Compression Ramp Research Articles

Shock boundary layer interactions in a low aspect ratio duct

Experimental data acquired using high resolution two-component particle image velocimetry (PIV) are presented for shock boundary layer interactions (SBLIs) generated by a compression–expansion ramp geometry. The incident oblique shock wave is generated by a sub-boundary layer height ramp inclined 20° to the M∞=2.05 inflow. Results are presented for two different ramp sizes (hramp/δ0=0.56 and hramp/δ0=0.93), and compared to a previously documented hramp/δ0=0.20 case. For each case, mean velocitiy and turbulent statistics for both the SBLI at the foot of the compression ramp and the incident/reflected SBLI on the opposite wall are analyzed. Data are acquired in several streamwise-vertical planes across the span of the low aspect ratio duct in order to document spanwise non-uniformities and confinement effects, with the specific goal of producing a high quality experimental database for CFD validation. The angles of the incident and reflected shock waves become steeper as the side wall is approached, due to the lower velocities within the side wall boundary layer. Mean flow reversal is observed near the spanwise centerline of the duct, but only instantaneous flow reversal is seen closer to the side walls. These spanwise non-uniformities are more prominent for the stronger interactions caused by the larger ramp geometry. For this case there is no nominally two-dimensional region of the flow, and a Mach stem occurs in the core of the flow, causing a significant subsonic wake. The shock excursion length scale relative to the incoming boundary layer thickness, Lex/δ0, is measured for all of the shock features and found to be significantly lower than values reported in the literature for similar flows. Furthermore, Lex/δ0 for the reflected shock does not depend on the strength of the incident shock or the size of the separated zone.

Numerical study on the standing morphology of an oblique detonation wave under the influence of an incoming boundary layer

Abstract The influence of an incoming boundary layer to the standing morphology of an oblique detonation wave (ODW) induced by a compression ramp is numerically studied in this paper. The Spalart-Allmaras (SA) turbulence model is used to perform simulation of detonationboundary- layer interactions. Three different wall conditions are applied to realize control on the boundary-layer separation scales. Accordingly, different standing morphologies of the ODWs are obtained, including smooth ODW (without transverse wave) under no-slip, adiabatic wall condition with large-scale separation, abrupt ODW (with transverse wave) under no-slip, cold wall condition with moderate-scale separation, and bow-shaped detached ODW under slipwall condition without a boundary layer.

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.

Control of mean separation in shock boundary layer interaction using pulsed plasma jets

The current study investigates the use of pulsed plasma jets (spark jets) to reduce the separation induced by shock wave-boundary layer interaction generated by a $$20^{\circ }$$ compression ramp in a Mach 3 flow with a Reynolds number of 5,400, based on the undisturbed boundary layer momentum thickness. Surface oil streak visualization is used in a parametric study to determine the optimum pulsing frequency of the jet, the optimum distance of the jet from the compression corner, and the optimum configuration of the jets. Several 3-jet actuator configurations are tested, including those where the jets are pitched, and pitched and skewed. The jet pulsing frequency is varied between 2 and 4 kHz, corresponding to a Strouhal number based on separation length of 0.012 and 0.023. Particle image velocimetry is used to characterize the effect that the actuators have on the reattached boundary layer profile on the ramp surface. Results show that plasma jets pitched at $$20^{\circ }$$ from the wall, and pulsed at a Strouhal number of 0.018, can reduce the distance between the separation line and the compression ramp corner by up to 40 % and increase the integrated momentum in the downstream reattached boundary layer, albeit with a concomitant increase in the shape factor.

Numerical Simulation on Air Mass Capture Control of Hypersonic Inlet by Laser Energy

Mechanism of hypersonic inlet performance promoting by “virtual cowl” induced by laser energy is introduced, and the physical model of interaction between laser energy and shocks in hypersonic flow field is established. Through comparing results in this paper with the work performed by Macheret et al.(Princeton University) and analyzing effects of laser energy deposited in hypersonic flow, numerical program and the feasibility of model are validated. At Mach 6, with no laser energy addition, air mass capture ratioKmis 71.87% in this paper, whileKmis 73.17% in the reference. Energy addition module can simulate the formation of shocks induced by laser energy, and the interaction process between laser energy and compression ramp shocks properly. Curves of mass capture ratioKmagrees well with that of the reference.Kmis a little higher than data in the reference, but the variable trend is consistent with each other when laser power is 1kW. When laser power is 0.5kW,Kmvaries a little, and the peak value is 77.36%, which only increases 7.64%, while the value is 7.28% in the reference.

