Experimental studies of shock wave/wall jet interaction in hypersonic flow

Holden, Michael; Rodriguez, Kathleen; Nowak, Robert J.; Olsen, G. C.

doi:10.2514/6.1990-607

Cited by 32 publications

(14 citation statements)

References 3 publications

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“…For impulse test facilities, especially for shock tunnels as used in this study for performing experiments, a transformation of the adiabatic temperatures to the measurable wall heat fluxes is necessary, because the adiabatic wall temperature is not reached during the short testing time. This transformation is similar to the approach given in [11]. In general, the heat flux to the surface can be expressed as follows:…”

Section: General Description Of the Flow Problemmentioning

confidence: 96%

“…For the same blowing and freestream conditions, the attainable cooling length is about an order of magnitude larger for the laminar case than for the turbulent case. To compare the data for the laminar and turbulent cases, that is, to relate them to the same Mach and Reynolds numbers, it has been assumed that the correlations given in [2,11] are also valid for the flow conditions considered in this paper.…”

Section: G Comparison With Data From Literaturementioning

confidence: 99%

“…However, for supersonic flow conditions and turbulent boundary layers, film cooling has been investigated by employing tangential slot injection [11][12][13][14]. One possible application of this is the cooling of scramjet combustion chambers.…”

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confidence: 99%

“…In these experiments, the ratio of the specific cooling gas mass flow to the corresponding specific freestream mass flow was less than the ratio for turbine blade cooling. In [11], a correlation factor has been found based on Goldstein's approach [2] to correlate the results for different blowing parameters.…”

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confidence: 99%

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Experimental and Numerical Study of Cooling Gas Injection in Laminar Supersonic Flow

Heufer

Olivier

2008

AIAA Journal

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In this study, we investigate film cooling on an inclined flat plate in a laminar supersonic flow. The influence of the most relevant parameters, the Mach and Reynolds numbers of the freestream, the blowing ratio, and the blowing geometry, is examined using numerical simulations and in experiments. Importantly, a correlation factor that describes the effect of the most important parameters for a large range of flow conditions has been developed from a simple heat balance model. Unlike the results for turbulent conditions that are typical for cooling turbine blades, interestingly, we have found that considerably less cooling gas mass flows are needed to cool the surface over a large area. In laminar boundary layers no turbulence appears, so that neither the cooling gas strongly mixes with the main stream nor does the blowing momentum influence the cooling effectiveness, provided that a boundary-layer transition is not induced by the cooling gas injection. Our results, presented here, agree with the results of previous investigations of laminar supersonic flow conditions. Nomenclature C = Chapman-Rubesin factor c p = specific thermal capacity F = blowing ratio FU = flux vector in the x direction GU = flux vector in the y direction HU = vector of the viscosity terms h = enthalpy k = thermal conductivity M = momentum ratio Ma = Mach number _ m c = cooling gas mass flow _ m 0 e = mass flow entering the mixing layer from the main flow Pr = Prandtl number _ q = wall heat flux Re e = unit Reynolds number behind the front shock Re x 0 = Reynolds number based on x 0 s = slot width in the streamwise direction T = mean temperature in the mixing layer T aw;c = adiabatic wall temperature for the case with cooling T r = recovery temperature T 0;c = total temperature of the cooling gas t = slot width in the spanwise direction U = state vector u = velocity in the x direction v = velocity in the y direction x = distance from the leading edge, parallel to the model surface x ref = empirical reference length x s = distance from the leading edge to the center of the blowing opening x 0 = distance from the center of the blowing opening y = distance normal to the model surface = boundary-layer thickness = cooling effectiveness = heat transfer coefficient = dynamic viscosity = correlation factor = density = Mach angle = blowing angle Subscripts c = with cooling nc = no cooling e = boundary-layer edge 1 = freestream 0 = stagnation conditions Superscript * = values determined at the reference temperature

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Section: General Description Of the Flow Problemmentioning

confidence: 96%

Section: G Comparison With Data From Literaturementioning

confidence: 99%

mentioning

confidence: 99%

mentioning

confidence: 99%

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Experimental and Numerical Study of Cooling Gas Injection in Laminar Supersonic Flow

Heufer

Olivier

2008

AIAA Journal

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“…In this case the heat flux to the wall can be measured and has to be transferred into an equivalent temperature. The transformation is done by a similar approach given by Holden [13]. In general the heat flux to the surface can be expressed as follows…”

Section: Fig 1 Principle Of Film Coolingmentioning

confidence: 99%

Experimental Study of Active Cooling in 8 Laminar Hypersonic Flows

Heufer

Olivier

Notes on Numerical Fluid Mechanics and Multidisciplinary Design

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SummaryIn this study fundamental investigations concerning film cooling on an inclined flat plate and on a stagnation point region in laminar flow are performed. The influence of the relevant parameters, i.e. the Reynolds number of the free stream, the blowing ratio and the geometry of the blowing opening, is investigated by experiments. The results show the weak influence of the cooling gas injection on the outer flow. But the experimental data achieved in a shock tunnel demonstrate that already low blowing rates lead to a significant reduction of the thermal loads. Overall the results exhibit that the film cooling technique is an effective method for cooling bodies in laminar hypersonic flows.

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Numerical Simulation of Film Cooling in Supersonic Flow

Peter

Kloker

2020

Notes on Numerical Fluid Mechanics and Multidisciplinary Design

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High-order direct numerical simulations of film cooling by tangentially blowing cool helium at supersonic speeds into a hot turbulent boundary-layer flow of steam (gaseous H2O) at a free stream Mach number of 3.3 are presented. The stagnation temperature of the hot gas is much larger than that of the coolant flow, which is injected from a vertical slot of height s in a backward-facing step. The influence of the coolant mass flow rate is investigated by varying the blowing ratio F or the injection height s at kept cooling-gas temperature and Mach number. A variation of the coolant Mach number shows no significant influence. In the canonical baseline cases all walls are treated as adiabatic, and the investigation of a strongly cooled wall up to the blowing position, resembling regenerative wall cooling present in a rocket engine, shows a strong influence on the flow field. No significant influence of the lip thickness on the cooling performance is found. Cooling correlations are examined, and a cooling-effectiveness comparison between tangential and wall-normal blowing is performed.

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Experimental studies of shock wave/wall jet interaction in hypersonic flow

Cited by 32 publications

References 3 publications

Experimental and Numerical Study of Cooling Gas Injection in Laminar Supersonic Flow

Experimental and Numerical Study of Cooling Gas Injection in Laminar Supersonic Flow

Experimental Study of Active Cooling in 8 Laminar Hypersonic Flows

Numerical Simulation of Film Cooling in Supersonic Flow

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