Compression Ramp Flow Research Articles

Experimental investigation of the effects of leading edge bluntness on supersonic flow over a double compression ramp

Understanding the flow characteristics over a double compression ramp is crucial for high-speed vehicle design. Leading edge bluntness is a key factor influencing the formation of a separation region on a double compression ramp flow. In the present study, the effect of bluntness on a double compression ramp is investigated experimentally at a nominal Mach number of 4. The test model has 13° and 40° inclinations with respect to the freestream. Five different levels of leading-edge radius, varying from 0.0 to 2.0 mm, were subjected to supersonic wind tunnel tests. Shadowgraph and infrared thermography techniques were employed to visualize the flow features of the double ramp model. Measurements of surface heat-transfer along the centerline of the test model were obtained from the acquired infrared images. It is shown that the leading-edge radius alters the separation characteristics as well as the surface heat-transfer. Possible reasons for such flow characteristics are discussed.

Direct Numerical Simulation of the Pulsed Arc Discharge in Supersonic Compression Ramp Flow

Direct numerical simulation (DNS) of shock wave/turbulent boundary layer interaction (SWTBLI) with pulsed arc discharge is carried out in this paper. The subject in the study is a Ma=2.9 compression flow over a 24-degree ramp. The numerical approaches were validated by the experimental results in the same flow conditions. The heat source model was added to the Navier-Stokes equation to serve as the energy deposition of the pulsed arc discharge. Four streamwise locations are selected to apply energy deposition. The effect of the pulsed arc discharge on the ramp-induced flow separation has been studied in depth. The DNS results demonstrate the incentive locations play a dominant role in suppressing the separated flow. Results show that pulsed heating is characterized by a thermal blockage, which leads to streamwise deflection. The incentive locations upstream the interaction zone of the base flow have a better control effect. The separation bubble shape shows as “spikes”, and the downstream flow of the heated region is accelerated due to the momentum exchange between the upper boundary layer and the bottom boundary layer. The high-speed upper fluid is transferred to the bottom, and thus enhances its ability to resist the flow separation. More stripe vortex structures are also generated at the edge of the flat-plate. Furthermore, the turbulent kinetic disturbance energy is increased in the flow filed. The disturbances that originate from the pulsed heating are capable of increasing the turbulent intensity and then diminishing the trend of flow separation.

Direct Simulation of Fluid–Structure Interaction in a Hypersonic Compression-Ramp Flow

Sustained flight at hypersonic speeds presents an enduring challenge to vehicle design and control. An extreme aerothermal environment acting on geometrically thin, multifunctional structures can result in the significant structural response of key aerodynamic surfaces. Specifically, the oscillations of a compression corner shock-wave/boundary-layer interaction, which models a control surface deflection, can couple with the lowest natural frequencies of a proximal compliant surface, leading to a range of undesirable outcomes from reduced service life to structural failure. The present work investigates these phenomena through time-accurate high-fidelity aeroelastic simulations of a laminar two-dimensional flow at over a 35 deg compression ramp featuring an embedded compliant panel. Surface pressure, skin friction, and heat transfer associated with the corner shock-wave/boundary-layer interaction are directly compared between rigid and compliant configurations. The upstream separation zone is observed to synchronize with the compliant-panel deflection. A reduction in surface heat flux is observed for a majority of compliant cases relative to the rigid case, whereas entirely new flow features are observed in response to large-scale panel deflections. A modification is proposed to the local piston-theory aerodynamic model to improve the accuracy of surface pressure predictions.

Reattachment streaks in hypersonic compression ramp flow: an input–output analysis

We employ global input–output analysis to quantify amplification of exogenous disturbances in compressible boundary layer flows. Using the spatial structure of the dominant response to time-periodic inputs, we explain the origin of steady reattachment streaks in a hypersonic flow over a compression ramp. Our analysis of the laminar shock–boundary layer interaction reveals that the streaks arise from a preferential amplification of upstream counter-rotating vortical perturbations with a specific spanwise wavelength. These streaks are associated with heat-flux striations at the wall near flow reattachment and they can trigger transition to turbulence. The streak wavelength predicted by our analysis compares favourably with observations from two different hypersonic compression ramp experiments. Furthermore, our analysis of inviscid transport equations demonstrates that base-flow deceleration contributes to the amplification of streamwise velocity and that the baroclinic effects are responsible for the production of streamwise vorticity. Finally, the appearance of the temperature streaks near reattachment is triggered by the growth of streamwise velocity and streamwise vorticity perturbations as well as by the amplification of upstream temperature perturbations by the reattachment shock.

