Expansion corner effects on hypersonic shock wave/turbulent boundary-layer interactions

White, Michael E.; Ault, David A.

doi:10.2514/3.24157

Cited by 15 publications

(7 citation statements)

References 4 publications

Supporting

Mentioning

Contrasting

Order By: Relevance

“…In the simulation, the flow condition at the left boundary of the domain is specified as M 1 3.5, P 1 21473.1 Pa, and T 1 527.3 K because the freestream flow is already compressed by the ramp shock. In addition, according to previous investigations [16,23], the boundary layer thickness at the duct entrance is typically between 15 and 20% of the duct height for hypersonic inlets. Therefore, the thickness of the coming boundary layer is specified as 0.15h in the current research.…”

Section: Description Of Simplified Inlet Modelmentioning

confidence: 96%

“…Chung and Lu [15] examined the interaction of a Mach 8 turbulent boundary layer with an externally generated shock impinging within the thickness of the boundary layer upstream and downstream of a convex corner. Later, White and Aultt [16] investigated the behavior of the shock wave/turbulent boundary layer interactions in the vicinity of an expansion corner at Mach 11.5. These investigations arrived at many valuable conclusions.…”

Section: Introductionmentioning

confidence: 99%

See 1 more Smart Citation

Influence of Expansion Waves on Cowl Shock/Boundary Layer Interaction in Hypersonic Inlets

Zhang

Tan

Zhuang

et al. 2014

Journal of Propulsion and Power

View full text Add to dashboard Cite

The cowl shock/boundary layer interaction is of great importance to the performance and operability of hypersonic inlets. However, it is usually affected by the expansion waves originated from the convex corner of the ramp surface (often called the shoulder), making the commonly used separation prediction relation unreliable. Therefore, the cowl shock/boundary layer interaction under the interference of the expansion waves is investigated by both theoretical analysis method and computational method in this paper. The results show that the expansion waves bring significant influence on the cowl shock/boundary layer interaction and four kinds of interaction processes exist when their relative position changes. When different interaction processes dominates the flowfield, the expansion waves may bring positive or negative effects on the shock/boundary layer interaction. In particular, while the cowl shock impinges near the shoulder, the interaction processes of shock-shock-expansion wave and shock-expansion waveshock bring positive effects on the shock/boundary layer interaction and suppress the boundary layer separation. Nomenclature d = distance between the shock impingement point and the shoulder d 1 = distance between the shoulder and the first demarcation point of the cowl shock/boundary layer interaction d 2 = distance between the shoulder and the second demarcation point of the cowl shock/boundary layer interaction h = height of the isolator M 0 = upstream Mach number M 1 = Mach number at the duct entrance M 3 = Mach number at the flow region downstream from the corner expansion fan and upstream from the cowl shock p sep = minimum pressure downstream of the shock when the boundary layer separation occurs p sh = pressure downstream of the incident shock p 0 = static pressure of the coming flow x = X-axis coordinate α = angle between the bottom wall of the duct and the cowl shock downstream from the expansion waves β = cowl contracting angle β 0 = angle of the shock generator δ = thickness of the boundary layer δ 0 = thickness of the boundary layer at the shoulder δ = displacement thickness of the boundary layer θ = momentum thickness of the boundary layer φ = ramp angle

show abstract

Section: Description Of Simplified Inlet Modelmentioning

confidence: 96%

Section: Introductionmentioning

confidence: 99%

Influence of Expansion Waves on Cowl Shock/Boundary Layer Interaction in Hypersonic Inlets

Zhang

Tan

Zhuang

et al. 2014

Journal of Propulsion and Power

View full text Add to dashboard Cite

show abstract

“…To take into account the viscous flow, to some extent, the Mach number profile of the boundary layer is applied to calculate the viscous cowl shock shape near the corner. The incoming flow is discretized by a number of parallel layers with a certain Mach number, 11 and then the cowl shock/nonuniform flow interaction is successively solved from outer to the wall by constructing twodimensional Riemann problems at the intersection points. To avoid singular shock, the calculation ends at an incoming Mach number of 1.95 in the boundary layer.…”

Section: Figure 13 Stretch Of Inviscid Flow In the Near Wall Regionmentioning

confidence: 99%

Leading Edge Bluntness Effects on Shock Wave/Boundary Layer Interactions Near a Convex Corner

