Wing sweep effects on laminar separated flows

The paper presents Navier-Stokes turbulent simulations of gust responses of a rigid airfoil and of a rigid wing configuration about high mean angles of attack, close to stall, in Reynolds numbers in the order of one million. Responses show that flow separation associated with high angles of attack significantly affect the lift build-up dynamics. Airfoil sharp-edge gust response about high mean angle of attack does not resemble the Kussner function. It involves an initial sharp rise of the lift, which overshoots the steady state value. The maximum lift force that develop in response to a one-minus-cosine gust about high mean angle of attack can be either higher or lower than the response about zero mean angle of attack, depending on the gust gradient length. For short gusts, a dynamic stall effect is observed, where the maximum lift response is higher than that computed about small mean angles of attack, while for longer gusts the maximum lift response is smaller. For the rigid HALE wing case, the lift that develop in response to a one-minus-cosine gust about high mean angle of attack is lower than the response about low mean angles of attack. Since the gust response about high angles of attack does not scale linearly, convolution with sharp-edge gust response does not serve well for prediction of the gust response, and full simulation is required for accurate estimation of the gust loads.

The flow around two infinite swept wings and an infinite swept cylinder model is computed using a Reynolds averaged Navier-Stokes method coupled to a boundary-layer and transition prediction method based on the e(N) approach. The configurations were wind-tunnel tested, including transition location measurements. In all test cases, transition is provoked purely by crossflow instabilities. The infinite swept wing experiments represent critical test cases because they contradict two assumptions made in the conventional e(N) approach, where the limiting cross-flow N factor is supposed to be independent of the surface curvature and the quality of the wing surface finish. It will be shown that the extended version of the e(N)- method can resolve this problem and produce acceptable results. The infinite swept cylinder model tested at high sweep angles between 53 and 71 deg is excellently suited to demonstrate that the e(N)-method is capable of describing the leading-edge contamination problem. For sweep angles between 53 and 58 deg, a relative constant limiting crossflow N factor exists, whereas for higher sweep angles the crossflow N factors at the measured transition location decrease gradually. Applying the leading-edge contamination criterion of Pfenninger, the e(N)-method predicts the sweep angle, where leading-edge contamination starts to influence the laminar boundary-layer flow. Furthermore, the sweep angle, which produces, according to an empirical correlation, fully turbulent leading-edge flows, is well predicted by the e(N)-method.

Roughness Based Crossflow Transition Control for a Swept Wing Airfoil Design

Effect of an end plate on surface pressure distributions of two swept wings

An experimental study of the breakdown of leading-edge vortices on diamond, cropped, delta, and double delta wings during dynamic pitching

This study presents the first comprehensive simulations of dynamic flows past a permeable semi-circular cylinder, exploring wide parameter space of porosity (10−4–0.8) and angle of attack (0°–180°) at a Reynolds number of 150. Employing lattice Boltzmann method with block-structured mesh refinement, we utilized a high-resolution Cartesian grid to correlate the complex vortex structures to fluid mechanics. Five distinct flow regimes were identified as overshoot (labeled as OS), overshoot-primary (OP), primary–meandering (PM), primary-two-layered (PT), and primary-two-layered-secondary (PTS). These regions are distributed within the variation of porosity and angle of attack, revealing a systematic transition in wake dynamics. The OS region occurs at high porosity, characterized by double elongating vortex structures. Otherwise, the PM region appears at moderate porosity and low or high angles of attack, indicating oscillatory behavior of wake vortex. In particular, the PTS region is observed at low or high angle of attack, indicating identical wake patterns to those of rigid semi-circular cylinder. The angle of attack significantly influences the time-averaged drag coefficient map, showing its symmetric distribution around the angle of attack of 90°, with the PTS regime having the highest drag coefficient. The Strouhal number peaks in the OP and PT regimes, signifying the angle of attack's role in the stretching of upper and lower shear layers, which leads to the highest frequency of primary vortex shedding. An increase in the porosity coefficient delays the transition of primary-two-layered-secondary vortex shedding modes.

