Nosetip bluntness effects on cone frustum boundary layer transition in hypersonic flow

Stetson, Kenneth F.

doi:10.2514/6.1983-1763

Cited by 122 publications

(63 citation statements)

References 5 publications

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“…Small leading edge bluntness moves the boundary layer transition backwards, but an increase of bluntness radius leads to some shift of transition forward. Reverse of boundary layer transition at the cone blunting is known [6,7]. Backward shift of boundary layer transition at the bluntness increase can be explained easily: it is caused by the decrease of Reynolds number calculated with entropy layer characteristics and by entropy layer elongation (before its absorption by boundary layer) at the increase of bluntness radius r. But it is more complicated to explain reversive shift of boundary layer transition forward at further increase of bluntness radius.…”

Section: Resultsmentioning

confidence: 99%

Temperature sensitive paint application for investigation of boundary layer transition in short-duration wind tunnels

Borovoy

Mosharov

Noev

et al. 2012

Progress in Flight Physics

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Application of Temperature Senstive Paint (TSP) technology for investigation of boundary layer transition in short-duration wind tunnels is presented. Investigations were made on sharp and blunted §at plates in UT-1 wind tunnel of TsAGI operated in Ludwieg scheme at Mach numbers 5, 6, and 8 (Reynolds numbers from 5.5 · 10 6 to 26 · 10 6 ). Both natural and induced boundary layer transitions were investigated. BACKGROUNDTemperature Sensitive Paint method is a very e¨ective optical technique for heat §ux measurements. It provides the pattern of heat transfer rates on the whole visible model surface that is essential for three-dimensional (3D) §ows. In few recent years, this technique is successfully developed in TsAGI for short-duration wind tunnels [1,2].Temperature Sensitive Paint is a specially developed paint containing speci¦c luminophore which luminescence depends on temperature. Being excited by the light of appropriate spectral range, the TSP emits the luminescent light of longer wavelength range, and luminescent light intensity decreases with an increase of temperature. Thus, the temperature distribution can be determined in each point of investigated surface from noncontact optical measurement of TSP luminescence intensity by acquiring an image of the whole investigated surface. But luminescence intensity depends not only on temperature but also on excitation light intensity and TSP thickness. To exclude an e¨ect of excitation light intensity and TSP thickness, it is necessary to acquire second reference image of investigated surface at known temperature. Normalization of active image (acquired at unknown temperature distribution) on reference image exclude an e¨ects of Progress in Flight Physics 3 (2012) 15-24

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Section: Resultsmentioning

confidence: 99%

Temperature sensitive paint application for investigation of boundary layer transition in short-duration wind tunnels

Borovoy

Mosharov

Noev

et al. 2012

Progress in Flight Physics

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“…This implies that the nose Reynolds number determines the stability and the transition of boundary layers over blunt bodies. Most of the wind tunnel experiments discussed in the introduction [2][3][4][5] were performed in conventional tunnels where the freestream disturbance levels were high. The computed transition N-Factor for these tunnels varies in the range of 3.5 to 4.5 [5,8,10].…”

Section: A Linear Stability: Similarity Profilesmentioning

confidence: 99%

“…Stetson [2] studied the effects of nosetip bluntness on the transition onset for sphere-cone boundary layers at freestream Mach numbers of M 6 and 9. The investigation mainly concentrated on the transition in the entropy-swallowing region.…”

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confidence: 99%

“…In supersonic and hypersonic boundary layers, one important geometrical parameter is nose bluntness. The effects of bluntness on transition in supersonic and hypersonic boundary layers have been studied experimentally [1][2][3][4][5][6] and numerically [7][8][9][10][11][12][13] by many researchers. It was identified that the entropy layer that is formed near the bow shock region persists for a long distance downstream and makes the boundary layer more stable when compared to the sharpcone case.…”

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confidence: 99%

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Effects of Nose Bluntness on Hypersonic Boundary-Layer Receptivity and Stability over Cones