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.

Validation and application of an intervertebral disc finite element model utilizing independently constructed tissue-level constitutive formulations that are nonlinear, anisotropic, and time-dependent

Finite element (FE) models are advantageous in the study of intervertebral disc mechanics as the stress–strain distributions can be determined throughout the tissue and the applied loading and material properties can be controlled and modified. However, the complicated nature of the disc presents a challenge in developing an accurate and predictive disc model, which has led to limitations in FE geometry, material constitutive models and properties, and model validation. The objective of this study was to develop a new FE model of the intervertebral disc, to validate the model׳s nonlinear and time-dependent responses without tuning or calibration, and to evaluate the effect of changes in nucleus pulposus (NP), cartilaginous endplate (CEP), and annulus fibrosus (AF) material properties on the disc mechanical response. The new FE disc model utilized an analytically-based geometry. The model was created from the mean shape of human L4/L5 discs, measured from high-resolution 3D MR images and averaged using signed distance functions. Structural hyperelastic constitutive models were used in conjunction with biphasic-swelling theory to obtain material properties from recent tissue tests in confined compression and uniaxial tension. The FE disc model predictions fit within the experimental range (mean±95% confidence interval) of the disc׳s nonlinear response for compressive slow loading ramp, creep, and stress-relaxation simulations. Changes in NP and CEP properties affected the neutral-zone displacement but had little effect on the final stiffness during slow-ramp compression loading. These results highlight the need to validate FE models using the disc׳s full nonlinear response in multiple loading scenarios.

Characterization of the shear layer in a Mach 3 shock/turbulent boundary layer interaction

A detailed characterization of the shear layer in a direct numerical simulation of a Mach 3, separated shock/turbulent boundary layer interaction over a 24° compression ramp is presented. Similarity solutions are identified in the time-averaged shear layer and the growth of the shear layer is found to be approximately linear. In addition to the time averaged spatial organization of the shear layer, the characteristic frequencies and time scales of the large Kelvin- Helmholtz vortices are determined and are found to be consistent with the energized frequency content in the pre-multiplied power spectra of the wall pressure. The three-dimensional nature of the mixing layer vortices are identified using two-dimensional correlation contour plots and a technique for the identification of individual vortices. The large vortices shed from the separated region are found to lie in a plane parallel to the ramp surface, but angled at ±45° to the freestream direction.

Luminescent measurement systems for the investigation of a scramjet inlet-isolator.

Scramjets have become a main focus of study for many researchers, due to their application as propulsive devices in hypersonic flight. This entails a detailed understanding of the fluid mechanics involved to be able to design and operate these engines with maximum efficiency even at their off-design conditions. It is the objective of the present cold-flow investigation to study and analyse experimentally the mechanics of the fluid structures encountered within a generic scramjet inlet at M = 5. Traditionally, researchers have to rely on stream-thrust analysis, which requires the complex setup of a mass flow meter, a force balance and a heat transducer in order to measure inlet-isolator performance. Alternatively, the pitot rake could be positioned at inlet-isolator exit plane, but this method is intrusive to the flow, and the number of pitot tubes is limited by the model size constraint. Thus, this urgent need for a better flow diagnostics method is addressed in this paper. Pressure-sensitive paint (PSP) has been applied to investigate the flow characteristics on the compression ramp, isolator surface and isolator sidewall. Numerous shock-shock interactions, corner and shoulder separation regions, as well as shock trains were captured by the luminescent system. The performance of the scramjet inlet-isolator has been shown to improve when operated in a modest angle of attack.

