Study on shock wave and turbulent boundary layer interactions in a square duct at Mach 2 and 4

Sugiyama, Hiromu; Minato, Ryojiro; Mizobata, Kazuhide; Tojo, Akira; Muto, Yohei

doi:10.1007/s11630-006-0037-3

Cited by 24 publications

(8 citation statements)

References 9 publications

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“…The duct must also be long enough to sustain the full pressure rise. The data of Sullins and Mclafferty [14] (M 0 1.92) correspond to a relatively thin boundary layer (A ∕A 5.8 × 10 −2 ), as does the data of Sugiyama et al [9] (M 0 1.96; A ∕A 6.5 × 10 −2 ); and both fall just below the M 0 2 line of Fig. 8.…”

Section: Maximum Backpressure In a Constant-area Ductsupporting

confidence: 63%

“…A schlieren image of a λ-type pseudoshock at Mach 2 is shown in Fig. 3a, taken from [9]. Flow is from left to right, and the shock train consists of multiple shocks that are normal in the center of the square duct.…”

Section: Flow Structure In Backpressured Ductsmentioning

confidence: 99%

“…It has no normal shock, but it consists of multiple crossing oblique shocks. Figure 3b, also taken from [9], shows a schlieren image of the same duct at Mach 4, showing a set of crossing oblique shocks with no normal component typical of the X-type interaction. Figure 4 shows wall static pressure distributions in pseudoshocks at different Mach numbers, taken from [10].…”

Section: Flow Structure In Backpressured Ductsmentioning

confidence: 99%

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Flow Modeling of Pseudoshocks in Backpressured Ducts

Smart

2015

AIAA Journal

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Scramjet operation in the lower hypersonic regime between Mach 4 and 7 is characterized by what is called dualmode combustion. In this situation, the backpressure generated by heat release in the combustor separates the wall boundary layer upstream of fuel injection, forming a highly complex flow known as a pseudoshock. The pseudoshock phenomenon has been studied extensively in the experimental literature, and its structure has been found to be affected by the duct Mach number, the extent of the wall boundary layer, and the streamwise area distribution of the duct. This paper presents a quasi-one-dimensional analysis technique for calculating the flow through a pseudoshock. It is quite general, and it can be used for ducts with any supersonic inflow, cross-sectional shape, and streamwise area distribution. Comparison of the results of this analysis with experimental data indicates that it provides a good estimate of both the interaction length and the pressure distribution for X-type pseudoshocks that occur in ducts with Mach numbers above two.

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Section: Maximum Backpressure In a Constant-area Ductsupporting

confidence: 63%

Section: Flow Structure In Backpressured Ductsmentioning

confidence: 99%

Section: Flow Structure In Backpressured Ductsmentioning

confidence: 99%

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Flow Modeling of Pseudoshocks in Backpressured Ducts

Smart

2015

AIAA Journal

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“…This is accomplished by a so called shock train, a system of oblique or normal shocks that has been investigated in detail for symmetric inflow conditions (e.g., by Billig and Kothari [7] and Billig [8]) and is also described in the standard textbook by Heiser and Pratt [9]. Depending on the combustionchamber static pressure, the flow structure in the isolator varies (Sugiyama et al [10]). For a scramjet compression setup with nonsymmetric inflow conditions to the isolator, the flow structure is visualized (i.e., by Hermann and Koschel [11] and Reinartz et al [12]) for different backpressures.…”

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confidence: 99%

Experimental Investigation of a Two-Dimensional and a Three-Dimensional Scramjet Inlet at Mach 7

Häberle

Gülhan

2008

Journal of Propulsion and Power

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The current experimental study at Mach 7 was carried out to investigate the viscous flow effects in a scramjet inlet with a fixed geometry operating at Mach 7.5-8. To reduce the risk of inlet unstart, an optional passive boundarylayer bleed was integrated at the throat and tested for different bleed dimensions. The passive bleed significantly reduces the lip-shock-induced separation bubble on the ramp. But the correct location of the bleed gap is important for the reduction of the separation bubble. To investigate the compression behavior, different backpressures were applied and the effects on the internal flow structure were analyzed by means of static pressure along the isolator and pitot pressure profiles at the exit of the isolator. Because the reference configuration has a very small internal contraction ratio to guarantee self-starting, the internal compression ratio was increased in three steps by means of additional sidewall compression along the internal flowpath. This includes internal contraction ratios above the starting limit. This measure led to a significant increase of the static pressure ratio at the end of the isolator. The complex flow structures are explained by means of detailed static and pitot pressure plots. Nomenclature= mass flow, kgs 1 p = pressure, Nm 2 q = dynamic pressure, Nm 2 R = specific gas constant for air 287.15, Jkg 1 K 1 Re = Reynolds number T = temperature, K = calibration value = throttle degree, 1 A mfm;th A 1 0 100% = density, kgm 3 = total pressure recovery Subscripts CC = combustion chamber entry condition mfm = mass flow meter SW = sidewall compression inlays 1 and 2 th = throat t0 = wind-tunnel total conditions w = wall 0 = freestream plane ahead of the inlet 1 = wind-tunnel freestream conditions

