Global Mode Analysis in the L/T Transition of a Hypersonic Boundary Layer Forced by Wall Injection

Thewis, André; Orlik, Evgeniy; Durant, Antoine; Fedioun, Ivan

doi:10.1016/j.piutam.2015.03.024

Cited by 3 publications

(3 citation statements)

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“…As a first step, and to setup the procedure, the DMD of 2D snapshots has been performed for ILES data of the T-313 configuration. 17 In the present study, we extend the analysis to 3D snapshots of the transitional BL in the BAM6QT, and compare the results to the experimental ones. The ILES methodology and results are described in section II.…”

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confidence: 87%

“…a It is not clear in the experiment whether the wall of the model is adiabatic or isothermal b See Ferrier et al 19 and Orlik et al 15 for a detailed analysis of the flow structure Contrarily to the procedure retained in previous studies 14, 17 where a developed LL profile was applied as boundary condition at the inlet of the computational domain, the ILES solution is entirely computed starting from the flat-plate leading edge. Numerical details are given hereafter.…”

Section: A Computational Setupmentioning

confidence: 99%

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Hypersonic Boundary-Layer Transition Tripped by Wall Injection: Global Mode Analysis

Thewis¹,

Durant²,

Fedioun³

2015

20th AIAA International Space Planes and Hypersonic Systems and Technologies Conference

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The boundary layer transition of a generic hypersonic forebody at Mach 6, tripped by wall injection, is investigated by means of Implicit Large Eddy Simulation supplemented by global mode analysis. An academic flat-plate configuration with boundary layer edge conditions matching those of the forebody is considered. Different transition mechanisms are observed downstream of the injection port, depending on the level of injection pressure. The global mode analysis is performed by Dynamic Mode Decomposition of numerical data. The spatial structure, temporal frequency and growth rate of 3D dynamic modes indicate some transient growth followed either by an exponential decay, or by a modal growth. These modes, that start in the jet counter-rotating vortices, are either varicose or sinusoidal, the later being more efficient than the former for effective tripping at low injection pressure. Recent experiments in the Boeing/AFOSR Mach-6 quiet tunnel at Purdue university show similar transition patterns, but need to be further analyzed. Nomenclaturef = frequency, s −1 h = height of the Mach disc, m J = momentum ratio P , T = pressure, Pa, temperature, K u, v, w = flow velocity components along x, y and z, m.s −1 x, y, z = streamwise, normal to the wall and spanwise coordinates, m δ = boundary layer thickness, m γ = ratio of specific heats ρ = density, kg.m −3 Subscript stat = static value tot = total (or stagnation) value inj = value at injection u = unit length ∞ = static value at infinity Superscript e = value in the free-stream, outside the BL H = conjugate-transpose

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mentioning

confidence: 87%

Section: A Computational Setupmentioning

confidence: 99%

Hypersonic Boundary-Layer Transition Tripped by Wall Injection: Global Mode Analysis

Thewis¹,

Durant²,

Fedioun³

2015

20th AIAA International Space Planes and Hypersonic Systems and Technologies Conference

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“…The interaction between the jet and the incoming BL results, in some conditions, in turbulent flow . To obtain more details on the flow structure, numerical simulations were previously carried out on the Jet In Supersonic Crossflow [15][16][17] .…”

Section: Jet-induced Transitionmentioning

confidence: 99%

Mach 6 Quiet Tunnel Laminar to Turbulent Investigation of a Generic Hypersonic Forebody

Durant¹,

Thewis²,

Schneider³

et al. 2015

20th AIAA International Space Planes and Hypersonic Systems and Technologies Conference

