Transition In Hypersonic Flows Research Articles

Transport Equation Transition Modeling in Cerans For Hypersonic Flows

Scope of present work covers implementation to validation of the Langtry-Menter two equation γ-Reθt Local Correlation based Transition Model (LCTM) in CERANS code for modeling subsonic to hypersonic flow transition. The LCTM is coupled with SST k-ω turbulence model. The governing equations of γ-Reθt model is discretized in finite volume framework similar to the RANS model and implicitized using the point Jacobi method. Transition correlations based on freestream as well as local turbulence intensity and critical momentum thickness Reynolds number were integrated with the model, and the code is validated for several standard transition test cases involving low subsonic to hypersonic high enthalpy flows covering a wide range of turbulent intensities.

The stabilizing effect of grooves on Görtler instability-induced boundary layer transition in hypersonic flow

Görtler vortex-induced hypersonic boundary layer transition controlled by grooves is investigated using direct numerical simulations and spatial bi-global stability analysis. In the simulations, Görtler vortices are excited by wall steady blowing and suction with spanwise wavelengths of 3 mm. It is found that when the wall is covered with grooves, the Görtler streaks keep more regular even at the end of the model. In addition, the skin friction coefficient is reduced efficiently. Furthermore, the wall-normal and spanwise velocity shear are both reduced, suppressing growths of secondary instabilities. In conclusion, grooves can delay Görtler vortex-induced transition by modifying the Görtler streaks structure and instability, which would shed light on hypersonic boundary layer transition control.

Electrically-heated flat plate testing in a free-piston driven shock tunnel

Reflected shock tunnels can reproduce realistic pressures, speeds, and temperatures of atmospheric hypersonic flight, however, their short test time does not lead to the wall temperatures experienced in flight. This work outlines the design challenges and testing methods of a hot-wall flat plate model. The test model aimed to reliably and rapidly achieve flight-realistic wall temperatures in the T4 free-piston driven shock tunnel at the University of Queensland. This was achieved using a rectangular graphite plate that was electrically heated to temperatures above 800 K. Infrared thermography was employed to measure the surface temperature immediately before a test. Visualised flow-fields from the high-speed schlieren showed that the heated wall conditions made a significant difference in the growth of the boundary layer, and the preliminary observation suggests the heated wall possibly accelerated the boundary layer transition in hypersonic flows.

Wavelet analysis ultra-thin metasurface for hypersonic flow control

Metasurfaces are one of the new frontier of material science and engineering, finding widespread applications in a number of fields, including acoustics and aeroacoustics. In the present paper we propose a novel metasurface specifically fabricated for the hypersonic flow transition control, together with a new strategy for metasurface characterization in the ultrasonic regime. Instead of a conventional porous layer, the metasurface here presented consists in a flat plate with a set of regularly distributed sharp slots. We experimentally observed that such a geometry significantly reduces the wall reflection coefficient, which is known to play a fundamental role in the boundary layer transition phenomenon. Numerical simulations led us to interpret the incident wave scattering as the underlying mechanism related to the observed reflection coefficient reduction. The metasurface characterization has been carried out by comparing the conventional reflection coefficient in the Fourier domain with an innovative wavelet transform-based strategy. As an overall result, the surface geometry here proposed has been shown to offer a twofold advantage: i) it is more effective in ultrasonic incident wave control with respect to conventional porous layers, ii) it needs less manufacturing time, a primary requirement of this type of technology. More interesting, we have put in evidence how a multiresolution approach, like that here proposed, can be highly promising as characterization tool for further metasurfaces with a more complex and multi-scale geometry, being wavelets able to capture the multiscale behaviour of the reflected wave, overcoming the well-known limits of Fourier-based strategies of data analysis.

Impedance-Near-Zero Acoustic Metasurface for Hypersonic Boundary-Layer Flow Stabilization

The laminar-turbulent transition in hypersonic flows is associated with the acoustic-wave-like Mack second mode. Understanding the physics and suppression of this mode is not only of basic scientific interest, but is also essential to future high-speed transportation. This study reveals that, counterintuitively, an almost nondissipative boundary with near-zero surface acoustic impedance can suppress this mode, and performs even better than conventional dissipation-based ultrasonic absorptive coatings. A practical realization, based on an acoustic metasurface, is furthermore presented, pointing toward full control of the hypersonic boundary-layer transition. Full speed ahead!

