Primary Separation Line Research Articles

On the synchronisation of three-dimensional shock layer and laminar separation bubble instabilities in hypersonic flow over a double wedge

Linear three-dimensional instability is studied in the shock layer and the laminar separation bubble (LSB) induced by shock-wave/boundary-layer interactions in a Mach 7 flow of nitrogen over a double wedge with a$30^{\circ }\text {--}55^{\circ }$cross-sectional profile. At a free-stream unit Reynolds number$Re=5.2\times 10^{4}\,{\rm m}^{-1}$this flow exhibits rarefaction effects and has shock thicknesses comparable to the thickness of the boundary layer at separation. Flow features have been fully resolved using a high-fidelity massively parallel implementation of the direct simulation Monte Carlo method that captures the flow evolution from the inception of three-dimensionality, through linear growth of instabilities, to the early stages of nonlinear saturation. It is shown that the LSB sustains self-excited, small-amplitude perturbations that originate past the primary separation line and lead to spanwise-periodic wall striations inside the bubble and downstream of the primary reattachment line, as known from earlier experiments, simulations and instability analyses. A spanwise-periodic instability, synchronised with that in the separation zone, is identified herein for the first time, which exists in the internal structure of the separation and detached shock layers, and manifests itself as spanwise-periodic cats-eyes patterns in the global mode amplitude functions. The growth rate and the spanwise-periodicity length of linear disturbances in the shock layers and the LSB are found to be identical. Linear amplification of the most unstable three-dimensional flow perturbations leads to synchronised low-frequency unsteadiness of the triple point, with a Strouhal number of$St\approx 0.028$.

Preliminary Study on Aerodynamic Control of High-Angle-of-Attack Slender Body Using Blowing from Penetrating Flow Channels

The objective of this study is to experimentally verify a new aerodynamic control concept of a high-angle-of-attack slender body. In the concept, penetrating flow channels are installed to the apex of the slender body. The blowing or suction is generated at the channel exits in response to the surface pressure distribution. First, the effects of the flow channels on the aerodynamic characteristics are experimentally investigated in a low-speed wind tunnel. The result shows the Suction-Blowing type channel is the most effective because its control effect does not reduce even in higher mainstream flow velocity. The peak value of the side force and yawing moment can be reduced by up to 64% and 49%, respectively. In addition, visualization of the surface flow pattern by the oil flow method shows that the Suction-Blowing type channel makes not only the primary separation line on the body side but also the secondary separation line on the body back become symmetric.

Experimental flow study over a blunt-nosed axisymmetric body at incidence

The flowfield over a blunt-nosed cylinder was examined experimentally at a low subsonic speed for Re=1.88×10 5 and angles of attack up to 40°. Velocity measurements were carried out (employing a seven-hole Pitot tube) as well as wall static pressure and wall shear-stress measurements. Surface flow visualization was applied using liquid crystals and a mixture of oil–TiO 2. For all the examined cases no flow asymmetries were found. For high angles of attack (20° and above) a separation “bubble” appears at the leeside of the nose area (streamwise flow separation). The basic feature of the circumferential pressure distribution at the after body area for these angles of attack is a plateau close to the suction peak and a fast recovery next to it. One streamwise vortex on each side of the symmetry plane is formed as well as a separation bubble about 90° far from this plane, where the cross-flow primary separation line is located. Each cross-flow primary separation line starts at the leeside nose area and moves towards the windward side along the cylindrical after body. The space between the two primary separation lines close to the wall is characterized by high flow fluctuations on the leeside, compared to the low fluctuations of the windward side.

Prediction of Separation Angles Induced by Sharp Fins with Attack Angle

Introduction G LANCING shock-wave/turbulent-boundarylayer interaction, in which an oblique shock wave glances across a boundary layer along an adjacent wall (Fig. 1), constitutes one of the most important phenomena of three-dimensional interference. The primary separation line is one of the most easily discerned features of the interactionfootprintobtainableby surface  ow visualization.In the interaction induced by a sharp Ž n or a semicone, the separation line is usuallystraightexcept the “inceptionregion.”Hence once the angle of the separation line ( s : deŽ ned as the angle formed by the separation line and the incoming freestreamdirection) is speciŽ ed, one can grasp the extent of interaction region. Koide et al.1 previouslyproposedan empirical predictionmethod for theprimary separationanglescausedby a seriesof unsweptsharp Ž ns (USF) with attack angle® (Fig. 1a), semicones (SC), and sharp triangle Ž ns (STF) at zero ® (Figs. 1b and 1c). In the prediction, only the inviscid shock-wave characteristics such as shock angle and pressure rise across the shock were employed. (Inviscid means the imaginary condition that would exist if no boundary layer were presented on the wall.) In this Note, the prediction method is extended to include swept sharp Ž n (SSF) with ® (Fig. 1d) and STF with ®. Using the extended method, the separation angles can be predicted for almost all types of sharp shock generators with various angles of attack ®, sweep angles , half-apex angles under a wide range of supersonicMach numbers.

菱形三角フィンによる衝撃波と境界層の干渉

This paper presents an experimental and numerical study of shock wave/turbulent boundary layer interaction induced by rhombic delta fins placed on the side wall of the test section. The experiments were carried out in 80×100mmmm2 supersonic wind tunnel at free stream Mach number of 2.50. Corresponding free stream unit Reynolds number was 3.74×107/m. Five models were used in this study. The fin half apex angles were 12°, 18°, 24° at 30° sweepback angle and for the fin sweepback angles 15°, 45° at 18° half apex angle. Schlieren photographs, oil flow measurements and wall surface static pressure measurements were made. It was found that the upstream influence line runs parallel to the primary separation line and obeys the Reynolds number scaling law. Furthermore, the primary separation angle βs1 can be accounted for by referencing βs1 and β0 to Mach angle μ∞.

