Separation In Hypersonic Flow Research Articles

Unsteady interaction mechanism of transverse stage separation in hypersonic flow for a two-stage-to-orbit vehicle

Due to the complex aerodynamic interaction, the safe separation of two stages is one of the challenges for the successful launch of a two-stage-to-orbit (TSTO) vehicle. The unsteady hypersonic flow past the parallel-staged TSTO model during stage separation at Ma = 6.7 and Re = 8.86 × 105 m−1 is numerically studied using laminar flow simulation. The TSTO model consisted of a waverider and a spaceplane as booster and orbiter, respectively. The effect of the center of gravity (CoG) of the orbiter on the unsteady aerodynamic interference during stage separation of TSTO is analyzed in detail with 0.65 ≤ lCoG/lo ≤ 0.80. In addition, the aerodynamic characteristics, dynamic behaviors, and unsteady wall pressure variation are compared in different cases. The results show that the CoG regime is limited to 5% of the orbiter length for absolutely safe separation, i.e., 0.70 < lCoG/lo < 0.75. As for the unsuccessful separation, the orbiter tends to fly nose-down if lCoG/lo ≤ 0.70 while tending to pitch or somersault when lCoG/lo = 0.80. Furthermore, the pitching moment of the orbiter, which is influenced by the interstage shock wave–boundary layer interaction and shock–shock interaction, dominates the separation safety, and the specific flow mechanisms concerning the separation behavior associated with aerodynamic interference in different cases are analyzed in detail.

CFD investigation of shock boundary layer interaction in hypersonic flow and flow control using micro ramps

PurposeThis study aims to investigate the prevalence of shock boundary layer interaction (SBLI) in air-breathing intake system is highly undesirable since this leads to high pressure gradients, typical stream mutilation and pressure drop. A novel flow control mechanism is incorporated in this research holding an array configuration of passive flow control device (micro ramps [MR]) that is adapted to improve the boundary layer stability.Design/methodology/approachTwo geometric variants of the MR, namely, MR40 and MR80 is considered which reduce the pressure drop during SBLI. The incidence oblique shock wave angle of 34° is considered for the modelling. Large eddy simulation (LES) turbulence model was used with subgrid models of Wall modelled LES, Smagorinsky–Lilly to compute the unsteady effects of SBLI control using micro vortex generators. The unsteady results are compared with steady Reynold’s average Naviers–Stoke’s equation for calibrating the turbulence models.FindingsThe array configuration of MR80 reduces the pressure drop by 22% as compared with no ramp configuration and also reduces the flow distortion in hypersonic inlet. The most affected region of the MR is in the vicinity of center-line. Quantitative results prove that the upstream influence of the shock waves has been largely reduces by MR80 array configuration as compared to single MR80 pattern configuration. Different vortex structures found in the experiments was exclusively predicted using LES.Originality/valueThis paper substantiates the requirement of MR array configuration for transferring the momentum from free stream to the boundary layer and thereby energizing the boundary layer. This process of energization delays the flow separation in hypersonic flow.

Numerical Investigation of the Transient Nature of a Laminar Separation Bubble in Hypersonic Flow

Shock-wave/boundary-layer interaction is a prime research topic in the design of hypersonic vehicles. For the proper designing of hypersonic vehicles, especially, their thermal protection systems, it is necessary to understand the time-dependent behavior of the wall properties and the corresponding pressure and thermal loads. Hence investigation is carried out to understand the unsteady nature of shock-induced laminar boundary layer separation for a simple canonical configuration of a two-dimensional ramp. A density-based non-reactive Navier-Stokes solver named rhoCentralFoam in openFOAM is employed in the present investigation. The detailed physics of laminar boundary layer separation in hypersonic flow is investigated through timewise behavior of the streamline patterns and the surface properties, such as the pressure, heat flux, and friction coefficients. It is found that at the onset of fluid flow it is its inviscid characteristics that are predominant and there is occurrence of very small separation bubble. The separation bubble grows, as the involvement of viscous characteristics increases gradually, and ultimately it attains a steady state. Because of separated boundary layer the heat flux and skin friction coefficients are found to follow the diffused V and deformed W shapes, respectively. The peaks of pressure and thermal loads are found to exist in the vicinity of the reattachment region. These peaks are higher in the initial stage of the process and attain steady state eventually. The high pressure and thermal loads may cause structural damage to the vehicle and hence their correct prediction is necessary. Thus, the current investigation is helpful in the design of thermal protection systems of hypersonic vehicles.