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This paper presents an experimental boundary layer transition investigation of the windward side of a generic hypersonic forebody performed in April 2015 in theBoeing/AFOSR Mach 6 quiet tunnel facility at Purdue university (BAM6QT). At 0 and 4 degrees of angle of attack, =11x10 6 /m, flow was fully laminar in quiet conditions. Under noisy conditions, an early transition front (Re θ~2 00) was observed, even when dividing the unit freestream Reynolds number by 6. In quiet conditions, several diamond-shaped roughness trips were found to efficiently trip the laminar boundary layer when > . Temperature-Sensitive Paint (TSP) enabled a global measurement of the heat flux distribution and detection of the transition front. PCB sensors confirmed the state of the boundary layer : laminar, turbulent or transitional. Transition results with a single continuously blowing sonic air jet was also collected at various pressure ratios, giving the laminar to turbulent threshold. NomenclatureCFD = Computational Fluid Dynamics BAM6QT = Boeing/AFOSR Mach 6 Quiet Tunnel NT = Natural Transition BLT = Boundary Layer Transition WT = Wind-Tunnel JISC = Jet In Supersonic Crossflow PR = Pressure Ratio LT = Laminar to Turbulent transition TSP = Thermal Sensitive Painting Re u = Unit Reynolds number, /m St = Stanton number δ = Laminar boundary layer thickness, mm μ = Dynamic viscosity, m2/s ρ = Density, kg/m3 u = Streamwise velocity, m/s M e = Edge boundary layer mach number P e = Edge boundary layer pressure T w = Wall temperature, K 2 k = Roughness element height, mm h = Mach disk height Re hh = Mach disk height Reynolds number Re kk = Roughness element height Reynolds number Re θ = Momentum thickness Reynolds number P i0 = Stagnation pressure, Pa T i0 = Stagnation temperature, K P ∞ = Freestream pressure, Pa T ∞ = Freestream temperature, K J = Jet momentum flux ratio

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Receptivity to Freestream Acoustic Noise in Hypersonic Flow over a Generic Forebody

Cerminara

Durant

Thewis

et al. 2019

Journal of Spacecraft and Rockets

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Direct numerical simulations (DNS) of the Navier-Stokes equations have been performed to investigate the receptivity and breakdown mechanisms in a Mach 6 flow over a generic forebody geometry with freestream acoustic disturbances. The simulations are based on transition experiments carried out in April 2015 in the Boeing/AFOSR Mach 6 facility at Purdue University. A three-dimensional model for both fast and slow freestream acoustic waves with multiple frequencies and spanwise wavenumbers has been adopted in the numerical simulations, for which high-amplitude disturbances have been considered in order to simulate noisy wind tunnel conditions. The numerical results reveal similarities in comparison to the experimental observations, especially when slow acoustic waves are considered as freestream disturbances. In particular, slow acoustic waves have been found to induce the breakdown process via crossflow instabilities located in the off-centerline region, with formation of streamwise streaks. Fast acoustic waves, in contrast, appear more efficient in inducing earlier nonlinear growth through destabilisation of the boundary layer along the symmetry plane of the body.Receptivity is the internalization process of the external disturbances into the boundary layer in the form of instability modes, and determines the initial conditions for the downstream linear growth of the unstable modes. In the case of high-amplitude disturbances, nonlinearities can become predominant in the 3 of 31 American Institute of Aeronautics and Astronautics receptivity process, which would lead directly to nonlinear growth and rapid transition to turbulence. In some particular conditions, an intermediate route to transition (between a linear modal and a fully nonlinear process) is possible, i.e. transient growth, consisting of a non-modal growth of linear instabilities, as in the case of lift-up-induced streamwise streaks in conjunction with secondary instabilities and consequent nonlinear breakdown [1,2].The body leading edge is a highly-receptive zone, due to the non-parallel effects and the related shortscale streamwise variations of the mean flow, which, in turn, cause a wavelength-conversion process from the scale of the external forcing to that of the induced boundary-layer disturbances [3]. At hypersonic Mach numbers, however, the small difference in phase speed between the forcing waves and the boundary-layer dominant modes can lead to a direct excitation of these modes via a resonance mechanism at the leading edge [4,5,6,7], without the need of a wavelength-conversion mechanism. By applying Fedorov's notation [5], the internal mode generated at the leading edge through the receptivity to fast acoustic waves is called the fast mode, or mode F, while the mode associated to the slow acoustic waves is known as the slow mode, or mode S, which is the mode pertaining to the class of the unstable boundary-layer modes. In hypersonic boundary layers, different unstable modes can coexist, including the first mode, corresponding to the Tollmien-...

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Global Mode Analysis in the L/T Transition of a Hypersonic Boundary Layer Forced by Wall Injection

Cited by 3 publications

References 31 publications

Hypersonic Boundary-Layer Transition Tripped by Wall Injection: Global Mode Analysis

Hypersonic Boundary-Layer Transition Tripped by Wall Injection: Global Mode Analysis

Mach 6 Quiet Tunnel Laminar to Turbulent Investigation of a Generic Hypersonic Forebody

Receptivity to Freestream Acoustic Noise in Hypersonic Flow over a Generic Forebody

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