Shock Impingement Location Research Articles

Numerical study on the nonlinear characteristics of shock induced two-dimensional panel flutter in inviscid flow

For an aeroelastic system of a two-dimensional elastic panel subjected to an impinging inviscid oblique shockwave, the nonlinear flutter characteristics are affected by many factors such as shock impingement location, cavity pressure and initial perturbation. The effects of the above factors on the variation of system bifurcation type and dynamic behaviors are investigated numerically. A low-fidelity computational method coupled with local piston theory and van Karman plate model, and a high-fidelity computational method coupled with Euler equations and finite element model are used for fluid-structure interaction simulations. Two sets of new findings are unveiled. First, either the variation of shock impingement location or cavity pressure can induce the aeroelastic system to transition between a subcritical bifurcation and a supercritical bifurcation. For some cases, the system bifurcation characteristics exhibit strong sensitivity to these two factors. Second, it is found that in addition to the limit cycle oscillation in the form of a combination of the second and third structural modes, multiple stable limit cycle oscillations due to the coupling of higher-order modes can be triggered by suitable initial perturbations. These limit cycle oscillations have high frequencies and amplitudes, which suggest a higher risk of structural fatigue failure.

High-Speed Fluid-Structure Interactions on a Compliant Panel Under Shock Impingement

This paper reports experimental results of fluid/structure interactions induced by shock impingement on a clamped-free/clamped-free compliant panel in a Mach 5.8 Ludwieg tube. A shock generator is used to statically impinge a shock on the panel with the aim of producing flow-induced vibrations. In turn, these vibrations create an aeroelastic coupling between the panel and the flowfield, which is dominated by the shock-wave/boundary-layer interactions. The influence of the structural compliance and shock impingement strength is assessed by testing a nominally rigid panel and compliant panels with thickness-to-length ratios of 0.0034 and 0.0068 with a shock generator at flow deflection angles of 0, 2, 3, 8, and 10 deg. One of the primary findings of the present work is that the aeroelastic response of the panel tends to shift from a structurally dominated to a fluid-dominated regime, which is sustained by either the freestream fluctuations or the fluid/structure coupling, depending on the structural compliance and shock impingement strength and location. This work is of fundamental nature, and its main goal is to analyze the interactions between high-speed flow and a thin compliant panel of canonical geometry. However, studies like this are crucial to characterize those complex couplings and provide experimental data that can help enhance the accuracy of aeroelastic modeling tools that feed into the design process of supersonic and hypersonic airframes.

Shock-induced leading-edge separation in hypersonic flows

We investigate the leading-edge separation, also called negligible boundary-layer thickness separation, induced by an impinging shock on a sharp flat plate. The canonical impinging shock wave/boundary-layer interaction configuration consisting of a wedge and a plate wall, placed in hypersonic free stream, is used for the investigations. We first construct a theoretical model for the leading-edge separated flow field (LESF) which predicts separation bubble geometry and surface pressure distribution as a function of three parameters: free-stream Mach number, wedge angle and the reattached flow turning angle. Markedly different predictions of the separated flow field are obtained for an oblique and a near normal reattachment on the plate surface. Experiments in a shock tunnel at a nominal Mach $6$ flow, with an impinging shock generated by a wedge of angle $26.6^{\circ }$ , are used to validate the model. Schlieren flow visualization using a high-speed camera and surface pressure measurements using fast response piezoelectric sensors are the diagnostics employed. For a range of shock impingement locations, the LESFs are observed to be geometrically similar and in good agreement with the LESF model. When the impingement location gets closer to the leading edge, it is observed from experiments that the flow field is no longer geometrically similar, and the separation angle increases as the impingement gets closer to the leading edge beyond the range of similarity. The work thereby offers an elaborate description of the leading-edge separated flow when shock impingement occurs near the plate leading edge.

Evaluation of reduced-order aeroelastic simulations for shock-dominated flows

Aeroelastic predictions of a thin panel in supersonic flow with a shock/boundary layer interaction (SBLI) are evaluated relative to wind tunnel experiments. Measured time histories of the panel displacement and velocity show stationary deflections as well as post-flutter oscillations for a range of shock impingement strengths. The fully coupled, computationally efficient aeroelastic modeling framework is formulated with a nonlinear structural reduced-order model and quasi-steady enriched piston theory aerodynamics. The aeroelastic model predicts the observed shape of the stationary responses with correlation coefficients larger than 0.8 for all SBLI conditions. In regards to dynamic responses, the model captures flutter onset as well as the post-flutter periodic oscillations for attached SBLI. The system’s sensitivity to different initial conditions and shock impingement location is also investigated for attached SBLI. Shock-induced flow separation emerges as the SBLI strength increases. For the strongest shock, the computations predict flutter but do not capture the frequency content of the measured post-flutter response. This discrepancy is tied to the breakdown of the quasi-steady flow assumption in piston theory with increasing flow separation and highlights the need for unsteady flow models.

