|
Abstract:
The flow through a terminating shock-wave and subsequent subsonic diffuser typically found in supersonic inlets has been simulated using a small-scale wind tunnel. Experiments have been conducted at an inflow Mach number of 1.4 using a dual channel working section to produce a steady near-normal shock-wave. The setup was designed so that the location of the shock-wave could be varied relative to the diffuser. As the near-normal shock-wave was moved downstream and into the diffuser, an increasingly distorted, three-dimensional and separated flow was observed. Compared to the interaction of a normal shock-wave in a constant area duct, the addition of the diffuser resulted in more prominent corner interactions. A variety of control philosophies were then employed to improve pressure recovery and reduce flow distortion. When control was employed along the centre-line, improvements in this region were masked by enlarged corner separations, and even asymmetric solutions in some configurations. On the other hand, when control was applied in the corners, although separation from the corners was removed this was replaced by a more two-dimensional floor separation. Only when control of both the centre-line and corners were addressed simultaneously was an overall improvement in the pressure recovery and distortion obtained. Using an appropriate control configuration (VGs on the centre-line and bleed in the corners) a highly separated flowfield is drastically improved, leading to the conclusion that there is significant potential for the use of nonbleed supersonic control when implemented in the correct manner.
| Limitations: |
 APPROVED FOR PUBLIC RELEASE |
| Description: |
Final rept. 1 Oct 2008-31 Mar 2011 |
| Pages: |
34 |
| Report Date: |
May 2011 |
| Contract Number: |
FA8655-08-1-3091 FA86550813091 |
| Report Number: |
A663345 |
|
|
|
Email This Abstract