Transition Prediction in Hypersonic Boundary Layers Using Receptivity and Freestream Spectra

Boundary-layer transition in hypersonic flows over a straight cone can be predicted using measured freestream spectra, receptivity, and threshold values for the wall-pressure fluctuations at the transition onset points. Simulations are performed for hypersonic boundary-layer flows over a 7 deg half-angle straight cone with varying bluntness at a freestream Mach number of 10. The steady and the unsteady flowfields are obtained by solving the two-dimensional Navier–Stokes equations in axisymmetric coordinates using a fifth-order-accurate weighted essentially nonoscillatory scheme for space discretization and using a third-order total-variation-diminishing Runge–Kutta scheme for time integration. The calculated factors at the transition onset location increase gradually with increasing unit Reynolds numbers for flow over a sharp cone and remain almost the same for flow over a blunt cone. The receptivity coefficient increases slightly with increasing unit Reynolds numbers. They are on the order of four for a sharp cone and are on the order of one for a blunt cone. The location of transition onset predicted from the simulation including the freestream spectrum, receptivity, and the linear and the weakly nonlinear evolutions yields a solution close to the measured onset location for the sharp cone. The simulations overpredict transition onset by about 20% for the blunt cone.

Blunt-body paradox and transient growth on a hypersonic spherical forebody

The laminar-turbulent transition in hypersonic flow on a blunt hemisphere is studied with Navier-Stokes and parabolized stability equations. Maximum transient kinetic energy growth is found close to the sonic point, which could explain the transition observed in wind tunnel and flight experiments.

A laminar kinetic energy transition model appropriate for hypersonic flow heat transfer

A laminar kinetic energy transition model appropriate for hypersonic flow heat transfer is proposed in this paper. Based on the laminar kinetic energy transition model at low speed, three rectifications are made to make it possess the ability to predict hypersonic flow transition. Firstly, turbulence length scale is revised to consider boundary layer information directly and reflect the kinematic wall effect correctly at hypersonic speed. Secondly, large-scale eddy viscosity is reconstructed to take hypersonic unstable modes into account to describe the procedure of flow instability. Additionally, for the sake of a better results of the full-turbulence simulation and to be compatible with the revised transition model, the eddy viscosity in sheer stress transport model is adopted to substitute for small-scale eddy viscosity. The model has been applied to a number of relevant test cases, including flat plate boundary layers and boundary layer transition over a blunt cone at different Reynolds numbers. The test cases demonstrate the capacity of the model to reproduce transition flow behavior with a reasonable degree of accuracy and reflect the effect of Reynolds number successfully. While it still needs further investigation to capture the overshoot of heat transfer and take more factors showing significant influence on transition into consideration.

Физические механизмы образования продольных вихрей, появления зон высоких тепловых потоков и раннего перехода в гиперзвуковом течении около треугольного крыла с затупленными передними кромками

Физические механизмы образования продольных вихрей, появления зон высоких тепловых потоков и раннего перехода в гиперзвуковом течении около треугольного крыла с затупленными передними кромками

Experiments on the effect of laminar–turbulent transition on the SWBLI in H2K at Mach 6

This paper presents the results of the experiments performed in the hypersonic wind tunnel H2K in the framework of the ESA technology research project “laminar to turbulent transition in hypersonic flows”. The investigations include the free boundary-layer transition on a flat plate as well as the influence of a shock wave–boundary layer interaction on the transition. The shock is created by a wedge with a small angle of attack resulting in a moderate shock intensity. The experiments were performed at Mach 6.0, at three different unit Reynolds numbers and with a translational displacement of the shock generator. Besides the optical methods—Schlieren photography and infrared thermography—several intrusive sensors were used. High-speed measurements were carried out using PCB and atomic layer thermo pile sensors. Kulite sensors were used for low- and mid-speed pressure measurements. The data analysis includes the comparison of the absolute values, the frequency spectra and wavelets and their distributions in time and space.

Improved-Delayed-Detached-Eddy Simulation of cavity-induced transition in hypersonic boundary layer

Hypersonic flow transition from laminar to turbulent due to the surface irregularities, like local cavities, can greatly affect the surface heating and skin friction. In this work, the hypersonic flows over a three-dimensional rectangular cavity with length-to-width-to-depth ratio, L:W:D, of 19.9:3.57:1 at two angles of attack (AoA) were numerically studied with Improved-Delayed-Detached-Eddy Simulation (IDDES) method to highlight the mechanism of transition triggered by the cavity. The present approach was firstly applied to the transonic flow over M219 rectangular cavity. The results, including the fluctuating pressure and frequency, agreed with experiment well. In the hypersonic case at Mach number about 9.6 the cavity is seen as “open” at AoA of −10° but “closed” at AoA of −15° unconventional to the two-dimensional cavity case where the flow always exhibits closed cavity feature when the length-to-depth ratio L/D is larger than 14. For the open cavity flow, the shear layer is basically steady and the flow maintains laminar. For the closed cavity case, the external flow goes into the cavity and impinges on the bottom floor. High intensity streamwise vortices, impingement shock and exit shock are observed causing breakdown of these vortices triggering rapid flow transition.