Supersonic Turbulent Boundary-Layer Separation Control Using a Morphing Surface

Separated flows arising due to shock-wave/turbulent boundary-layer interactions can cause problematic low-frequency unsteadiness with potentially severe structural response. High-fidelity large-eddy simulations are employed to examine surface morphing as a way to reduce the size of the separation region, and thus favorably alter the unsteadiness characteristics. The configuration considers a turbulent Mach 2.7 flow at a Reynolds number of subjected to an impinging shock system with a pressure ratio of , which results in separation and the presence of structurally relevant low-frequency unsteadiness. The control surface, centered about the shock impingement location and extending over the separation region, is allowed to deform under material property-based realizability constraints until an asymptotic state is achieved. The criterion for deformation uses a measure proportional to the directional surface shear stress. At an asymptotic state, the deformed surface reveals a shape consistent with aerostructural optimization and a maximum height of . Control mitigates the sharp initial pressure gradient of the uncontrolled flow to delay and reduce separation extent (by 50%) with diminution of low-frequency content and turbulent kinetic energy. Modal decomposition highlights these effects in the energy content of the prominent modes. Morphing may thus provide a means to adjust the local surface deflection in a manner that reduces some of the problems associated with turbulent separation.

Control of transitional shock wave boundary layer interaction using structurally constrained surface morphing

The potential of surface morphing techniques, including passive shock control bumps (SCB) and active surface morphing, is explored to control transitional shock wave boundary layer interactions (SWBLI). In addition to reducing the size of the separation bubble, a key objective is to mitigate the low-frequency unsteadiness that can cause detrimental structural response. To this end, three-dimensional flow simulations are performed using direct numerical simulations (DNS) at Mach 2 and Reynolds number based on inflow boundary layer thickness Reδin=996. An incident oblique shock of angle σ=35deg and strength p2/p1=1.4 impinges on a laminar boundary layer that evolves from a Blasius profile. The resulting boundary layer separation leads to transition and a Görtler-like instability is observed; the nominally two-dimensional and steady flow becomes three-dimensional and unsteady. The goal of this work is to develop a surface modification method to mitigate separation and low-frequency unsteadiness, which can trigger structural response and flow distortion. An aero-structural solver framework is developed and employed to examine both passive SCB and active surface morphing. To avoid unrealistic structural deformation in the transient or final states, the structural integrity is concurrently monitored so that the intermediate morphing solutions are restricted to achievable elastic deformations. The results indicate that transitional SWBLI can be controlled in this manner, to essentially inhibit transition and thus eliminate the separation and unsteadiness associated with the Görtler-like vortices. The mechanism modulates sharp increases in surface pressure at separation and shock-impingement locations encountered in uncontrolled SWBLI and results a lower specific entropy rise.

Effects of Shock Location on Aeroelastic Stability of Flexible Panel

Focusing on the aeroelastic stability of heated flexible panel in shock-dominated flows with arbitrary shock impingement location condition, a systematic theoretical analysis model is established. The von-Karman large deflection theory is used to account for the geometrical nonlinearity. Local first-order piston theory is employed in the region before and after shock to estimate the aerodynamic pressure. The results show that the shock impingement location has a significant effect on the stability of the flexible panel. The change of the critical dynamic pressure (for which flutter onset occurs) is not a monotonic function of the shock impingement point from one endpoint to the other. Specifically, the nondimensional critical flutter dynamic pressure increases as the shock impingement points move from the left to the right in the middle region, and the value of the critical flutter dynamic pressure varies substantially as the shock is located near the boundary of the panel. The variation is larger for higher shock strength. For a large shock strength, panel can buckle (or diverge) rather than only flutter even without thermal stresses. The sensitivity of the panel dynamics to the shock impingement location may have interesting implications for the case of a moving incident shock.

A parametric study of laminar and transitional oblique shock wave reflections

High-resolution particle image velocimetry measurements were performed on laminar and transitional oblique shock wave reflections for a range of Mach numbers ($M=1.6{-}2.3$), Reynolds numbers ($Re_{x_{sh}}=1.4\times 10^{6}{-}3.5\times 10^{6}$) and flow deflection angles ($\unicode[STIX]{x1D703}=1^{\circ }{-}5^{\circ }$ or $p_{3}/p_{1}=1.11{-}1.64$). The laminar interactions revealed a long, flat and triangular shaped separation bubble. For relatively strong interactions ($p_{3}/p_{1}>1.2$), the bubble grows linearly in the upstream direction with increasing shock strength. Under these conditions, the boundary layer keeps an on average laminar velocity profile up to the shock impingement location, followed by a quick transition and subsequent reattachment of the boundary layer. For weaker interactions ($p_{3}/p_{1}<1.2$), the boundary layer is able to remain laminar further downstream of the bubble, which consequently results in a later reattachment of the boundary layer. The pressure distribution at the interaction onset for all laminar cases shows excellent agreement with the free-interaction theory, therefore supporting its validity even for incipiently separated laminar oblique shock wave reflections.

