Effect Of Leading Edge Bluntness Research Articles

Effect of Leading-Edge Bluntness on Shockwave–Boundary-Layer Interaction in a Scramjet Intake

An experimental and numerical study was carried out to investigate the effect of leading edge bluntness on the flowfield characteristics of a Mach 6 scramjet inlet. The numerical study was performed using Ansys Fluent 18.0 software. A steady two-dimensional Reynolds-Averaged Navier–Stokes simulation was carried out using k-ε-Re-Normalisation Group model for a sharp edge and leading edge with bluntness of 0.5 and 1 mm (ramp and cowl tip). Significant changes in the flowfield characteristics and parameters were observed for a blunt leading edge compared to the sharp leading edge case. The experimental campaign was carried out in the Mach 6 S1 Hypersonic Ludwieg Tunnel to validate the numerical results, and a good match was observed up to the separation bubble, and deviation in the reattachment is observed. Furthermore, the unsteadiness associated with the shockwave–boundary-layer interaction region inside the isolator was studied for the sharp and blunt leading edge cases. It was observed that changing the leading edge bluntness modifies the flowfield oscillations, separation bubbles, shear layer oscillations, etc.

Effects of Small Leading-Edge Bluntness on High-Speed Boundary-Layer Instabilities on Flat Plates

A numerical investigation of the effect of small leading-edge bluntness on the linear stability regime in high-speed boundary layers for flat plates was carried out using linear stability theory. While it is known that increasing leading-edge bluntness leads to stabilization of the linear instability, it is not completely understood when leading-edge bluntness effects become relevant; in other words, for which leading-edge radii the primary instability regime is affected for flat plates at higher Mach numbers. For the so-called “sharp” geometries, it is assumed that the influence of the nose radius on stability and transition is negligible. For cones, it is generally accepted that nose radii resulting in Reynolds numbers (based on nose radius) smaller than 1000 can be considered sharp. However, no clear guidelines supported by numerical or experimental data exist to conclusively determine what constitutes a sharp leading edge for a flat plate. In this work, bluntness effects are assessed by considering the influence of the leading-edge radius on the linear stability characteristics for high-speed flat-plate boundary layers. It was found from linear stability calculations that surprisingly small leading-edge radii have significant effects on the critical Reynolds number and spatial amplification rates, thus resulting in substantially reduced maximum -factors.

Leading-edge bluntness effects on the hypersonic flow over the double wedge at multiple aft-wedge angles

The present numerical investigation focuses on the leading-edge bluntness effects on the double wedge with varied aft-wedge angles exposed to low enthalpy hypersonic free stream conditions. The bluntness ratio in this study varies, ranging from R/L1 = 0 (sharp leading edge) to R/L1 = 0.577 (maximum allowable bluntness), along with the aft-wedge angle varying between θ2 = 45° and 60°. Noticeably, even a small bluntness ratio can completely change the shock interaction pattern compared to its sharp geometrical counterpart due to a detached leading-edge shock, enlarged separation bubble, and location of various shock waves concerning it. Critical bluntness ratios exist for the low aft-wedge θ2 = 45° angle, but increasing the aft-wedge angle makes the flow field highly unsteady for some bluntness ratios. Nevertheless, these bluntness ratios for such double-wedge configurations are reported using the mean of separation bubble size. Moreover, this work unravels the cause of such unsteadiness for the unsteady flow fields using the spatial-temporal evolution of the wall pressure distribution and fast Fourier transform of the pressure fluctuation signal at the compression corner and supports the deduced observation with the help of energy-based proper orthogonal decomposition. The increased shock–boundary layer interaction strength moves the separation point upstream beyond the junction of cylindrical bluntness and inclined fore-wedge surface, accompanying sudden change in its direction of motion that perturbs the shear layer that set to a self-sustained, highly unsteady flow field.