Reynolds Number Effects on Low-Speed Aerodynamics of a Hypersonic Configuration

A Collaborative Research Center project is focused on fundamental design aspects of a hypersonic research conŽ guration that consists of a lifting body with a delta planform of aspect ratio 1.1 and rounded leading edges. The aerodynamics of the conŽ guration at Mach numbers below 0.3 are described. The conclusions are based on tests in different subsonicwind tunnels. Pressure distributionmeasurementswere carried out at aReynoldsnumber range of 3 £ 106 <– Re <– 40 £ 106 using different wind-tunnel models. Above an angle of attack of 8 deg, the  ow at the suction side changes and a concentrated vortex system (primary and secondary vortices) begins to form, causing local pressure minima in the pressure distribution.On rounded leading edges, the primary separation line is not a priori Ž xed and the vortex system has less intensity than for delta wings with sharp leading edges. The present investigations show that, in this case, the location of the separation line and the strength and location of the vortex system depend on the Reynolds number. Pressure distribution measurements at higher Reynolds numbers with a largemodel in theGerman–DutchWind Tunnel supported these observations. In particular, the high spatial resolution allowed a good physical interpretation of the  owŽ eld. Knowledge of the  ow processes was improved by surface oil- ow patterns, the laser-light-sheet technique, and particle image velocimetry.

Control of 3D turbulent boundary layer separation caused by a wing–body junction

Abstract In three-dimensional flows, boundary layer separation leads to the formation of vortical structures formed by the rolling up of the viscous flow sheet previously confined in a thin layer attached to the wall. Several characteristics of the vortex flow development have negative effects; for example, it may be responsible for unsteadiness or generating unwanted vibration or noise. There are several ways to alter the properties of the vortex or to eliminate it as an organized structure. The objective of this paper is to present two different vortex control approaches. The first solution involves a modification of the pressure field induced by the obstacle by altering its geometry with a strake. The second solution, which can be called active control, consists of suction through an aperture located near the separation line. In this study, we have considered the incompressible three-dimensional boundary layer separation induced by an obstacle mounted perpendicular to a flat plate. The tests were conducted in the F2 subsonic wind tunnel of the ONERA Fauga-Mauzac center. The test section has a cross sectional area of 1.8 × 1.4 m2 and the upstream velocity of the flow was 50 m/s. The surface flow properties have been characterized by using a viscous coating which allows the visualization of the skin friction line patterns. The essential part of the study has been devoted to Laser Doppler Velocimetry (LDV) measurements. The plane of symmetric flow ahead of the leading edge has been qualified by using 3D laser velocimetry in the forward scatter mode. Results show that the geometrical extent of the separation region is reduced and the primary separation line is much closer to the model's leading edge when considering the configuration with a strake. Using an active control system, with suction through a hole located in the plane of symmetry in front of the obstacle's leading edge, the vortex size is reduced and its location displaced towards the origin of the obstacle.

Supersonic separation with obstructions

The complex problem of the supersonic flow along a flat plate as it interacts with the blunt obstructions has recieved considerable attention. The presence of a blunt obstruction in supersonic flow causes the formation of a bow shock wave ahead of the obstruction. The adverse pressure gradient resulting from the bow shock wave propagates upstream in the subsonic flow within the boundary layer and results in the separation of the boundary layer and formation of vortices ahead of the obstruction. These vortices scavenge the boundary layers leading to increased pressures and heat transfer loads. Our attempts are made to correlate the separation distance S, which is defined as the distance from the leading edge of the obstacle to the primary separation line, ahead of the obstacle to the primary separation line, ahead of the obstacle to the height H of the obstacle.

鈍頭突起物に誘起された衝撃波‐乱流境界層干渉場の実験的研究

Three-dimensional shock wave/turbulent boundary layer interaction region induced by a blunt protuberance is experimentally investigated. Experiments are performed under the testing condition of freestream Mach number of 2, 3 and 4 with almost adiabatic wall condition. The primary separation lines locate almost same position irrespective of freestream Mach number, and freestream Mach number independence on the primary separation region is revealed. However, detailed flow structures for each Mach number are different. Pressure begins to rise at slightly ahead of the primary separation line where stream lines begin to deflect outward. As freestream Mach number is increased, the maximum pressure becomes higher but locations, where the pressure begins to rise and reaches to its maximum level, exist at almost same locations respectively regardless of freestream Mach numbers. The oscillating separated shock waves are observed in the interaction region by instantaneous shadow graph technique and significant unsteady flow structure is captured.

A214 三次元衝撃波−乱流境界層干渉における二次剥離現象の可視化

Secondary separation phenomena in three-dimensional shock wave/turbulent boundary layer interaction regions induced by sharp fin and blunt fin have been investigated. A color oil source method, in which various color oil are oozed through many holes on the interaction surface, are applied for details visualization of separated region as well as conventional oil flow technique.For sharp fin separated region increases as attach angle is increased. Also at higher attack angle a separated region is divided into two parts. The outer part is located next to the primary separation line and oil flow lines are almost parallel to the primary separation line. The inner part is located between the outer part and fin and 'S' type oil flow lines are quite significant.For blunt fin separated region at freestream Mach number of 2, 3 and 4 are visualized. The results show separated flow patterns changes between Mach number 3, and 4. Also at Mach number of 4, secondary separation region become quite obvious by using the color oil source method.