Shock-tunnel investigations on the evolution and morphology of shock-induced large separation bubbles

ABSTRACTShock-tunnel experiments are carried out to study the strong interaction between an impinging shock wave and boundary layer on a flat plate, accompanied bylargeseparation bubble with a length comparable to the distance of the location of shock impingement from the leading edge of the plate. For nominal freestream Mach numbers ranging from 6 to 8.5, moderate to high total enthalpies of 1.3MJ/kg to 6MJ/kg are simulated in the Indian Institute of Science's hypersonic shock tunnels HST-2 (a conventional Hypersonic Shock Tunnel) and Free Piston Shock Tunnel (FPST) with freestream Reynolds numbers ranging from 4 × 106/m to 0.3 × 106/m. The strong impinging shock is generated by a wedge (or shock generator) at an angle of 30.96° to the freestream. From the time-resolved Schlieren flow visualisations using a high-speed camera and surface pressure measurements on the flat plate using fast response sensors, a statistically steady flow field with a large separation bubble was established within the short test time of the shock tunnels (around 600µs in HST-2 and 300µs in FPST). The role of various parameters on the interaction – Mach number, location of shock impingement and flow total enthalpy – are investigated from the measured separation length and surface pressure distribution. For the nominal Mach number of 8.5, with shock impingement at 100mm from the leading edge, the separation length increased from 60mm to 70mm as the total enthalpy is increased from 1.6MJ/kg to 2.4MJ/kg; whereas it dropped drastically to 30-40mm at 6MJ/kg. This is due to the prominence of real gas effects at higher enthalpies.

Aerodynamic Interference on Finned Slender Body

Aerodynamic interference can occur between high-speed slender bodies when in close proximity. A complex flowfield develops where shock and expansion waves from a generator body impinge upon the adjacent receiver body and modify its aerodynamic characteristics in comparison to the isolated case. The aim of this research is to quantify and understand the multibody interference effects that arise between a finned slender body and a second disturbance generator body. A parametric wind tunnel study was performed in which the effects of the receiver incidence and axial stagger were considered. Computational fluid dynamic simulations showed good agreement with the measurements, and these were used in the interpretation of the experimental results. The overall interference loads for a given multibody configuration were found to be a complex function of the pressure footprints from the compression and expansion waves emanating from the generator body as well as the flow pitch induced by the generator shockwave. These induced interference loads change sign as the shock impingement location moves aft over the receiver and in some cases cause the receiver body to become statically unstable. Overall, the observed interference effects can modify the subsequent body trajectories and may increase the likelihood of a collision.

Colour surface flow visualisation of interfering slender bodies at Mach 3

A study to understand the flow physics produced by two slender bodies in close proximity in high-speed airflow was undertaken. The interference flow field generated by the bodies is dominated by shock and expansion waves, and of particular significance is the complex interaction of the bow-shock wave emanating from the disturbance generator, striking the surface of the disturbance receiver. To gain insight into the shock wave-body interaction, the traditional surface oil flow visualisation technique was extended to include colour, which assists the eye in tracking the streak lines back to their separation and reattachment regions. In addition, the fine particle sizes of the colour pigment produced crisper and more definitive separation lines over the body, in comparison to the traditional monochrome pigments, such as lamp-black or titanium-dioxide. Subsequently, the dried surface pattern was lifted off the body using matte-acetate tape, digitised and then straightened using datum markings along the sting. This allowed the shock impingement location and shock diffraction path over the body to be established quantitatively, which was used to validate numerical simulations. The experimental and computational data showed good agreement for all configurations considered, providing complementary information about the disturbance-induced effects generated in the interference flow field, and provided detailed insight into the near-surface flow topology produced by the shock wave-slender body interaction.

Transitional shock-wave/boundary-layer interactions in hypersonic flow

AbstractStrong interactions of shock waves with boundary layers lead to flow separations and enhanced heat transfer rates. When the approaching boundary layer is hypersonic and transitional the problem is particularly challenging and more reliable data is required in order to assess changes in the flow and the surface heat transfer, and to develop simplified models. The present contribution compares results for transitional interactions on a flat plate at Mach 6 from three different experimental facilities using the same instrumented plate insert. The facilities consist of a Ludwieg tube (RWG), an open-jet wind tunnel (H2K) and a high-enthalpy free-piston-driven reflected shock tunnel (HEG). The experimental measurements include shadowgraph and infrared thermography as well as heat transfer and pressure sensors. Direct numerical simulations (DNS) are carried out to compare with selected experimental flow conditions. The combined approach allows an assessment of the effects of unit Reynolds number, disturbance amplitude, shock impingement location and wall cooling. Measures of intermittency are proposed based on wall heat flux, allowing the peak Stanton number in the reattachment regime to be mapped over a range of intermittency states of the approaching boundary layer, with higher overshoots found for transitional interactions compared with fully turbulent interactions. The transition process is found to develop from second (Mack) mode instabilities superimposed on streamwise streaks.