Effect of leading edge bluntness on the interaction of ramp induced shock wave with laminar boundary layer at hypersonic speed

Present investigations deal with ramp induced shock wave and boundary layer interaction (R-SWBLI) in the presence of leading edge bluntness. A second order accurate finite-volume compressible flow solver is employed to assess the effectiveness of leading edge bluntness in reducing the separation bubble size. Boundary layer edge Mach number and temperature, sonic height, boundary layer thicknesses, skin friction coefficient and pressure difference are considered for this assessment. Present studies reveal the existence of two critical radii of leading edge bluntness associated with R-SWBLI. Increase in separation bubble size has been observed with increase in leading edge radius until the first critical radius called as ‘inversion radius’ which corresponds to maximum extent of separation. Swallowing of the entropy layer by the boundary layer is seen to increase the separation zone size while the separation bubble length decreases when the boundary layer is immersed within the entropy layer. The inversion radius is shown to be associated with equal thicknesses of the entropy and boundary layers. Increase in leading edge radius beyond the second critical radius, called as ‘equivalent radius’, is found to decrease the size of separation zone in comparison with that of the sharp leading edge case. This reduction is attributed mainly to the existence of a wide over pressure region. Physics-driven predictive strategies, to determine the critical radii for given geometry and freestream conditions, are derived successfully from the present numerical simulations. Current studies on a 50mm long plate with 15° compression ramp for Mach number 6 and wall temperature of 300K show that the inversion radius lies between 0.3 and 0. 6mm while the equivalent radius lies between 1 and 1.2mm.

OH PLIF visualization of the UVa supersonic combustion experiment: configuration A

Hydroxyl radical (OH) planar laser-induced fluorescence (PLIF) measurements were performed in the University of Virginia supersonic combustion experiment. The test section was set up in configuration A, which includes a Mach 2 nozzle, combustor, and extender section. Hydrogen fuel was injected through an unswept compression ramp at two different equivalence ratios. Through the translation of the optical system and the use of two separate camera views, the entire optically accessible range of the combustor was imaged. Single-shot, average, and standard deviation images of the OH PLIF signal are presented at several streamwise locations. The results show the development of a highly turbulent flame structure and provide an experimental database to be used for numerical model assessment.

Shock wave-boundary layer interaction on the non-adiabatic ramp surface

The boundary layer separation of the supersonic air flow over the heated compression ramp has been investigated. Particle Image Velocimetry and Shlieren non-invasive visualization methods were used to study the influence of surface heating on the structure of separation zone. The linear dependence of separation length on the temperature ratio was found. The influence of surface heating on separation region instability was studied. It was found experimentally that increase of wall temperature causes intensification of separation region fluctuations.

Ramp Injector Scale Effects on Supersonic Combustion

The combustion field downstream of a 10 deg compression ramp injector has been studied experimentally using wall static pressure measurement, OH planar laser-induced fluorescence, and 2 kHz intensified video filtered at 320 nm for OH emission. Nominal test section entrance conditions were Mach 2, 131 kPa static pressure, and 756 K stagnation temperature. The experiment was equipped with a variable-length inlet duct that facilitated varying the boundary-layer development length for nearly constant injector shock structure conditions. By varying the boundary-layer thickness from 40% of the ramp height to 150% of the ramp height, changes in the combustion flowfield downstream of the injector could be diagnosed. It was found that flame shape changed, the persistence of the vortex cores was reduced, and combustion efficiency rose as the incident boundary layer grew. Unsteadiness within the flameholding region also varied greatly over the range of conditions studied, and largest unsteadiness was observed at intermediate boundary-layer heights.

Elastic, permeability and swelling properties of human intervertebral disc tissues: A benchmark for tissue engineering

The aim of functional tissue engineering is to repair and replace tissues that have a biomechanical function, i.e., connective orthopaedic tissues. To do this, it is necessary to have accurate benchmarks for the elastic, permeability, and swelling (i.e., biphasic-swelling) properties of native tissues. However, in the case of the intervertebral disc, the biphasic-swelling properties of individual tissues reported in the literature exhibit great variation and even span several orders of magnitude. This variation is probably caused by differences in the testing protocols and the constitutive models used to analyze the data. Therefore, the objective of this study was to measure the human lumbar disc annulus fibrosus (AF), nucleus pulposus (NP), and cartilaginous endplates (CEP) biphasic-swelling properties using a consistent experimental protocol and analyses. The testing protocol was composed of a swelling period followed by multiple confined compression ramps. To analyze the confined compression data, the tissues were modeled using a biphasic-swelling model, which augments the standard biphasic model through the addition of a deformation-dependent osmotic pressure term. This model allows considering the swelling deformations and the contribution of osmotic pressure in the analysis of the experimental data. The swelling stretch was not different between the disc regions (AF: 1.28±0.16; NP: 1.73±0.74; CEP: 1.29±0.26), with a total average of 1.42. The aggregate modulus (Ha) of the extra-fibrillar matrix was higher in the CEP (390kPa) compared to the NP (100kPa) or AF (30kPa). The permeability was very different across tissue regions, with the AF permeability (64 E−16m4/Ns) higher than the NP and CEP (~5.5 E−16m4/Ns). Additionally, a normalized time-constant (3000s) for the stress relaxation was similar for all the disc tissues. The properties measured in this study are important as benchmarks for tissue engineering and for modeling the disc's mechanical behavior and transport.