Unsteady pressure distributions at the wind tunnel model of a pitching Lambda wing with development of vortical flow

A refined mathematical definition for the aerodynamic center of a wing or complete airplane is presented, which allows for inclusion of the trigonometric and aerodynamic nonlinearities. From this definition, analytical relations are developed for both the axial and vertical positions of the aerodynamic center at any angle of attack, independent of whether or not the nonlinearities are included. Results show that, when all nonlinearities are included, the position of the aerodynamic center can change with angle of attack. Because the traditional approximation provides only a fixed axial coordinate, this analysis provides two additional pieces of information about the aerodynamic center, that is, its vertical coordinate and its movement. Two examples are presented, which separate these two effects. The first example uses computational fluid dynamics to show that, when the effects of the nonlinearities are combined with wing sweep at high angles of attack below stall, the aerodynamic center of a planar wing moves significantly aft and below the wing. The second example, which uses an approximate closed-form solution for two lifting surfaces, emphasizes the importance of knowing the vertical position of the aerodynamic center and shows that, in the absence of sweep, its movement with angle of attack is slight.

Results are presented of the calculation of the laminar boundary layer on infinitely long elliptic cylinders in a supersonic perfect gas flow at an arbitrary angle of attack. It was assumed that the Prandtl number is constant and equal to 0.7, the dynamic viscosity coefficient follows a power-law variation (μ ≈ T0.76) with temperature, and there is high heat transfer at the body surface (H1w=0.05). The calculations showed that a change of the body shape—the ellipticity coefficient δ=b/a—has a significant effect on the nature of the distribution and the magnitude of the local heat flux. In evaluating the thermal fluxes at the blunt leading edges, swept wings are usually considered as infinitely long yawed cylinders. In studying heat transfer at the surface of bodies of small aspect ratio at high angles of attack, wide use is made of the hypothesis of plane sections, when each section, orthogonal to the longitudinal axis of the body, is considered equivalent to a corresponding yawed infinite cylinder. By now quite detailed studies have been made of the behavior of the boundary layer on an infinitely long yawed circular cylinder with both the laminar and turbulent flow regimes for a compressible gas [1, 2]. However, there are no data on the heat transfer at the surface of a yawed infinite cylinder with arbitrary cross section, although the availability of such data is urgently needed, for example, for the proper selection of the form of the leading edges of the swept wing. This article presents the results of the calculation of the characteristics of the laminar boundary layer on the surface of infinite elliptic cylinders in a supersonic perfect gas flow. The calculations were made over a quite wide range of flight Mach number M, yaw angle λ, and ellipticity factor δ. The data presented on the distribution of the relative heat flux along the cylinder directrix may be used also for estimating the heat flux with account for the real properties of air if we know the corresponding value of the heat flux in the vicinity of the stagnation line.

In this paper the physics of the flow past a submerged vane at high angles of attack is presented. The data presented were collected during an experimental study of flow past vanes at various angles of attacks in a deformable-bed straight rectangular channel. The first few experimental runs were undertaken to establish the optimal angle of attack required to generate the strongest secondary circulation in the flow. At this optimal angle of attack, the flow features responsible for the transverse movement of bed sediment are clearly identified. The differences between the flow features observed near the vanes at both low and high angles of attack, are also outlined. These include, among others, two counterrotating vortices for vanes at high angles of attack as opposed to the widely acclaimed hypothesis of a single vortex for low angles of attack.

Investigations have been made adopting experiments and computations on an ogive-nosed slender body at different angles of attack and Reynolds number of 29,000 based on the model base diameter diameter. The results indicated an increase in the side force at large angles of attack, which is mainly due to the presence of asymmetric vortices in the leeward of the body. The inclusion of a rectangular cross-sectioned ring in the initial portion of the body reduced the side force at higher angles of attack. However, significant side force was experienced at lower angles of attack (30° < α < 40°). Use of a ring of 3% height was found to be suitable for reducing the side force at a higher angle of attack. From the results obtained it was observed that a ring if placed at a different axial location alters the flow field and changes the side force at higher angles of attack. Further studies indicated that placing of rings pair at an axial location of 3.5 and 4.5 times the base diameter reduced the side force to a very low value at all the angles of attack for the present shape of body and flow conditions.