2011

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The receptivity to freestream acoustic disturbances and the stability properties of hypersonic boundary layers are numerically investigated for boundary-layer flows over a 5 straight cone at a freestream Mach number of 6.0. To compute the shock and the interaction of the shock with the instability waves, the Navier-Stokes equations in axisymmetric coordinates were solved. In the governing equations, inviscid and viscous flux vectors are discretized using a fifth-order accurate weighted-essentially-non-oscillatory scheme. A third-order accurate total-variationdiminishing Runge-Kutta scheme is employed for time integration. After the mean flow field is computed, disturbances are introduced at the upstream end of the computational domain. The appearance of instability waves near the nose region and the receptivity of the boundary layer with respect to slow mode acoustic waves are investigated. Computations confirm the stabilizing effect of nose bluntness and the role of the entropy layer in the delay of boundary-layer transition. The current solutions, compared with experimental observations and other computational results, exhibit good agreement. Nomenclature = streamwise wave number C recpt = receptivity coefficient c p = specific heat at constant pressure c v = specific heat at constant volume E = total energy e = molecular internal energy F, G = inviscid flux vector in the axial and radial directions

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“…The boundary layer is for instance strongly stabilized by slight nose bluntness as already observed at lower hypersonic Mach numbers. 11,[40][41][42] For R N = 4.75 , the free-stream unit Reynolds numbers tested were too low to reach a turbulent boundary layer by the end of the model, only a departure from the laminar state is visible at the rear of the model for the largest free-stream unit Reynolds number. With the 10 mm nosetip, the flow remains laminar at least until R B = 10R N .…”

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confidence: 99%

Hypersonic Boundary Layer Transition on a 7 Degree Half-Angle Cone at Mach 10

Grossir

Regert

Pinna

et al. 2014

7th AIAA Theoretical Fluid Mechanics Conference

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Hypersonic boundary layer transition experiments are performed in the low-enthalpy Longshot wind tunnel with a free-stream Mach number ranging between 12 ≥ M∞ ≥ 9.5 and Reynolds number between 12 × 10 6 /m ≥ Reunit,∞ ≥ 3.3 × 10 6 /m. The model is an 800 mm long 7• half-angle cone with nosetip radii between 0.2 and 10 mm. Instrumentation includes flushmounted fast-response thermocouples and pressure sensors. Boundary layer transition onset location is determined from the wall heat flux distribution. Nose bluntness has a strong stabilizing effect. No transition reversal could be observed at RB = 10RN for a Reynolds number based on the nosetip radius of ReR N ,∞ = 123, 000. Increasing freestream unit Reynolds number results in larger Rex B ,e. Wavelet analysis of the boundary layer fluctuations shows that numerous wave packets are present during the transition process. Comparison with Linear Stability Theory results for second mode waves shows an excellent agreement for the most amplified frequencies. The N -factor of the wind-tunnel is 5 based on these computations and on the transition location measured experimentally. The convection velocity of the disturbances is closely approximated by the local boundary layer edge velocity for all conditions investigated. Schlieren flow visualization of the instabilities exhibits the typical rope shape of second mode disturbances for the sharpest nosetips. For nose bluntness larger than 4.75 mm, disturbances are mainly present at the edge of the boundary layer and within the inviscid shock layer. Their shape no longer presents the second mode typical structure although a frequency analysis of the disturbances is still compatible with second mode instabilities. Present results confirm the dominance of second mode waves in the transition process along a conical geometry for Mach numbers larger than 10.

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Nosetip bluntness effects on cone frustum boundary layer transition in hypersonic flow

Cited by 122 publications

References 5 publications

Temperature sensitive paint application for investigation of boundary layer transition in short-duration wind tunnels

Temperature sensitive paint application for investigation of boundary layer transition in short-duration wind tunnels

Effects of Nose Bluntness on Hypersonic Boundary-Layer Receptivity and Stability over Cones

Hypersonic Boundary Layer Transition on a 7 Degree Half-Angle Cone at Mach 10

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