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“…As described in the previous nominally two-dimensional LES, inflow turbulence profiles are computed simultaneously in a separate mesh; however, since the flow is no longer homogeneous in the spanwise direction, recycling/rescaling with dynamic reflection is no longer an appropriate method for prescribing inflow turbulence. Instead, we make use of the digital filtering method developed by Klein et al 39 and improved by Touber and Sandham 40 expanded to produce turbulent boundary layer profiles along adiabatic walls at both y/δ r = 0 and z/δ r ** While the flow is generally symmetric for normal shock trains at moderate Mach number such as the one under consideration, it is important to remember that in more intense interactions including oblique shock trains, the flow may be asymmetric 38 .…”

Section: Les Of Channel With Sidewallsmentioning

confidence: 99%

Large-Eddy and RANS Simulations of a Normal Shock Train in a Constant-Area Isolator

Morgan

Duraisamy

Lele

2012

50th AIAA Aerospace Sciences Meeting Including the New Horizons Forum and Aerospace Exposition

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Large-eddy simulations (LES) of a normal shock train in a constant-area isolator (M ∞ = 1.61, Re θ = 1660) are carried out with a high-order compact differencing scheme using localized artificial diffusivity (LAD) for shock capturing. We examine sensitivities of the solution to a variety of physical modeling assumptions. Simulations with spanwise periodic boundary conditions are first performed, the results of which are compared to experiment and validated with a three-level grid refinement study. Due to the computational cost associated with resolving near-wall structures, the LES is run at a Reynolds number lower than that in the comparison experiment; thus, the confinement effect of the turbulent boundary layers is not exactly duplicated. While this discrepancy affects the location of the first normal shock, the overall structure of the shock train and its interaction with the boundary layers matches the experiment quite closely. Observations of pertinent physical phenomena in the experiment such as a lack of reversed flow and the development of secondary shear layers are captured by the simulation. Next, a series of linear eddy viscosity based Reynolds Averaged Navier Stokes (RANS) simulations are performed to assess the ability of reduced-order models to accurately capture the complicated physics of multiple shock/boundary layer interactions. It is found that the RANS solutions exhibit a wide range of variations, indicating a strong sensitivity -particularly with respect to streamwise location of the initial shock --to model formulation. Finally, three-dimensional effects due to the side walls of the isolator are investigated by performing LES of the same shock train interaction in a three-dimensional, low-aspect-ratio rectangular duct geometry. A significant three-dimensional flow feature is observed in the region of the strong initial shock, which agrees well with experimental oil flow visualizations. NomenclatureC f = skin friction coefficient k = turbulence kinetic energy h = isolator half-height l = turbulent length scale L x , L y , L z = extent of the computational mesh M = Mach number N x , N y , N z = number of grid points in computational mesh p = static pressure Re = Reynolds number S ij = Strain-rate tensor s L = length of shock train t = time T = temperature u = velocity vector u = streamwise velocity component u τ = friction velocity w = value at wall v = wall-normal velocity component w = spanwise velocity component δ = boundary layer thickness δ ij = Dirac delta θ = momentum thickness μ t = eddy viscosity ρ = density τ = shear stress ω = specific dissipation rate Superscript + = wall unit quantity ′ = fluctuation quantity inn = inner boundary layer value out = outer boundary layer value 2 Subscript ∞ = freestream value in = value at isolator inlet out = value at isolator outlet r = reference value ahead of initial shock VD = van Driest-transformed quantity

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Study on shock wave and turbulent boundary layer interactions in a square duct at Mach 2 and 4

Cited by 24 publications

References 9 publications

Flow Modeling of Pseudoshocks in Backpressured Ducts

Flow Modeling of Pseudoshocks in Backpressured Ducts

Experimental Investigation of a Two-Dimensional and a Three-Dimensional Scramjet Inlet at Mach 7

Large-Eddy and RANS Simulations of a Normal Shock Train in a Constant-Area Isolator

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