Experimental and numerical investigation for hypersonic performance of double swept waverider

The planform-customized waverider developed from the osculating-cone method enables the design of a double swept waverider to remedy some deficiencies for the waverider. To analyze the aerodynamic performances, two models of the double swept waveriders, with cusp head and bend head shape, respectively, were fabricated for the wind-tunnel test, and a single swept waverider model was constructed as the compared configuration. A set of experiments at the design Mach number were performed in the hypersonic wind tunnel FD-07 of our institute. Combining computational fluid dynamics (CFD) simulations, the effect of leading-edge bluntness, the aerodynamic forces and longitudinal stability were studied experimentally and numerically. The bluntness of leading-edge area reduced the lift-to-drag ratio (L/D) significantly. The aerodynamic performances obtained from wind-tunnel tests and CFD simulations were very close, marking a difference in L/D of less than 1% in the design state. The data from both wind-tunnel experiments and CFD simulations implied that the double swept waverider maintained “wave-riding” performance in the hypersonic state, and their longitudinal stability improved dramatically compared with the single swept configuration. The results verified the design method of the double swept waverider and its advantages in aerodynamic performance.

Leading-edge bluntness effects on hypersonic three-dimensional flows over a compression ramp

Abstract

Leading-edge bluntness effects in hypervelocity flat plate flow

The combined effects of leading-edge bluntness and high enthalpy are examined for hypersonic flat plate flow. Experimental pressure and heat transfer data are presented for both sharp and blunt leading-edge flat plates. For the sharp leading-edge flows, the data are in agreement with perfect gas theory. For the blunt leading-edge flows, the low and high enthalpy pressure data approach the perfect gas blast wave theory as the flow proceeds downstream, consistent with earlier studies at low enthalpy. Toward the front of the plate, the heat transfer data lie above the corresponding values obtained with the sharp leading-edge configuration. The difference between sharp and blunt leading-edge heat transfer levels appears to be smaller at high enthalpy. This seems to be due to dissociation which occurs in the nose region, thus reducing the shock stand-off distance and increasing the chemical potential enthalpy of the flow. When the presence of dissociated species in the flow is accounted for, the data close to the leading-edge are seen to compare well with perfect gas bluntness–viscous similitudes. Farther downstream, the measured heat transfer values are greater than the perfect gas theoretical predictions for blunt leading-edge flow, and closer to both the corresponding predictions at chemical equilibrium, and sharp leading-edge theory for perfect gas flow.

Probing Real Gas and Leading-Edge Bluntness Effects on Shock Wave Boundary-Layer Interaction at Hypersonic Speeds

AbstractThe present investigations are centered on understanding the discrepancies in shock wave boundary-layer interaction (SWBLI) for perfect and real gas laminar flows. In view of this, the in-h...

Computational Investigation of the Effects of Leading-Edge Bluntness on Drag at Supersonic Speeds

Computational Investigation of the Effects of Leading-Edge Bluntness on Drag at Supersonic Speeds

Experimental Investigation of Leading Edge Bluntness Effects on Hypersonic Tow-dimensional Inlet

The leading edge bluntness effects on inlet boundary layer transition had been studied based on a typical hypersonic two-dimensional inlet. Four leading edge radii(R=0.05mm, R =0.1mm, R =0.2mm, R =0.25mm) had been experimented in FD-07 wind tunnel, including natural transition and artificial transition. The location of boundary layer transition was identified through the corner pressure distribution characteristics and inlet starting, by this method, we had obtained the regular of boundary layer transition location with the leading edge radii. Experimental results showed that, under the wind tunnel condition, the boundary layer transition location moved down stream with increasing the leading edge radius. The inlet would not start for leading edge radius R = 0.25mm and we had designed artificial transition strip based onlinear stability theory(LST) to make the inlet start successfully.

Receptivity to free-stream vorticity of flow past a flat plate with elliptic leading edge

Receptivity of the two-dimensional boundary layer on a flat plate with elliptic leading edge is studied by numerical simulation. Vortical perturbations in the oncoming free stream are considered, impinging on two leading edges with different aspect ratio to identify the effect of bluntness. The relevance of the three vorticity components of natural free-stream turbulence is illuminated by considering axial, vertical and spanwise vorticity separately at different angular frequencies. The boundary layer is most receptive to zero-frequency axial vorticity, triggering a streaky pattern of alternating positive and negative streamwise disturbance velocity. This is in line with earlier numerical studies on non-modal growth of elongated structures in the Blasius boundary layer. We find that the effect of leading-edge bluntness is insignificant for axial free-stream vortices alone. On the other hand, vertical free-stream vorticity is also able to excite non-modal instability in particular at zero and low frequencies. This mechanism relies on the generation of streamwise vorticity through stretching and tilting of the vertical vortex columns at the leading edge and is significantly stronger when the leading edge is blunt. It can thus be concluded that the non-modal boundary-layer response to a free-stream turbulence field with three-dimensional vorticity is enhanced in the presence of a blunt leading edge. At high frequencies of the disturbances the boundary layer becomes receptive to spanwise free-stream vorticity, triggering Tollmien–Schlichting (T-S) modes and receptivity increases with leading-edge bluntness. The receptivity coefficients to free-stream vortices are found to be about 15% of those to sound waves reported in the literature. For the boundary layers and free-stream perturbations considered, the amplitude of the T-S waves remains small compared with the low-frequency streak amplitudes.