Laminar-Turbulent Transition in a Mach 8 Elliptic Cone Flow

To examine the influence of crossflow on transition in a hypersonic boundary layer, an experimental and computational study was made of a Mach 8 flow over a sharp-nosed elliptic cone of 2:1 cross section. Extensive flow visualization was carried out experimentally, including schlieren photography and surface oil-flow visualization. Mean pressure and heat flux were measured at the wall of the wind-tunnel model. For comparison with the experiments, a computation of the basic state flow was carried out using the parabolized Navier-Stokes equations. The basic state laminar flow was found to be relatively complex, with highly skewed velocity profiles and a pair of symmetric vortices on the top centerline of the cone. Transition in the elliptic cone flow was found to be significantly different than transition in hypersonic flow over planar and axisymmetric configurations, which is driven by the second mode instability. In the elliptic cone flow the transition front was nonuniform, with early transition near the top centerline and delayed transition near the leading edge. Rope waves were observed on both the leading edge and top centerline; unusual elongated streaks were observed on the top centerline

Smooth leading edge transition in hypersonic flow

The boundary layer transition along the attachment line of a smooth swept circular cylinder in hypersonic flow is investigated in a blowdown wind tunnel. A wide range of spanwise Mach numbers Me (3.28 to 6.78) is covered with the help of different models at several sweep angles (60°?Λ?80°). The transition is indirectly detected by means of heat flux measurements. The influence of the wall to stagnation temperature ratio is investigated by cooling the model with liquid nitrogen.

Influence of roughness on attachment line boundary layer transition in hypersonic flow

Attachment line boundary layer transition on swept cylinders is studied in a low enthalpy hypersonic wind tunnel at M ∞=7.14. Sweep angles of 60° and 70° are used and transition is detected by means of heat flux measurements. The influence on attachment line transition of single 2D-roughness elements, in the form of tripwires or slots, as well as 3D obstacles is determined and the results are analyzed with respect to Poll’s criterion.

A132 極超音速流中での物体表面温度の測定法

A new type of measurement for surface temperature and heat transfer in hypersonic flow with use of temperature-sensitive paint by the fluorescent phosphor, which is excited by ultraviolet rays, is presented.At first, the relation between temperature and fluorescent intensity is calibrated in static condition.Then, the calibration experiment is applied to a wedge and a cone model in hypersonic flow of Mach 7, and accuracy of this measurement system is evaluated. Simultaneously, surface temperature is visualized by the fluorescence pattern.Solving an unsteady, one-dimensional heat conduction equation, the distribution of heat flux is obtained from the surface temperature distribution, and the boundary layer transition in hypersonic flow is detected and visualized using the cone model with tripping wire.

Reply by Author to K. F. Stetson

Stetson, K. F., Nosetip Bluntness Effects on Cone Frustum Boundary-Layer Transition in Hypersonic Flow, AIAA Paper 831763, July 1983. Reshotko, E. and Khan, M. M. S., ''Stability of the Boundary Layer on a Blunted Plate in Supersonic Flow, presented at IUTAM Symposium on Laminar-Turbulent Transition, Stuttgart, FRG, Sept. 1979. Wright, R. L. and Zoby, E. V., Flight Boundary-Layer Transition Measurements on a Slender Cone at Mach 20, AIAA Paper 77719, June 1977. Stetson, K. F., Thompson, E. R., Donaldson, J. C., and Siler, L. G., Laminar Boundary-Layer Stability Experiments on a Cone at Mach 8, Part 2: Blunt Cone, AIAA Paper 84-0006, Jan. 1984. Stetson, K. F., Thompson, E. R., Donaldson, J. C., and Siler, L. G., Laminar Boundary-Layer Stability Experiments on a Cone at Mach 8, Part 1: Sharp Cone, AIAA Paper 83-1761, July 1983. via the change in the boundary-layer edge conditions through the ''entropy wake generated by the curved bow shock. This change of inviscid flow characteristics is well predicted by the embedded Newtonian theory,' as has been demonstrated recently. Changes in wall temperature and unit Reynolds number do, of course, affect the viscous flow characteristics for both the sharp and the blunted cone. However, the effect of a geometric change from a sharp to a blunted cone is still inviscid in nature. Keeping the above in mind, a careful reading of Ref. 1 will provide all of the answers to the questions raised by Mr. Stetson, keeping in mind that the ratio dN/dB tells whether the conic frustum is long enough to experience the inviscid flow effects imprinted on the boundary layer by the entropy wake.