Hypersonic Boundary-Layer Transition Experiments in a Mach-6 Quiet Tunnel

Abstract

The Boeing/AFOSR Mach-6 Quiet Tunnel has achieved quiet flow to stagnation pressures of 146 psia, and intermittently quiet flow between 146 and 169 psia. In an attempt to measure natural transition under quiet flow, a 3-m-circular-arc compression cone was tested with a nearly sharp nosetip. Using temperature-sensitive paint, hot streaks were observed to develop near the rear of the cone at high pressures under quiet flow. The streaks do not appear under noisy flow. The cause of the hot streaks remains unknown, though they may be instabilities or artifacts of nonlinear breakdown. Under quiet flow, the cone boundary layer remained laminar up to N factors of at least 15 and possibly as high as 19. Transition occurred at N = 9 under noisy flow. It is unknown why laminar flow persisted to such high N factors. As part of an investigation of crossflow vortices, a 7 • half-angle cone was tested at 6 • angle of attack with temperature-sensitive paint finishes of varying roughness. The roughness of the paint finish was observed to have an effect on crossflow vortices, in some cases inducing transition under noisy flow. Heat-transfer measurements were made at the stagnation point of a hemisphere to observe the effect of freestream noise; no effect was evident.

Numerical Investigation of Laminar-Turbulent Transition in a Cone Boundary Layer at Mach 6

“…These streaks look similar to the streamwise streaks observed in the Purdue experiments using temperature sensitive paint (Berridge et al, 2010). The qualitative similarity of the "hot" streaks observed in our simulations and the streamwise streaks …”

“…Figure 5.14: Time-averaged skin friction from CFUND 1 (top), CFUND 2 (center) and CFUND 3 (bottom). The streamwise alligned "hot" streaks observed in our simulations looks qualitatively similar to the streamwise streaks observed in the Purdue experiments using temperature sensitive paints for a flared cone (Berridge et al, 2010). The time-and azimuthally-averaged streamwise velocity profiles plotted in wall coordinates can provide more insight on how far transition has progressed.…”

“…Figure 5.13: Time-averaged heat transfer (dT /dy) at the wall from CFUND 1 (top), CFUND 2 (center) and CFUND 3 (bottom). The streamwise alligned "hot" streaks observed in our simulations looks qualitatively similar to the streamwise streaks observed in the Purdue experiments using temperature sensitive paints for a flared cone (Berridge et al, 2010). Figure 5.14: Time-averaged skin friction from CFUND 1 (top), CFUND 2 (center) and CFUND 3 (bottom).…”

Numerical Investigation of Transition in a Flared Cone Boundary Layer at Mach 6

“…The flow parameters chosen are those of the experiments by Schneider and coworkers 5,6 in the Boeing/AFOSR Mach 6 quiet-flow Ludwieg Tube at Purdue University. The approach flow has a Mach number of 6 and a unit Reynolds number of 9.276 x 10 6 m −1 .…”

Experimental Study of Second-Mode Instabilities on a 7-Degree Cone at Mach 6

“…It is uncertain how pressure is transmitted to the sensing element through this epoxy, so the active sensing area is unknown. It has been stated earlier that the sensing element size was 1 × 1.6 mm, [18][19][20][21][22] but further communication with the sensor manufacturer revealed that to be an error.…”