Effect of Forcing on a Supersonic Compression Ramp Flow

Simulations were carried out of a 24° compression ramp in Mach 2.25, turbulent flow. In addition to simulations of the baseline flow, calculations were carried out with flow control applied, in the form of a streamwise body force with a wall-normal profile chosen to match a perturbation velocity profile previously identified through conditional averaging. Forcing in the upstream boundary layer with this particular form was found to drive large-scale separation unsteadiness downstream. Power spectral densities of the wall pressure fluctuations in the separated region displayed a low-frequency component that increased in proportion to the amplitude of forcing. The separation location and separation shock location were both observed to be synchronized with the forcing. These results are consistent with the Plotkin model of separation unsteadiness, and present an example of a case where large-scale separation motion is sensitive to external disturbances.

Görtler Vortices in Hypersonic Flow on Compression Ramps

Direct numerical simulation of Görtler vortices in hypersonic compression ramp flow is performed. Three ramp angles of 15, 20, and 25 deg are considered for a freestream Mach number of 7.7 and a unit Reynolds number of . The results show that a mesh resolution of 151 million points is sufficient to capture the flow details. The validation tests present excellent agreement with theoretical values and experimental data. Görtler vortices show unsteady features for the considered ramp flows. The induced heat flux variation and the vortex structure of Görtler vortices are discussed. Two methods of determining the Görtler number are proposed. The first one gives the variation of the Görtler number near reattachment based on the computational results. Additionally, based on a simplified model, a formula depending only on the flow condition and the geometry of the ramp model is deduced to predict the Görtler number. It is shown that the latter method can quantitatively predict the Görtler number near reattachment, and the predictions are in qualitative agreement with a theoretical stability analysis. In the flowfield, corrugated reattachment resulting from the three-dimensionality of the separation bubble is found. Görtler vortices originating in the shear layer as a result of streamline curvature are enhanced at and after flow reattachment.

Selective upstream influence on the unsteadiness of a separated turbulent compression ramp flow

A study of the shock-wave/boundary-layer interaction induced by a compression ramp was carried out using high-fidelity simulations. The objective was to investigate the influence of upstream disturbances on low-frequency separation unsteadiness. Two computations were performed for a 24° compression ramp at Mach 2.25, one highly resolved case and one reduced-resolution case. The reduced-resolution case was run for an extended duration to capture many cycles of low-frequency unsteadiness. Basic flow characteristics, including statistics on the boundary layer, wall pressure, and skin friction, were computed. Frequency spectra were calculated to confirm the presence of low-frequency unsteadiness. The influence of upstream disturbances on large-scale separation unsteadiness was investigated using correlations, filtering, and conditional averaging based on the position of the primary separation shock. Low-frequency unsteadiness was found to be related to structures near the wall (y/δ < 0.5) with a time scale greater than 20δ/U∞, whereas higher frequency separation motion could be attributed to turbulent boundary layer structures with a time scale on the order of δ/U∞. The finding that the separation region responds selectively to certain large-scale, near-wall perturbations in the incoming flow supports a model of separation unsteadiness in which external forcing by certain components of boundary layer turbulence drives a weakly damped global mode of the separation bubble. This contrasts with suggestions that have been made in the literature that the separation region oscillates on its own, as in an amplified global mode.

Numerical analysis of exhaust jet secondary combustion in hypersonic flow field

The interaction effect between jet and control surface in supersonic and hypersonic flow is one of the key problems for advanced flight control system. The flow properties of exhaust jet secondary combustion in a hypersonic compression ramp flow field were studied numerically by solving the Navier–Stokes equations with multi-species and combustion reaction effects. The analysis was focused on the flow field structure and the force amplification factor under different jet conditions. Numerical results show that a series of different secondary combustion makes the flow field structure change regularly, and the temperature increases rapidly near the jet exit.

Spectrum study on unsteadiness of shock wave–vortex ring interaction

Shock oscillation with low-frequency unsteadiness commonly occurs in supersonic flows and is a top priority for the control of flow separation caused by shock wave and boundary layer interaction. In this paper, the interaction of the shock caused by the compression ramp and the vortex rings generated by a micro-vortex generator (MVG) in a supersonic flow at Ma = 2.5 is simulated by the implicit large eddy simulation method. The analysis of observation and the frequency of both the vortex ring motion and the shock oscillation is carried out. The results show that the shock produced by a compression ramp flow at Ma = 2.5 has a dominant non-dimensional low frequency, which is around St = 0.002, while the vortex rings behind the MVG have a dominant high frequency which is around St = 0.038. The dominant low frequency of the shock, which is harmful, can be removed or weakened through the shock-vortex ring interaction by the vortex rings which generate high frequency fluctuations. In the shock and vortex ring interaction region, a dominant high frequency St = 0.037–0.038 has been detected rather than the low frequency St = 0.002, which indicates that the vortex ring is stiff enough to break or weaken the shock. This analysis could provide an effective tool to remove or weaken the low frequency pressure fluctuation below 500 Hz, which has a negative effect on the flight vehicle structures and the environmental protection, through the high frequency vortex generation.