Studies on the Influence of Steady Microactuators on Shock-Wave/Boundary-Layer Interaction

An array of high-momentum microjets are used upstream of a compression corner to control the shock-wave/boundary-layer interaction on a 24 deg unswept compression ramp in a Mach 2 flow. Measurements include schlieren flow visualization and unsteady pressure measurements using fast-response pressure sensors of the interaction region. Results show that the array of microjets issuing in the supersonic crossflow create oblique shocks, which effectively reduce the incoming Mach number at the compression corner. This leads to a modified separation shock of significantly reduced strength. The location of the modified shock is moved upstream by as much as from its mean undisturbed location. The mean pressure distribution on the surface is altered with microjet control leading to a more gradual compression of the incoming flow relative to the separation shock without control. The wall-pressure fluctuations in the interaction region are reduced by approximately 50%, and the flow near the compression corner appears to be energized with control, based on the unsteady surface-pressure measurements. The pressure spectra show that microjet control results in a redistribution of energy on the wall and the ramp surfaces.

충격파 경계층 상호작용에서 난류모델 및 난류점성의 효과

Two compression ramp problems and an impinging shock problem are computed to investigate influence of turbulence models and eddy viscosity on the shock-wave / boundary layer interaction. A Navier-Stokes boundary layer generation code was applied to the generation of inflow boundary conditions. Computational results are validated well with the experimental data and effects of turbulence models are investigated. It is shown that the behavior of turbulence (eddy) viscosity directly affects both the extent of the separation and shock-wave positions over the separation.

Control of oblique shock wave/boundary layer interactions using plasma actuators

Localized arc filament plasma actuators (LAFPAs) are used for shock wave/boundary layer interaction induced separation control in a Mach 2.3 flow. The boundary layer is fully turbulent with a Reynolds number based on the incompressible momentum thickness of 22,000 and shape factor of 1.37, and the impinging shock wave is generated by a 10° compression ramp. The LAFPAs are observed to have significant control authority over the interaction. The main effect is the displacement of the reflected shock and most of the interaction region upstream by approximately one boundary layer thickness (~5 mm). The initial goal of the control was to manipulate the low-frequency (St~0.03) unsteadiness associated with the interaction region. A detailed investigation of the effect of actuator placement, frequency, and duty cycle on the control authority indicates the actuators’ primary control mechanism is not the manipulation of low-frequency unsteadiness. Detailed measurements and analysis indicate that a modification to the boundary layer through heat addition by the actuators is the control mechanism, despite the extremely small power input of the actuators.

Aerodynamic design and experimental modelling of an innovative supersonic three-dimensional air-intake

AbstractAerodynamic design of an innovative supersonic three-dimensional air-intake including a starting device has been carried out. The intake concept is based on the use of an external compression ramp in the form of a caret waverider. In a design flow regime, this ramp generates a plane oblique shock wave lying on its swept leading-edges and a subsequent two-dimensional isentropic compression wave. The intake starting is provided with a device special flaps of which in the throat are opened slightly in order to enlarge it and to form here both longitudinal and cross slots for air bleed and boundary-layer diversion. The same device can also serve for regulation of intake operating regimes depending on the flight velocity. The experimental model intake with a design Mach number MD = 2 has been developed and tested at flow Mach numbers 1·5, 1·75, and 2·0 in the T-313 supersonic wind tunnel based at ITAM. Flow patterns about the intake have been observed, and flow parameters characterising the intake performance have been measured. The experimental results obtained have confirmed the workability of the starting/regulating device; the data on the efficiency of the model intake have demonstrated its good characteristics in terms of the total pressure recovery.

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.