Numerical study of bluntness effects on laminar leading edge separation in hypersonic flow

Bluntness effects on laminar hypersonic leading edge separation are investigated numerically at Mach number $M\approx 10$, unit Reynolds number $Re=1.3\times 10^{6}~\text{m}^{-1}$, specific enthalpy $h_{o}=3.1~\text{MJ}~\text{kg}^{-1}$ and wall-to-stagnation temperature ratio $T_{w}/T_{o}=0.1$. Such effects are important from an experimental point of view and because bluntness can affect a separated flow favourably or adversely. In this study, two blunt leading edge cases of small radius ($15~\unicode[STIX]{x03BC}\text{m}$) and large radius ($100~\unicode[STIX]{x03BC}\text{m}$) are investigated. A comparison with the idealised sharp leading edge case is also given. General flow features and surface parameters such as the shear stress, pressure and heat flux are presented and analysed. The results have also been interpreted in terms of Cheng’s displacement-bluntness similitude parameter affecting the size of separation. Previous experiments by Holden delineated small and large bluntness effects based on Cheng’s parameter and considering small to moderate separated regions. In this study, leading edge separation was found to suppress the favourable effect of bluntness in delaying separation. Bluntness, furthermore, seemed to promote the appearance of secondary vortices within a main separated region. Analysis of reverse flow boundary layer profiles such as velocity, pressure and temperature is also given. It is shown that bluntness accentuates large transverse gradients. This in turn adversely effects the reverse flow boundary layer leading to the appearance of secondary vortices.

Experimental and numerical study of the laminar separation in hypersonic flow

This article is devoted to an experimental and numerical study of shock wave/boundary layer interaction in hypersonic laminar flow ( M = 10). The experimental was performed in the ONERA R5Ch wind tunnel on a hollow cylinder flare. The flow stream delivered by the R5Ch wind tunnel produces physical conditions which justify both the theoretical approach using classical Navier-Stokes equations and the approach by Direct Simulation Monte-Carlo. So the aim of this study is to improve the capacity of Navier-Stokes and DSMC codes to predict high Mach number interactions. Pressure and heat flux have been measured at the wall and compared with the results obtained with Navier-Stokes and DSMC codes. Two different meshes have been considered with the three Navier-Stokes codes used. The conclusion is that the codes can give a good evaluation of the physical quantities at the wall.

Separation phenomenon in a hypersonic flow with strong wall cooling: Subcritical regime

This paper is concerned with a hypersonic flow past a cooled curved wall. The separation phenomenon at high Reynolds numbers is described using an interactive boundary-layer approach. For a problem of this type, a distinction must be made between two regimes subject to whether the global interaction between the boundary layer and the external flow is strong or weak. This depends on a hypersonic interaction parameter x ∼ MeRe−1/2: if x is small, the interaction is weak; if it is large, the interaction is strong. The present study is applied to a flow with a weak global interaction regime when x → 0. There are three principal motivations for the study. First, the problem of hypersonic boundary-layer separation is obviously of great practical interest. Second, the desire to demonstrate that marginal separation theory (Ruban, 1982; Stewartson et al., 1982) may be successfully applied to the description of a hypersonic flow separation. Third, the interest to verify the theoretical prediction by Smith (1988) of a local breakdown or stall occurring in any interactive boundary-layer solution at a finite value of the controlling parameter.

Second-order longitudinal curvature effects in compressible laminar boundary layers

11 Kaufman, L. G., II et al., of Hypersonic Flow Separation and Control Characteristics, ASD-TDR-62-168, March 1962, Grumman Aircraft Engineering Corp., New York. 12 Werle, M. J., A Critical Review of Analytical Methods for Estimating Control Forces Produced by Secondary Injection/' NOLTR 68-5, Jan. 1968, United States Naval Ordnance Lab., White Oak, Md. 13 Hawk, N. E. and Amick, J. L., Two-Dimensional Secondary Jet Interaction with a Supersonic Stream, AIAA Journal, Vol. 5, No. 4, April 1967, pp. 555-660. 14 Zakkay, V., Erdos, J., and Calarese, W., An Investigation of the Interaction Between a Transverse Sonic Jet and a Hypersonic Stream, NYU AA-68-27, Aug. 1968, Aerospace Labs., New York University, N.Y.

Stability of compressible boundary layers.

with emphasis on the structure of the jet and separation ahead of the jet, Separated Flow, Part 2, Proceedings of a Specialists Meeting, AGARD Fluid Dynamics Panel (AGARD, Paris, France 1966), pp. 667-700. 16 Needham, D. A. and Stallery, J. L., Boundary layer separation in-hypersonic flow, AIAA Paper 66-455 (1966). 17 Maurer, F., Interaction effects of lateral control jets expanding into a supersonic flow, Deutsche Versuchsanstalt fur Luftund Raumfahrt, Munich, Germany Rept. 65-04 (1965). 18 Wu, J.-M., personal communication, based on R. E. Wilson, An experimental investigation of separated turbulent supersonic flow at M = 3.0, M.S. Thesis, Univ. of Tennessee (June 5, 1966). 19 Reshotko, E. and Tucker, M., Effect of a discontinuity on turbulent boundary-layer-thickness parameters with application to shock-induced separation, NACA TN 3454 (May 1955). 20 Mager, A., On the model of the free shock-separated turbulent boundary layer, J. Aeronaut. Sci. 23, 181-184 (1956). 21 Zukoski, E. E., Review of data concerning turbulent boundary layer separation in front of a forward facing step, Guggenheim Jet Propulsion Center, California Institute of Technology, Pasadena, Calif., Rept. 66-8-15-1 (August 1966).

Laminar boundary-layer separation in hypersonic flow.

Laminar boundary-layer separation in hypersonic flow.

A heat-transfer criterion for the detection of incipient separation in hypersonic flow

A heat-transfer criterion for the detection of incipient separation in hypersonic flow