Yang²

2015

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

View full text Add to dashboard Cite

The cowl shock/ingested forebody boundary layer interaction is usually encountered near the hypersonic inlet shoulder, a convex corner, and sometimes causes the inlet to unstart. The bluntness of the forebody leading edge also affects the downstream convex corner flow, and that should be carefully considered to improve the inlet performance. A twodimensional inlet with a sharp leading edge and a series of blunt leading edges at Mach 5 were numerically studied. The relative positions of the cowl shock and the convex corner were changed to reveal the characteristics of the shock wave/boundary layer interactions with the corner expansion wave interference. Flow patterns and the distributions of surface pressure, heat flux, and friction coefficient involved were thoroughly studied. The results indicate that the boundary layer is even thicker and the flow is easier to separate when the bluntness of the leading edge increases. The variation trends of the separation extent with a sharp leading edge agree reasonably well with the estimations on the basis of inviscid analysis. However, the inviscid formulas established in the sharp leading edge case are not suitable for the blunt leading edge cases. With the aid of viscous simulations, it is found that the upstream effect of the corner extend a little farther and the dimensionless relative position range that can largely reduce the separation narrows with the increasing of the leading edge bluntness. To take into account the nonuniform flow in the boundary layer, a rapid calculation method of the incident shock (B. L.) was established. It is shown that the shock (B. L.) impinges on the corner can minimize the separation in the sharp leading edge case. However, the shock (B. L.) has to move downstream to minimize the separation when the leading edge is blunt. NomenclatureC * = Chapman-Rubensin constant, (µ/µ∞)(T∞/T) CD = drag coefficient of a hemicylinder cp = specific heat capacity of air at constant pressure d1 = distance between the convex corner and the first demarcation point of the inviscid shock impingement d2, d2', d2" = distance between the convex corner and the second demarcation point of the inviscid shock impingement d3 = distance between the convex corner and the third demarcation point of the inviscid shock impingement dn = diameter of the blunt leading edge Ls, Lsx = separation bubble length along the wall and along the x-axis, respectively Ma = Mach number M1, M2 = Mach number upstream and downstream of the incident shock, respectively M3 = Mach number downstream of the reflected shock at the flat plate Mc = Mach number downstream of the corner expansion M∞ = freestream Mach number p∞ = freestream static pressure q = heat flux rate Rn = radius of the blunt leading edge Re = Reynolds number 2 St = Stanton number, St = q/[ρ∞U∞cp (T0-Tw)] T0, T∞, Tw = Total temperature, freestream temperature, and wall temperature, respectively U∞ = freestream flow velocity x = x-axis coordinate xc = distance from the convex corner to the leading edge of the flat plate xi, xr,...

show abstract

“…Furthermore, Chung and Lu 7 reported that the expansion corner located in the vicinity of SWBLI significantly affects the upstream influence location and flow separation. A similar experimental study of different shock impingement locations near the expansion corner was carried out by White and Ault 8 . The investigation was mainly performed to generate a data set for SWBLI in the vicinity of an expansion corner to compare it with the numerical results.…”

Section: Introductionmentioning

confidence: 99%

“…A similar experimental study of different shock impingement locations near the expansion corner was carried out by White and Ault. 8 The investigation was mainly performed to generate a data set for SWBLI in the vicinity of an expansion corner to compare it with the numerical results. They reported a reduction in bubble size when its edge is closely located to the corner.…”

mentioning

confidence: 99%

Shock/expansion fan interaction with real gas effects for Earth and Mars atmospheric conditions

2022

View full text Add to dashboard Cite

Numerical simulations are performed to investigate the real gas effects on shock/expansion fan interaction. Initial perfect gas simulations at low enthalpy capture the flow structures efficiently and outcomes are found to have excellent agreement with the analytical calculations. Furthermore, the simulations with the real gas solver for different enthalpies showed that the variation in enthalpy significantly changes the flow structures. It is observed that an increase in enthalpy leads to a decrease and increase in the postshock and postexpansion fan Mach numbers, respectively. Another important observation is the decrement in the peak pressure ratio with an increment in the enthalpy. These effects are noted to be more pronounced for Mars's environment due to the higher dependency of specific heat on temperature.computational fluid dynamics, Earth and Martian atmospheric simulations, expansion fan, hypersonic flow, real gas flow, shock waves | INTRODUCTIONThe design of high-speed 1,2 aircraft necessitates the appropriate predictions of flow structures associated with it. These flights often encounter shocks and expansion fans. Furthermore, these shock waves and expansion fans may interact with each other. The presence of shock/ expansion fan interaction may complicate the flow structures by affecting the boundary-layer

show abstract

Expansion corner effects on hypersonic shock wave/turbulent boundary-layer interactions

Cited by 15 publications

References 4 publications

Influence of Expansion Waves on Cowl Shock/Boundary Layer Interaction in Hypersonic Inlets

Influence of Expansion Waves on Cowl Shock/Boundary Layer Interaction in Hypersonic Inlets

Leading Edge Bluntness Effects on Shock Wave/Boundary Layer Interactions Near a Convex Corner

Shock/expansion fan interaction with real gas effects for Earth and Mars atmospheric conditions

Contact Info

Product

Resources

About