Streamwise Vortex Instability and Transition on the Hyper-2000 Scramjet Forebody

Matsumura, Shin; Schneider, Steven P.; Berry, Scott A.

doi:10.2514/1.3959

Cited by 29 publications

(7 citation statements)

References 37 publications

(38 reference statements)

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“…The trace of vortex streaks are also visible on the windward side of the model at incidence angle of -12 degrees from the infrared thermography measurement, presented in Figure 18. The existence of streamwise vortices has also been reported by Matsumura et al 44 on the scramjet forebody test in Mach 6 flow. The generation of these vortices is attributed to the combination of freestream noise, leading edge imperfections, and the flow instabilities.…”

Section: Surface Pressure Mapping From Topviewsupporting

confidence: 70%

Pressure-sensitive paint on a truncated cone in hypersonic flow at incidences

Yang¹,

Erdem²,

Zare‐Behtash³

et al. 2012

International Journal of Heat and Fluid Flow

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Section: Surface Pressure Mapping From Topviewsupporting

confidence: 70%

Pressure-sensitive paint on a truncated cone in hypersonic flow at incidences

Yang¹,

Erdem²,

Zare‐Behtash³

et al. 2012

International Journal of Heat and Fluid Flow

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“…Some possible interactions between vortices emanating from the nose tip and CF or Görtler instabilities were also highlighted [30]. These geometries noticeably differ from the present forebody.…”

Section: A Modes: Terminologymentioning

confidence: 94%

Modal Linear Stability of the Near-Wall Flow on a Hypersonic Forebody

Ferrier

Fedioun

Orlik

et al. 2009

Journal of Spacecraft and Rockets

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The three-dimensional modal linear stability analysis is performed for the near-wall flow of a full-scale generic hypersonic vehicle forebody at flight Mach numbers 4, 6, and 8, at different angles of attack. The mean flow is computed with a Navier-Stokes commercial code. A physically sound, computationally efficient original method is proposed to define the integration path in the e N method. It has a significant impact on the computed N factors. The entropy-layer effect on the flow instability is analyzed in the framework of Lees and Lin's asymptotic theory. The entropy layer introduces an additional unstable mode, for which N 4. Results show that crossflow instability is dominant at Mach 6 and 8, whereas Mack's oblique first mode prevails at Mach 4. This mode is stabilized by a radiating wall, compared with an adiabatic one. Mack's second mode is also present at Mach 8. In any case, none of the instability modes that have been found is strong enough to provoke a natural transition in flight by itself: computed N factors do not exceed 10. Attempts to correlate the results with the Re =M e Const criterion are discussed.Nomenclature a = speed of sound, m=s C p , = heat capacity at constant pressure, J=kg K C v = heat capacity at constant volume, J=kg K f = frequency, Hz h = enthalpy per unit mass, J=kg k = thermal conductivity, W=m K k = wave vector (real), 1=m P = pressure, Pa r = C p C v , gas constant, J=kg K T = temperature, K t = time, s U, V, W = mean-flow components along x, y, and z, m=s V g = group velocity vector, m=s V ' = phase velocity, m=s W = molecular weight of species , kg=kmol X, Y, Z = coordinates in the global reference frame attached to the vehicle, m x, y, z = streamwise, transverse (normal to the wall), and spanwise coordinates, m X = mole fraction of species , = wave numbers (complex) in the x and z directions, 1=m = displacement thickness, m = momentum thickness, m g = direction of V g = viscosity, kg=m s = =, diffusivity, m 2 =s = density, kg=m 3 = amplification vector (real), 1=m = compressibility factor = direction of k = direction of ! = (real) pulsation, 1=sSubscripts M = maximum value (envelope method) u = unity w = value at the wall 1 = value at infinity, static value Superscripts e = value in the freestream, outside the boundary layer N = exponent in the e N method

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“…Liu and Sullivan 15 present a good background of the physics of TSP. Matsumura 16 successfully employed the TSP technique on a generic scramjet forebody in the BAM6QT. The methodology used in the present experiments is very similar to his.…”

Section: Experiments On the X-51a Forebodymentioning

confidence: 99%

Effect of Freestream Noise on Roughness-Induced Transition for the X-51A Forebody

Borg

Schneider

2008

Journal of Spacecraft and Rockets

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A 20%-scale X-51A forebody model was tested in the Boeing/AFOSR Mach-6 Quiet Tunnel. Repolishing the nozzle throat has restored quiet flow at high Reynolds numbers. The effect of a smooth blank and two different trip strips on windward-forebody transition was measured using temperature-sensitive paint and hot-wire anemometry. Reducing freestream noise from conventional to quiet levels increased the smooth-wall transition Reynolds number by a factor of at least 2.2. In addition, the transition Reynolds number based on the distance from the trips increased by a factor of 2.4 for the smaller trips and by a factor of 1.7 for the larger trips. Thus, tunnel noise had a substantial effect on roughness-induced transition. I. Hypersonic Laminar-Turbulent Transition Laminar-turbulent transition in hypersonic boundary layers is important for prediction and control of heat transfer, skin friction, and other boundary layer properties. Vehicles that spend extended periods at hypersonic speeds may be critically affected by the uncertainties in transition prediction, depending on their Reynolds numbers. Although slender vehicles are the primary concern, blunt vehicles are also affected by transition. 1 However, the mechanisms leading to transition are still poorly understood, even in low-noise environments. Many transition experiments have been carried out in conventional ground-testing facilities over the past 50 years. 2 However, these experiments are contaminated by the high levels of noise that radiate from the turbulent boundary layers normally present on the wind tunnel walls. 3 These noise levels, typically 0.5-1% of the mean, are an order of magnitude larger than those observed in flight. 4, 5 These high noise levels can cause transition to occur an order of magnitude earlier than in flight. 3, 5 In addition, the mechanisms of transition operational in small-disturbance environments can be changed or bypassed altogether in high-noise environments; these changes in the mechanisms change the parametric trends in transition. 4 Mechanismbased prediction methods must be developed, supported in part with measurements of the mechanisms in quiet wind tunnels.

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Streamwise Vortex Instability and Transition on the Hyper-2000 Scramjet Forebody

Cited by 29 publications

References 37 publications

Pressure-sensitive paint on a truncated cone in hypersonic flow at incidences

Pressure-sensitive paint on a truncated cone in hypersonic flow at incidences

Modal Linear Stability of the Near-Wall Flow on a Hypersonic Forebody

Effect of Freestream Noise on Roughness-Induced Transition for the X-51A Forebody

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