Gas kinetic scheme for turbulence simulation

The extended gas-kinetic scheme (GKS) for turbulence simulations is developed based on the generalized BGK equation with the effective relaxation time. This relaxation time can be computed from the turbulent viscosity, through which turbulence models can be directly combined. For engineering low-cost simulations of high Reynolds number flows, common-used RANS models are applied, while the LES and hybrid RANS/LES (DES and IDDES) models as well as the minimized dispersion and controllable dissipation (MDCD) reconstruction are adopted in high fidelity turbulence simulations. In addition, the turbulent transport equations are solved in a strongly coupled way by using GKS with scalar transport. The extended GKS is applied in typical turbulent flow predictions including the RANS simulation of hypersonic compression ramp flow and the detailed simulation of multiscale turbulent structures in low-speed cylinder flow with hybrid models. The predicted results agree well with existing experimental measurements and numerical studies, which shows the good accuracy, resolution and robustness of the extended GKS and reveals the wide prospects in turbulence simulations on different model scales, including the multiscale models such as hybrid RANS/LES methods. Furthermore, the multiscale turbulence simulation methods are worthy of further studies based on the generalized BGK model.

Letter: Transient interaction between plasma jet and supersonic compression ramp flow

The rapid flow evolution between a plasma jet and a 20° compression ramp flow is captured by a high-speed schlieren system at Mach 2.0. Several interesting flow phenomena are observed for the first time. The pulsed jet, which generates strong perturbations, forces the crossflow boundary layer to separate and forms a forward moving shock. A typical shock-on-shock interaction occurs when the precursor shock intersects with the original shock. The interaction is initially regular, and then it transforms into an irregular one with a Mach stem connecting the precursor shock and original ramp shock.

초음속 이중 압축 램프의 앞전 곡률에 따른 천이 유동 해석

Accurate prediction of supersonic transition is required for the heat transfer estimation over supersonic double compression ramp flows. Correlation-based transition models were assessed for a supersonic double ramp problem. Numerical results were compared with experimental data from RWTH Aachen University. A parametric study on the nose bluntness was performed using a selected transition model. As the nose bluntness increases, the boundary layer thickness is increased and the triple point of shock interactions moves downstream. The peak magnitude of the heat transfer is consequently decreased with the nose bluntness.

Experimental studies on flow visualization and velocity field of compression ramp with different incoming boundary layers

Experimental studies which focus on flow visualization and the velocity field of a supersonic laminar/turbulent flow over a compression ramp were carried out in a Mach 3.0 wind tunnel. Fine flow structures and velocity field structures were obtained via NPLS (nanoparticle-tracer planar laser scattering) and PIV (particle image velocimetry) techniques, time-averaged flow structures were researched, and spatiotemporal evolutions of transient flow structures were analyzed. The flow visualization results indicated that when the ramp angles were 25°, a typical separation occurred in the laminar flow, some typical flow structures such as shock induced by the boundary layer, separation shock, reversed flow and reattachment shock were visible clearly. While a certain extent separation occurred in turbulent flow, the separation region was much smaller. When the ramp angles were 28°, laminar flow separated further, and the separation region expanded evidently, flow structures in the separation region were complex. While a typical separation occurred in turbulent flow, reversed flow structures were significant, flow structures in the separation region were relatively simple. The experimental results of velocity field were corresponding to flow visualization, and the velocity field structures of both compression ramp flows agreed with the flow structures well. There were three layered structures in the U component velocity, and the V component velocity appeared like an oblique “v”. Some differences between these two compression ramp flows can be observed in the velocity profiles of the shear layer and the shearing intensity.

Instability of supersonic compression ramp flow.

The instability of supersonic compression ramp flow is investigated. It is assumed that the Reynolds number is large and that the governing equations are the unsteady triple-deck equations. The mean flow is first calculated by solving the steady equations for various scaled ramp angles α, and the numerical results suggest that there is no singularity for increasing ramp angles. The stability of the flow is investigated using two approaches, first by solving the linearized unsteady equations and looking for global modes proportional to e(λt). In the second approach, the linearized unsteady equations are solved numerically with various initial conditions. Whereas no globally unsteady modes could be found for the range of ramp angles studied, the numerical simulations show the formation of wavepacket type disturbances which grow and convect and reach large amplitudes. However, the numerical results show large variations with grid size even on very fine grids.

