Görtler-type vortices in hypersonic flows: the ramp problem

Chevalerie, D. Aymer de la; Fonteneau, A.; Luca, Luigi De; Cardone, Gennaro

doi:10.1016/s0894-1777(97)00051-4

Cited by 40 publications

(20 citation statements)

References 5 publications

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“…3) as lines with constant non-dimensional wavelength K, defined as a Gö rtler number based on the disturbance wavelength [4,33]:…”

Section: Gö Rtler Numbersmentioning

confidence: 99%

“…Disturbances associated to leading edge imperfections may trigger the instability, destabilising the flow and subsequently producing transition either by interaction with, e.g., Tollmien-Schlichting waves, or by the Gö rtler instability itself. Leading edge irregularities have been reported to be of great importance [4] for the distribution and formation of these vortices.…”

Section: Introductionmentioning

confidence: 99%

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Numerical simulation of Görtler vortices in hypersonic compression ramps

Navarro‐Martinez

Tutty

2005

Computers & Fluids

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“…3) as lines with constant non-dimensional wavelength K, defined as a Gö rtler number based on the disturbance wavelength [4,33]:…”

Section: Gö Rtler Numbersmentioning

confidence: 99%

Section: Introductionmentioning

confidence: 99%

Numerical simulation of Görtler vortices in hypersonic compression ramps

Navarro‐Martinez

Tutty

2005

Computers & Fluids

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“…IR thermography is widely applied in many hypersonic applications; for example, Cardone [24] described the use of IR thermography for heat flux measurements in hypersonic high enthalpy wind tunnels, Gulhan et al [25] estimated the aerothermal heating caused by jet-hypersonic crossflow interaction and Di Clemente et al [26] used IR measurements to validate numerical computations. IR thermography is particularly used in hypersonic flows to obtain the surface temperature distribution because of the presence of high spatial gradients caused by vortical structures: de Luca et al [16,27,28], Aymer de La Chevalerie et al [29] and Schrijer [30] analysed several aspects of shock wave boundary layer interaction (SWBLI) in the presence of Görtler vortices that were induced by surface irregularities installed on or near the leading edge.…”

Section: Introductionmentioning

confidence: 99%

Image resection and heat transfer measurements by IR thermography in hypersonic flows

Avallone

Greco

Ekelschot

2013

Quantitative InfraRed Thermography Journal

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This study describes a data reduction technique to estimate heat fluxes in hypersonic flows based on measurements of the temperature distribution on the surface of the model by means of infrared (IR) thermography. IR images are registered using a camera model based on perspective projection approach taking into account lens distortion. Relative displacements between model and IR camera, due to sting mechanical vibrations, can increase the noise. An approach to reduce it, based on single-step discrete Fourier transform and Speed-Up Robust Features methods, is presented. Heat flux is computed with an inverse heat transfer problem applying the trust region reflective algorithm. The proposed data reduction technique is numerically validated and applied to an experimental test carried out in the Hypersonic Test Facility Delft (HTFD) at Mach number equal to 7.5

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“…Experimental studies of the flow around plane and axisymmetric configurations at moderate hypersonic Mach numbers revealed the emergence of streamwise structures [1, 2] and a significant increase in the maximum values of the heat-transfer coefficient [3]. An apparent reason for these phenomena is the emergence and development of transversal nonuniformity induced by natural roughness of the leading edge of the model.…”

Section: Introductionmentioning

confidence: 99%

Effect of streamwise structures on heat transfer in a hypersonic flow in a compression corner

Bazovkin

Kovchavtsev

Kuryshev

et al. 2009

J Appl Mech Tech Phy

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Coefficients of heat transfer to the surface in a laminar hypersonic flow (M ∞ = 21) over plane and axisymmetric models with a compression corner are presented. These coefficients are measured by an infrared camera. The parameters varied in the experiments are the angle of the compression corner and the distance to the corner point. Characteristics of the flow with and without separation in the corner configuration are obtained. The measured results are compared with direct numerical simulations performed by solving the full unsteady Navier-Stokes equations. Experiments with controlled streamwise structures inserted into the flow are described. A substantial increase in the maximum values of the heat-transfer coefficient in the region of flow reattachment after developed laminar separation is demonstrated. Key words: hypersonic flow, laminar separation in the compression corner, streamwise structures, heat transfer to the surface. Introduction.A large amount of theoretical and experimental studies of heat transfer in a supersonic flow in a compression corner is currently available. These activities are important because corner surfaces are essential elements of inlets of supersonic aviation engines, which create conditions for flow compression and for static pressure increase. This portion of the air duct of the engine displays extremely high heat intensity, and its temperature stability is responsible for reliable operation of the aircraft engine as a whole. This problem is particularly relevant in hypersonic flight with several-fold increased stagnation temperatures and, hence, heat fluxes to the surface.One important feature of a hypersonic flow in a compression corner is the emergence of transversal nonuniformity of the flow, which can be enhanced in the field of centrifugal forces generated by variations of the flow direction. Experimental studies of the flow around plane and axisymmetric configurations at moderate hypersonic Mach numbers revealed the emergence of streamwise structures [1, 2] and a significant increase in the maximum values of the heat-transfer coefficient [3]. An apparent reason for these phenomena is the emergence and development of transversal nonuniformity induced by natural roughness of the leading edge of the model. Experimental and numerical studies without modeling transversal nonuniformity were performed for high Mach numbers, in particular, for conditions of a hypersonic viscous shock layer. It should be noted that numerical simulations (with allowance for three-dimensional effects) of hypersonic flows in corner configurations are rather difficult, and numerical algorithms require experimental verification. It seems of interest, therefore, to measure the heat-transfer coefficient in compression corners at high Mach numbers and in the presence of transversal nonuniformity.

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Görtler-type vortices in hypersonic flows: the ramp problem

Cited by 40 publications

References 5 publications

Numerical simulation of Görtler vortices in hypersonic compression ramps

Numerical simulation of Görtler vortices in hypersonic compression ramps

Image resection and heat transfer measurements by IR thermography in hypersonic flows

Effect of streamwise structures on heat transfer in a hypersonic flow in a compression corner

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