Review on active thermal protection and its heat transfer for airbreathing hypersonic vehicles

Experimental equilibrium temperatures and skin friction at the surface of a flat plate cooled by two-dimensional, tangential slot injection are presented for a Mach 6 mainstream. Effects of slot height, slot mass-flow rate, injection gas temperature, and heat conduction to the slot are investigated. Experimental skin friction and effectiveness data in general agree well with predictions from a finite-difference method. Film cooling at Mach 6 was found significantly more effective than at lower speeds, with large reductions in friction drag measured downstream of the slot These results indicate the necessity for a new evaluation of film-cooling systems for hypersonic vehicles.

Film-Cooling Effectiveness and Skin Friction in Hypersonic Turbulent Flow

The current work is an experimental investigation of the dependence of film cooling effectiveness on the injection Mach number, velocity, and mass flux. The freestream Mach number is 2.4, and the injection Mach numbers range from 1.2 to 2.2 for both air and helium injection. The adiabatic wall temperature is measured directly. The injection velocity and mass flux are varied by changing the total temperature and Mach number while maintaining matched pressure conditions between the injected flow and that of the freestream. The total temperature of the freestream is 295 K, and for the injection it ranges from 215-390 K. The results indicate an increase in film cooling effectiveness as the injection rate is increased. With the exception of heated helium runs, larger injection Mach numbers slightly increase the effective cooling length per mass injection rate. The results for helium injection indicate an increase in effectiveness as compared to that for air injection. Heated injection, with the injectant to freestream velocity ratios greater than 1, exhibit a rise in wall temperature downstream of the slot resulting in effectiveness values greater than 1. The experimental results are also compared with earlier studies in the literature.

Influence of injectant Mach number and temperature on supersonic film cooling

The use of film cooling for protecting a surface exposed to high temperature air at hypersonic speeds is investigated experimentally. The tests were conducted in a Mach 6 contoured axisymmetric nozzle with a streamlined centerbody. The Reynolds number in the test section was in the range of 1-3.6 X 10 6/in.; and a wall to freestream temperature ratio of 0.635. Heat-transfer distributions downstream of the slot were obtained for various mass flow rates and the effect of injection on the velocity, temperature and Mach number profiles was studied. Correlations for the cooling lengths with the blowing rate parameters X = pjUj/peUe for the various coolants-—air, helium, hydrogen, and argon were obtained. Correlations for the heattransfer rates in the form (1 — q/go.») an<l for the effectiveness [e = (Taw. — T0oo)/(T0j — T0oo)] of the adiabatic wall temperatures were obtained as a function of X and x/S (distance from the slot/slot width) for each of the injected coolant gases. The adiabatic wall temperatures are not a basic result of the measurements, but have been inferred from the heat transfer measured, using a local flat plate equation for the same wall temperature. Skin-friction coefficients were also obtained from the measured profiles. Finally the results of this study are compared with available results in the literature.

An Experimental Investigation of Turbulent Slot Injection at Mach 6

Research progress on active thermal protection for hypersonic vehicles

The effect of injecting a secondary gas stream through a rear ward-facing slot in a direction parallel to a wall into a supersonic air flow has been examined in a small Mach 3 wind tunnel. The experiments were carried out for three step heights, 0.121, 0.180, and 0.239 in. with slot openings of 0.064, 0.123, and 0.182 in., respectively. Air was used as the secondary fluid with these three different slot openings. The total temperature of the main flow was equal approximately to the room temperature, and the temperature of the wall at the point of injection varied between 418° and 655°R as determined by the temperature of the secondary air. The ratio M of the mass velocity in the secondary stream to that of the mainstream was in the range of 0 to 0.408. Helium also was injected through a slot opening of 0.063 in. for two different injection parameters M = 0.01 and 0.02 and a temperature of the wall at the point of injection ranging between 562.8° and 663.7°R. Schlieren photographs are presented and adiabatic wall temperature measurements are described.

A. Haji-Sheikh

Papers

Film cooling with air and helium injection through a rearward-facingslot into a supersonic air flow.