature on aerodynamic characteristics is investigated through wind-tunnel measurements and numerical simulations. Results show that in conventional hypersonic wind tunnels, the inner wall temperature of the high-speed vehicle model increases continuously with the wind-tunnel running time, yielding significant variations of the pitching moment and wall pressure. The maximum increment of the pitching moment requires an increase of the rudder angle by 2°. In addition, numerical simulation results indicate that the thickness of boundary layer in the internal flow increases with the rise of wall temperature. Moreover, the equivalent area of high-speed flow channel decreases so that the flow is compressed. The change of velocity profiles in boundary layers results in the forward moving of shockwaves in the internal flow channel, which causes the change of pitching moment of high-speed vehicles.

To induce the near-wall low-momentum fluids moving laterally, an aerodynamic design method of hypersonic forebody/compression surface with the controlled lateral surface pressure is proposed. The basic principle is that the lateral pressure distribution on one section after the external conical flow field is prescribed, and the spatial coordinates of the section can be derived inversely through the coordinate transform, then the forebody/compression surface can be obtained by the stream tracing method. The numerical results demonstrate that the lateral pressure gradient dominates the motion of the near-wall low-momentum fluids on the forebody/compression surface. For the conventional forebody/compression surface, it has strong lateral pressure gradient on the 1^st stage of the forebody, which can induce a lateral flow with a deflection angle below 3° at the design point (Ma = 7.0 and H = 28 km). However, it almost has no lateral pressure gradient on the subsequent compression surfaces, and the lateral flow is also fairly weak. The controlled lateral pressure distribution forebody can intensify the lateral pressure gradient about 7 times within the sector-angle ranging from 0° to 40°, the deflection angle increases about 5° on the 1^st stage of the forebody, and the lateral pressure gradient increases significantly with the deflection angle increased over 7° on the 2^st and 3^rd stages of the forebody. Consequently, the boundary layer thickness decreases about 20%, and the total-pressure recovery coefficient in the sector region of the inlet increases about 1.56%.

A flow control method using multiple micro-vane vortex generators is proposed to improve the fullness of boundary layer profiles over the forebody ramp of hypersonic vehicles and to reduce the potential risk of flow separation in the inlet. The flowfield characteristics and mixing mechanism at an incoming Mach number of 7.0 are numerically analyzed. The effect of the installation angle of vortex generators on the flowfield is studied as well. Results show that micro-vane vortex generators can generate a local large-sideslip-angle and low-pressure flow in the near-wall region. At the lateral sides of mocro vanes, glancing shocks and expansion waves are generated, which induce the lateral mixing of low-momentum fluid with the mainstream. When installation angles are positive, with the increase of installation angle, the mixing effect gets stronger but the pressure loss increases. Micro vanes with negative installation angles lead to the strongest mixing effect, however, they also bring significant total pressure loss. Compared with the one without flow control, the shape factor of the boundary layer gets smaller by using vortex generators, among which those with an installation angle of 15 degrees lead to the smallest shape factor and consequently the fullest boundary layer profile, which has better anti-separation capability under reverse pressure gradients.