Stability, separation and close body interactions. Introduction.

Stability, separation and close body interactions. Introduction.

Numerical simulation of ramp induced shock wave/boundary-layer interaction in turbulent flow

AbstractA computational study has been carried out to analyse the shock wave turbulent boundary layer interaction in a two-dimensional compression ramp flow for a free stream Mach number of 2·94. Ramp angles ranging from 8° to 24° are used to produce the full range of possible flow fields, including flows with minor separation, moderate separation, and significant amount of separation. The model has been analysed using 2D numerical simulations based on a commercially available Computational Fluid Dynamics (CFD) Code, Fluent, that employs k-ω Shear Stress Transport (SST) turbulence model. The computed data for surface pressure distribution indicated a good agreement with experiment. Numerical results obtained through the present series of computations indicate an increased extent of separated zone, and thus show increased upstream influence when compared to experiment. Total pressure loss has shown to increase downstream of separation location and increase when corner angle increases. Secondary separation has been observed for higher angles.

Improvement of Baldwin–Lomax turbulence model for supersonic complex flows

Entropy represents the dissipation rate of energy. Through direct numerical simulation (DNS) of supersonic compression ramp flow, we find the value of entropy is monotonously decreasing along the wall-normal direction no matter in the attached or the separated region. Based on this feature, a new version of Baldwin–Lomax turbulence model (BL-entropy) is proposed in this paper. The supersonic compression ramp and cavity-ramp flows in which the original Baldwin–Lomax model fails to get convergent solutions are chosen to evaluate the performance of this model. Results from one-equation Spalart–Allmaras model (SA) and two-equation Wilcox k–ω model are also included to compare with available experimental and DNS data. It is shown that BL-entropy could conquer the essential deficiency of the original version by providing a more physically meaningful length scale in the complex flows. Moreover, this method is simple, computationally efficient and general, making it applicable to other models related with the supersonic boundary layer.

A three-equation turbulence model for high-speed flows

Adding a new equation to the two-equation K-ω turbulence model framework, this paper proposed a three-equation turbulence model to determine the density variance for high-speed aero-optics and high-speed compressible turbulent flows. Simulations were performed with the new model for supersonic and hypersonic flat-plate turbulent boundary layer and hypersonic ramp flows. The results showed that the prediction with the present model agrees well with the experimental data and is significantly better than the Lutz’s model in predicting the density variance for the flat-plate flows. Furthermore, the present model can produce more accurate skin pressure and skin heat flux distributions than the original K-ω model in simulating hypersonic compression ramp flows with separation and reattachment and shock/boundary layer interactions. Without introducing a variety of ad hoc wall damping and wall-reflection terms, the proposed three-equation turbulence model is applicable to high-speed aero-optics and turbulent flows of real vehicles of complex configuration.

An Improved K-ω Model for High Speed Flows

An improved K-ω model, which allows for compressible corrections, is proposed in this paper. Numerical scheme was established utilizing the improved Total Variation Diminishing (TVD) scheme and applying implicit scheme to the negative source terms of the turbulence model. Hypersonic flat-plate boundary-layer flows and hypersonic compression ramp flows marked with separation, reattachment and shock/boundary layer interactions are then computed. Comparisons between the computational results, the experimental results and the semi-empirical formulations show that the compressible correction term of the K-ω turbulence model is a pressure-dilatation correlation. In addition, for flow with separation and without separation, calculation results of wall pressures, friction coefficients and wall heat transfer rate distributions using the improved model and established scheme agree better with the experimental results than that using the original K-ω model.

An Improved Shear Stress Transport(SST) Model for High Speed Flows

An improved Shear Stress Transport(SST) model, which allows for the compressible corrections, is proposed in this study. Numerical scheme was established by taking advantage of the improved Total Variation Diminishing (TVD) scheme and by applying implicit scheme to the negative source terms of the turbulence model. Hypersonic flat-plate boundary-layer flows and hypersonic compression ramp flows marked with separation, reattachment and shock/boundary-layer interactions are then computed. The comparisons between the computational results, the experimental results and the semi-empirical formulations show that the compressible correction term of the SST turbulence model is the scalar product of the weighted density average of the turbulent fluctuating velocity and the pressure gradients of the average flow field correlation quantities. In addition, for flow with separation and without separation, calculation results of wall pressures, friction coefficients and wall heat transfer rate distributions using the improved model and established scheme agree better with the experimental results than that using the original SST model.