Serpentine Inlet Performance Enhancement Using Vortex Generator Based Flow Control

In order to provide the line-of-sight blockage of the engine face for an advanced Uninhabited Combat Air Vehicle（UCAV）, a highly curved serpentine inlet is proposed and experimentally studied. Based on the static pressure distribution measurement along the wall, the flow separation is found at the top wall of the second S duct for the baseline inlet design, which yields a high flow distortion at the exit plane. To improve the flow uniformity, a single array of vortex generators （VGs） is employed within the inlet. In this experimental study, the effects of mass flow ratio, free stream Mach number, angle of attack and yaw on the performance of a serpentine inlet instrumented with VGs are obtained. Results indicate： （1） Compared with the baseline serpentine design without flow control, the application of the VGs promotes the mixing of core flow and the low momentum flow in the boundary layer and thus prevents the flow separation. Under the design condition, the exit flow distortion （-↑△σ0） decreases from 11.7% to 2.3% by using the VGs. （2） With the descent of the free stream Mach number the total pressure loss decreases. However, the circular total pressure distortion increases. When the angle of attack rises from - 4° to 8°, the total pressure recovery and the circular total pressure distortion both go down. In addition, with the increase of yaw the total pressure recovery is fairly constant, while the circular total pressure distortion ascends gradually. （3） When Mao = 0.6-0.8, a = -4°-8° and β = 0°-6°, the total pressure recovery varies between 0.936 and 0. 961, the circular total pressure distortion coefficient varies between 1.4 % and 5.4 % and the synthesis distortion coefficient has a ranges from 3.8 % to 7.0 %. The experimental results confirm the excellent performance of the newly designed serpentine inlet incorporating VGs.

Numerical Analysis and Experimental Validation of a Submerged Inlet on the Plane Surface

In order to get a deep insight of a submerged inlet on the plane surface, the integrated flow field of the inlet and fuselage has been numerically studied. The investigation is mainly focused on the formation of the total pressure distribution at the exit of the inlet, the structure of the inner flow and the effects of the boundary layer along the fuselage on the performance of the inlet. Moreover, in comparison with the experimental data at different angles of attack, yaws and mass flow ratios, the reliabilities of the computational fluid dynamics（CFD） studied are verified. Results indicate： （1） the CFD results agree well with the experiment results and the relative errors of the total pressure coefficient is less than 1% ; （2） at the inlet＇s exit, the contour of total pressure obtained by CFD is similar to the experiment result except the contour in the low total pressure zone in CFD is slightly larger; （3） the secondary flow at the cross section behave as two counter-rotating vortices. Along the flow direction, the fields influenced by the vortex pair transport downstream and expand to the whole section at the exit; （4） the total pressure loss at the exit of the submerged inlet can be divided into external loss and internal loss. Usually, the external loss is greater than the internal loss, and both decrease with the augment of the Mach number at the exit. In addition, when the angle of attack ranges from -2° to 8°, the total pressure coefficient ascends gradually, due to the reduction of the external loss caused by the less boundary layer flow captured and the invisible change of the internal loss.