Boundary-layer transition at high free-stream disturbance levels – beyond Klebanoff modes

We consider a nominally uniform flow over a semi-infinite flat plate and show how a small slowly modulated (predominantly streamwise) disturbance of the upstream flow is amplified by leading-edge bluntness effects and eventually develops into a small-amplitude but nonlinear spanwise motion far downstream from the edge. This motion is then imposed on the viscous boundary layer at the surface of the plate – causing an order-one change in its profile shape, which can reduce the wall shear to zero and thereby causes the boundary layer to separate. The present study is similar to an earlier steady flow analysis, but the unsteady effects now cause the upstream boundary layer to develop inflectional profiles which can support rapidly growing inviscid instabilities that give rise to transition before the separation can occur.

Hypersonic Boundary-Layer Transition: Application to High-Speed Vehicle Design

This paper reports the boundary-layer transition characteristics of airbreathing hypersonic configurations and their impact on design environments. It discusses the evolution of the National Aerospace Plane configuration from an axisymmetric baseline to a wedgelike configuration where the transition mechanism is dominated by “natural” transition. Thenatural transition is characteristic of the “quiet” environment in freeflight of smooth slender vehicles. This type of transition mechanism gradually evolves spatially in different modes that can be computed analytically and verified experimentally. The paper discusses the effect of leading-edge bluntness, surface wall temperature, and adverse pressure gradient in compression ramp on transition. The effect of freestream Mach number, Reynolds number, and angle of attack are also studied over the range of peak aerodynamic heating conditions of the National Aerospace Plane environment. The transition behavior was investigated using e-type calculations based on a linear stability code known as the eMalik code.

Leading-edge bluntness effects on aerodynamic heating and drag of power law body in low-density hypersonic flow

A numerical study is reported on power law shaped leading edges situated in a rarefied hypersonic flow. The sensitivity of the heat flux and drag coefficient to shape variations of such leading edges is calculated by using a Direct Simulation Monte Carlo method. Calculations show that the stagnation point heating on power law leading edges with finite radius of curvature follows the same relation for classical blunt body in continuum flow; it scales inversely with the square root of the curvature radius at the nose. Furthermore, for those leading edges with zero or infinity radii of curvature, the heat transfer behavior is in surprising agreement with that for classical blunt body far from the nose of the leading edge.

Correlation of laminar-turbulent transition data over flat plates in supersonic/hypersonic flow including leading edge bluntness effects

The paper discusses flat plate boundary layer transition in supersonic/hypersonic flow conditions. Examination of experimental infrared thermography data illustrates the importance of the leading edge thickness and (non-) uniformity to the transition process. Such observations have triggered the collection of a wide range of experimental data on supersonic/hypersonic flat plate boundary layer transition, and a number of attempts to correlate this data with characteristic parameters including leading edge thickness. Results indicate a strong dependence of the relevant transition parameters on the pressure field in the transition region, as this is determined by the combined effects of leading edge thickness and boundary layer growth/viscous interaction, and particularly on the relative importance of the two effects. In fact, two distinct correlation zones are established, depending on whether the pressure distribution at the onset of transition is dominated by leading edge bluntness effects or by boundary layer growth and viscous interaction, thus limiting the observed data scatter to reasonable levels.

