Shockwave-boundary layer interference heating analysis

Hung, F. T.; Barnett, D. O.

doi:10.2514/6.1973-237

Cited by 23 publications

(9 citation statements)

References 1 publication

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“…The peak heating rates for laminar, turbulent, and transitional shock impingement regions have been correlated by Hung and Barnett [4]. It was found that when 𝑅𝑒 𝐿 > 5 • 10 5 the boundary layer flow is sufficiently disturbed that SWBLI disturbance promotes transition.…”

Section: Resultsmentioning

confidence: 95%

“…incident shock causes the boundary layer to separate and rapidly transition to turbulence [2]. Hypersonic transitional SWBLIs can result in significantly greater peak surface heat transfer than laminar or turbulent interactions [3][4][5][6] and are a phenomenon of great interest, especially for the outer surfaces of hypersonic cruise vehicles where in many cases the boundary layer is laminar up to the shock interaction. This is a particular concern for hypersonic cruise vehicles wherein the actuation of control surfaces may lead to an SWBLI interaction, triggering transition or even boundary layer separation and leading to very high off-design thermal and pressure loads.…”

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confidence: 99%

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Experimental Studies of Hypersonic Shock Impingement on a Transpiration-Cooled Flat Plate

Naved

Hermann

Hambidge

et al. 2023

Journal of Spacecraft and Rockets

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Hypersonic vehicle design requires mitigation of the high heat fluxes present in regions of shock-wave/boundary-layer interactions. A candidate technology that may be applied locally to these regions is transpiration cooling. In this work, experiments were conducted in the University of Oxford’s high-density tunnel at Mach 6.1 in both laminar and turbulent undisturbed boundary-layer regimes where a 10 deg shock generator impinged a strong oblique shock wave onto a transpiration-cooled microporous injector. For the laminar boundary layer, due to the strength of the incident shock, a transitional shock-wave/boundary-layer interaction region was formed with peak heating over 50 times greater than the nominal laminar level. Both nitrogen and helium were used as coolants. Relatively low levels of helium injection of [Formula: see text] for the transitional and [Formula: see text] for the turbulent scenarios were sufficient to reduce the heat transfer downstream of shock interaction to approximately 50% of the value without cooling. In fact, helium is highly effective with a similar cooling performance achieved as eight times the equivalent mass flux of nitrogen. The experimental data are correlated, and both the turbulent and transitional shock-impingement scenarios display a similar trend of reduced surface heat transfer with higher blowing parameters. Empirical fits are proposed that may be used for initial systems design.

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Section: Resultsmentioning

confidence: 95%

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confidence: 99%

Experimental Studies of Hypersonic Shock Impingement on a Transpiration-Cooled Flat Plate

Naved

Hermann

Hambidge

et al. 2023

Journal of Spacecraft and Rockets

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“…Here, q w,max /q w,0 is the ratio between the maximum heat transfer and the upstream boundary layer, and p max / p ∞ is the pressure ratio across the shock wave. The value n was determined as 0.85 or 0.80 for turbulent boundary layers (Back & Cuffel 1970;Hung & Barnett 1973;Hankey & Holden 1975), and 0.70 for laminar boundary layers (Hankey & Holden 1975). While the n values from the literature were based on the separation of either fully laminar or turbulent boundary layers, the present investigation showed an upstream laminar boundary layer transitioning to turbulence by the shock impingement.…”

Section: Mean Surface Properties Of the 12mentioning

confidence: 99%

Hypersonic shock impingement studies on a flat plate: flow separation of laminar boundary layers

et al. 2022

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Interactions between shock waves and boundary layers produce flow separations and augmented pressure/thermal loads in hypersonic flight. This study provides details of Mach 7 impinging-shock-flat-plate experiments conducted in the T4 Stalker Tube. Measurements were taken at flow conditions of Mach 7.0 (2.44 MJ kg $^{-1}$ ) and Mach 7.7 (2.88 MJ kg $^{-1}$ ) flight enthalpies with a range of freestream unit Reynolds numbers from $1.43 \times 10^{6}$ m $^{-1}$ to $5.01 \times 10^{6}$ m $^{-1}$ . A shock generator at $12^{\circ }$ or $16^{\circ }$ to the freestream created an oblique shock which impinged on a boundary layer over a flat plate to induce flow separation. The flow field was examined using simultaneous measurements of wall static pressure, heat transfer and schlieren visualisation. Measured heat transfer along the flat plate without the shock impingement indicated that the boundary layer remained laminar for all flow conditions. The shock impingement flow field was successfully established within the facility test duration. The onset of separation was observed by a rise in wall pressure and a decrease in heat transfer at the location corresponding to the stem of the separation shock. Downstream of this initial rise, an increased pressure and higher heating loads were observed. The heat-transfer levels also indicated an immediate boundary layer transition due to the shock impingement. The separation data of the present work showed good agreement with our previous work on shock impingement on heated walls (Chang et al., J. Fluid Mech., vol. 908, 2021, pp. 1–13). A comparison with the previous scaling indicated that the separation also relates to the pressure ratio and the wall temperature parameter.

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“…Thermal protection design requires the severe peak heat flux in the reattachment region of separation states. There are several methods for predicting the peak heat flux 1,2,[27][28][29][42][43][44] , and the peak heat flux is generally correlated with peak pressure rise in these methods. One simple correlation is expressed in the following form 28,29…”

Section: A Wall Shear Stress In the Reversed-flow Regionmentioning

confidence: 99%

“…The peak heat flux in the reattachment region is an important feature of separated flow. Most empirical methods to predict the peak heat flux always correlate it with peak pressure rise 1,2,[27][28][29] . Studies on wall pressure distributions are abundant.…”

Section: Introductionmentioning

confidence: 99%

Hypersonic curved compression ramp flows with bistable states

Tang,

Wang,

et al. 2021

Preprint

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The bistable states and separation hysteresis in curved compression ramp (CCR) flows, and the corresponding aerothermal characteristics (including wall friction C f , pressure and heat flux St), are studied numerically and theoretically. Direct numerical simulations of separation hysteresis induced by variation of turning angle φ , as well as the influence of inflow Mach number and wall temperature on hysteresis loops, are carried out. Distributions of wall friction, pressure and heat flux are analyzed. Further, emergence of C f 's first and second minima in the separation bubble is interpreted, revealing it is dominated by the adverse pressure gradient (APG) induced by separation and reattachment shocks. The present results and analysis indicate that the reversed-flow singularity of Smith (Proceedings of the Royal Society of London. A. Mathematical and Physical Sciences, 1988, 420: 21-52) is less likely to occur in CCR flows. The prediction of peak pressure p pk of separation states confirms the model based on the minimum viscous dissipation theorem (Physics of Fluids, 2020, 32(10):101702). While the pressure overshoot p os can be analyzed by shock-polars with pressure match of compression and expansion process. The correlation between peak heat flux and peak pressure rise of both separation and attachment states is also discussed in terms of the classical power relations.

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Shockwave-boundary layer interference heating analysis

Cited by 23 publications

References 1 publication

Experimental Studies of Hypersonic Shock Impingement on a Transpiration-Cooled Flat Plate

Experimental Studies of Hypersonic Shock Impingement on a Transpiration-Cooled Flat Plate

Hypersonic shock impingement studies on a flat plate: flow separation of laminar boundary layers

Hypersonic curved compression ramp flows with bistable states

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