The effect of Mach number on unstable disturbances in shock/boundary-layer interactions

The effect of Mach number on the growth of unstable disturbances in a boundary layer undergoing a strong interaction with an impinging oblique shock wave is studied by direct numerical simulation and linear stability theory (LST). To reduce the number of independent parameters, test cases are arranged so that both the interaction location Reynolds number (based on the distance from the plate leading edge to the shock impingement location for a corresponding inviscid flow) and the separation bubble length Reynolds number are held fixed. Small-amplitude disturbances are introduced via both white-noise and harmonic forcing and, after verification that the disturbances are convective in nature, linear growth rates are extracted from the simulations for comparison with parallel flow LST and solutions of the parabolized stability equations (PSE). At Mach 2.0, the oblique modes are dominant and consistent results are obtained from simulation and theory. At Mach 4.5 and Mach 6.85, the linear Navier-Stokes results show large reductions in disturbance energy at the point where the shock impinges on the top of the separated shear layer. The most unstable second mode has only weak growth over the bubble region, which instead shows significant growth of streamwise structures. The two higher Mach number cases are not well predicted by parallel flow LST, which gives frequencies and spanwise wavenumbers that are significantly different from the simulations. The PSE approach leads to good qualitative predictions of the dominant frequency and wavenumber at Mach 2.0 and 4.5, but suffers from reduced accuracy in the region immediately after the shock impingement. Three-dimensional Navier-Stokes simulations are used to demonstrate that at finite amplitudes the flow structures undergo a nonlinear breakdown to turbulence. This breakdown is enhanced when the oblique-mode disturbances are supplemented with unstable Mack modes.

Experimental Investigation of a Transonic Aspirated Compressor

Abstract The experimental investigation of a transonic aspirated stage demonstrating the application of boundary layer aspiration to increase stage work is presented. The stage was designed to produce a pressure ratio of 1.6 at a tip speed of 750ft∕s resulting in a stage work coefficient of 0.88. The primary aspiration requirement for the stage is a bleed fraction 0.5% of the inlet mass flow on the rotor and stator suction surfaces. Additional aspiration totaling 2.8% was also used at shock impingement locations and other locations on the hub and casing walls. Detailed rotor and stator flow field measurements, which include time-accurate and ensemble-averaged data, are presented and compared to three-dimensional viscous computational analyses of the stage. The stage achieved a peak pressure ratio of 1.58 and through-flow efficiency of 90% at the design point. In addition, the stage demonstrated good performance with an aspiration lower than the design requirement, and a significant off-design flow range below that predicted by the computational analysis.

Unsteadiness of shock wave/boundary layer interaction in supersonic cascade

An investigation of the passive control of shock wave/boundary layer interaction for reducing the amplitude of the shock oscillation was conducted on the circular arc-wedge (CW) profile cascade in a 220×290 mm transonic compressor cascade wind tunnel. A perforated surface with a cavity beneath it was positioned on the suction surface of the blade at the location of shock impingement. The Schlieren and high-speed photographs for flow over perforated blade are presented and compared with the results for solid blades. With the perforated surface, the high-speed photographs indicated an significant suppression of shock oscillation.

Inviscid parametric analysis of three-dimensional inlet performance

The advantages and design requirements of propulsion/airtranie integration for the high Mach number flight of air-breathing vehicles have led to extensive study of the three-dimensional sidevv all-compression scramjet inlet in recent years. Inlets of this genre afford a relatively simple, generic geometry while producing a highly complex, three-dimensional flowfield dominated by shock/shock and shock/boundary-layer interactions. While the importance of the viscous effects in high-speed inlet interactions is recognized, the present work addresses in a parametric fashion the inviscid effects of leading-edge sweep (between 0-70 deg) and inflow Mach number (between 2-12) on the inlet performance. Two-dimensional oblique shock theory is appropriately modified to account for the three-dimensional effects of leading-edge sweep and is applied throughout the inlet configuration to obtain inviscid shock impingement locations, mass capture, inlet compression, total pressure recovery, and kinetic energy efficiency. Comparison of these results with CFD indicates that the parametric trends are identified by this computationally quick and inexpensive method for preliminary design applications.

Investigation of shock/turbulent boundary-layer bleed interactions

Investigation of shock/turbulent boundary-layer bleed interactions