Leading-edge bluntness effects in high enthalpy, hypersonic compression corner flow

An experimental study of the combined effects of leading-edge bluntness and real gas behavior on shock wave/boundary-layer interaction has been performed. Pressure and heat transfer distributions have been measured over a compression corner for a range of corner angles, including the datum case of flat plate flow. On the flat plate and upstream of the corner, the pressure and heat transfer for the blunt leading-edge configuration were found to lie above the corresponding sharp leading values. On the ramp, there was a considerable reduction in the pressure and heat transfer when the leading edge was blunt. Also, the extent of the interaction was seen to be smaller with the blunt leading edge. These results are similar to those from perfect gas studies. The differences between the sharp and blunt leading-edge data appear to be less pronounced at the higher enthalpy. This is thought likely to be because of the reduced shock standoff, which occurs as a result of dissociation. A comparison between the heat transfer data on the ramp face and predictions from the generalized reference enthalpy theory was seen to be reasonable. The upstream influence and plateau pressure were found to be in fair agreement with data from low enthalpy experiments. A C CD c d h hr L lu M n Pr p qw Rex

Hypersonic control flap effectiveness

SummaryThe effects of flap deflection, configuration incidence and leading edge bluntness on boundary layer separation and transition have been studied experimentally. A quasi two-dimensional flat plate equipped with a full span trailing edge control flap has been employed for these tests.The studies have been carried out in a hypersonic gun tunnel at M∞ = 8·2 and Re∞/cm = 9·0 × 104. The flap deflection angles studied were in the range 0° ≤ β ≤ 30°. The incidence range was from zero to α = 10° (positive α is nose down). Leading edge bluntness effects were simulated by the introduction of a hemi-cylindrical leading edge.The flow structure was studied using high speed Schlieren photography as well as surface pressure and heat transfer measurements. Liquid crystals were employed to study the threedimensionality of the flow structure for selected configurations. Analytical theories have been developed to estimate the flap pressure and heat transfer levels for the sharp and blunt configurations. These are compared against experimental measurements.

Distortion of a flat-plate boundary layer by free-stream vorticity normal to the plate

We consider a nominally uniform flow over a semi-infinite flat plate. Our analysis shows how a small streamwise disturbance in the otherwise uniform flow ahead of the plate is amplified by leading-edge bluntness effects and eventually leads to a small-amplitude but nonlinear spanwise motion far downstream from the leading edge of the plate. This spanwise motion is then imposed on the viscous boundary-layer flow at the surface of the plate – causing an order-one change in its profile shape. This ultimately reduces the wall shear stress to zero – causing the boundary layer to undergo a localized separation, which may be characterized as a kind of bursting phenomenon that could be related to the turbulent bursts observed in some flat-plate boundary-layer experiments.

Boundary-layer displacement and leading-edge bluntness effects on attached and separated laminar boundary layers in a compression corner. II - Experimental study

Boundary-layer displacement and leading-edge bluntness effects on attached and separated laminar boundary layers in a compression corner. II - Experimental study

Boundary-layer displacement and leading edge bluntness effects on attached and separated laminar boundary layers in a compression corner. I - Theoretical study

This paper describes a theoretical analysis of highly cooled attached and separated regions of shock wave-laminar boundary-layer interaction in the presence of strong streamwise pressure gradients generated by boundary-layer displacement effects at the leading edge. This method is an extension of an earlier analysis by Holden1'12 to conditions where the inviscid flow cannot be described by simple isentropic flow relationships, and where the boundarylayer upstream of the main interaction is subjected to a strong pressure gradient. The analysis is compared with measurements described in Part II of the study. For strong leading edge displacement effects (%z, > 1), the analysis predicts that highly cooled boundary layers in hypersonic flow will be supercritical; a supercritical-subcritical jump is therefore required to join the solution to the subcritical viscous layer at separation. An examination of the experimental measurements indicates that the supercritical-subcritical jump does not reflect a sudden and basic change in the flow mechanics of separation, but is an approximation necessary because the conventional boundary-layer equations cannot adequately describe the viscous interaction process leading to separation. For some high Mach number, low Reynolds number conditions, we were unable to obtain a unique solution, without recourse to experimental data, by locating a critical point in the throat region of the flow. As in the separated region, there is serious question whether the conventional boundary-layer equations can be used to adequately describe the mechanism of boundary-layer reattachment in these flows.

Mesures de flux thermiques dans un ecoulement hypersonique de gaz rarefie

In rarefied gas flow, measurements of low heat rates throughout the surface of models require high sensitivity heat probes. Presented results have been obtained by means of a thin film surface thermometer; an analog device enables the heat flux to be recorded directly. Heat transfer results are concerning flat plates and Delta wings located with angle of attack in a strong interaction regime at Mach 8; experimental data are compared with theoretical curves taking into account boundary-layer displacement, incidence and leading-edge bluntness effects.