Effect of transpiration cooling on boundary layer transition for hypersonic flight

Superseding ablative cooling techniques with active thermal protection could benefit the development of hypersonic flight technology by improving resistance of vehicles to thermal loads, re-usability and shape stability. Transpiration cooling could ensure protection of areas that are most affected by aerothermodynamic heating by injecting a coolant gas through a porous surface and thus shield the vehicle from hot surrounding gas. Problematically, gas effusion can advance boundary layer transition and resulting turbulence would cause excess heating and sabotage the local upstream cooling effect. In addition, turbulence reduces aerodynamic efficiency and influences critical variables of flight system dynamics. Transpiration cooling can only be applied if sufficient boundary layer transition control is achieved simultaneously or its influence on stability can be predicted accurately.
This thesis investigates the effect of transpiration cooling on hypersonic boundary layer transition experimentally by conducting tests in a low enthalpy hypersonic Ludwieg tube at Mach 7 using a 7-degree half-angle cone model. Coolant gas injection is performed at two locations on the model. Firstly, fully porous nose tips with 1 mm nose radius and 58.5 mm axial length are used. Secondly, helium, nitrogen, argon and carbon dioxide are injected through a small segment in the frustum, while a sharp impervious nose tip is employed. The porous frustum injector is 25 mm long and spans 60 degrees azimuth in the axial centre of the cone. Outflow profiles of all porous injectors were characterized using a hot wire, whereby three-dimensional maps at high resolution could be obtained. All model parts were optimized for smoothness. This work utilizes high frequency schlieren imaging, focusing laser differential interferometry and surface instruments to conduct instability measurements.
At a unit Reynolds number of 17.9 mio/m using a 1 mm porous nose tip helium injection was applied from said permeable nose cone segment. Thin film gauges show a delayed rise of the surface heat flux above laminar level, surface pressure measurements indicate a delay in growth and decay of the second mode instability and schlieren imaging shows a delayed onset of turbulence. Stabilization was stronger with 0.04 % helium injection than with 0.02 %. Nitrogen injection at a comparable rate was found to advance instability growth and transition. Helium injection at a blowing ratio of 0.1 % caused early transition, potentially due to boundary layer blow off at the stagnation point.
Employing a sharp, impervious aluminium nose tip, the cone was tested at a unit Reynolds number of 12.2 mio/m with injection from the frustum. At this freestream condition, strongly amplified second mode instabilities are present near the frustum injector segment. It was observed to experience a stronger reduction in second mode frequency with increasing blowing ratio and coolant molecular mass. In contrast, the thickening of the boundary layer increased with decreasing coolant molecular mass. Due to the waveguide nature of the second mode, it was inferred that a change in local speed of sound and boundary layer temperature profile must play a role.
Schlieren images and spectral analysis showed how turbulence and instability mechanisms are displaced from the surface by helium injection. FLDI confirmed how second mode instability oscillations are pushed away from the cone. Reduction in heat flux and attenuation of surface pressure disturbances were thus ascribed to a cooling and displacement effect instead of a true transition delay. Fluid turbulence moves upstream for all coolants as seen with schlieren. Heavier injection gases caused less cooling and less displacement of the boundary layer.
This thesis provides the first experimental demonstration of efficient hypersonic stagnation point cooling combined with a boundary layer transition delay. It furthermore makes experimental contributions to the fundamental understanding of the influence of different gas injections on the transition mechanisms on a hypersonic slender cone.

Keywords: boundary layer transition
hypersonic wind tunnel testing
hypersonics
transpiration cooling
second mode instability
Advisor: McGilvray, M
Wylie, S