A series of experiments was carried out to investigate the effects of active flow control on the aerodynamic characteristics of a cranked-delta-wing model at a subsonic speed. A dielectric barrier discharge (DBD) plasma actuator was used to improve the aerodynamic performance of the half-span cranked-delta-wing model at various angles of attack. The model had inboard and outboard sweepback angles of 57 and 38 deg, respectively. The surface pressure distribution was obtained at five streamwise stations and at two ray stations. All data were obtained at a constant Reynolds number of 0.27×106 and at various angles of attack: 0 deg≤α≤36 deg. The results showed that the pulsed DBD plasma actuator improved the aerodynamic performance of the cranked-delta-wing model by preventing flow separation at high angles of attack. The deep stall angle of attack was delayed up to 6 deg, and the poststall suction coefficient was increased approximately 30% at an angle of attack of about α=34 deg when the pulsed DBD plasma was applied at its optimum condition (F+≈1 and duty cycle=10%). It was found that the pulsed plasma actuator suppressed the wake flow in the forward portion of the wing at poststall angles of attack, thus delaying the flow separation. However, the effects of the DBD actuator at very high angles of attack on the flow reattachment process at the rear portion of the wing for the present DBD setting were seen to be minimal. The induced flow by the pulsed DBD plasma actuator influenced the vortex merging phenomena and delayed the onset of the outer vortex burst at prestall angles of attack.

Incompressible flow feature at near stall angle off attack on classical and wavy surface wing with aspect ratio λ = 3,7 at Reynolds number 1,5 · 105 was investigated. Such kinds of the wings at the same Reynolds number are employing on micro air vehicle. Vitality of the small aircraft at near ground atmospheric condition flight depends from predictable wing behavior at near stall angle of attack. Separation and attached flow structure at lee surface of wavy and classical wing was shown by oil-film visualization at sweep and no sweep angle condition. It was shown the wavy wing at no sweep condition has the stall angle of attack lager than classical one. At post-stall angle of attack α = 9º for classical wing sweep angle increasing from 15º to 45º results to changing from separation flow regime to attached one. On wavy wing at the same angle of attack and sweep angle from 0º to 45º the attached flow is observed only. In practice, this means that micro air vehicle will easy controlling by automatic system and will fall in stall regime rarely

Wake dynamics and linear modal instabilities of flow over NACA0015 finite-aspect-ratio, swept and tapered wings are investigated numerically. Two configurations are considered: a symmetric semi-aspect ratio (sAR) 4 wing at Reynolds number of 400 and a cantilevered sAR = 2 wing at Re = 600. The symmetric wing cases analysed at three angles of sweep and a range of taper ratios expand on the previous work on finite swept wings and complement the analysis of sAR = 2 wings reported in Part I. The cantilevered cases match the experimental configurations examined in Part II and allow cross-validation of results. Even small taper ratios are shown to have a significant effect on the wake characteristics. Base flows for global stability analysis are generated using selective frequency damping to obtain stationary linearly unstable flows. Linear global (TriGlobal) modes are computed for the purpose of flow control. These results are compared with earlier work on untapered wings at corresponding conditions to assess the effects of taper on the stability of these flows. Same families of global modes are found on both tapered and untapered wings. The spanwise location of the spatial peaks of the leading global mode supports the observation made in Part I that the trailing edge tapering concentrates unsteadiness near the root plane, whereas leading edge tapering has the opposite effect. These results give important insights into the effects of taper on the wake dynamics of tapered wings and form a first step towards understanding flows at higher Reynolds numbers.

Upper surface flaps commonly referred to as spoilers or drag brakes can increase maximum lift, and improve aerodynamic efficiency at high, near-stall angles of attack. This phenomenon was studied experimentally and computationally using a 0.307626 m chord length NACA 2412 airfoil in six different configurations, and one baseline clean configuration. A wind tunnel model was placed in the Ryerson Low Speed Wind Tunnel (atmospheric, closed-circuit, 3 ft × 3 ft test section) at a Reynold’s number of approximately 780,000 and a Mach number of 0.136. The wind tunnel study increased the lift coefficient by 0.393%-2.497% depending on the spoiler configuration. A spoiler of 10% chord length increased the maximum lift coefficient by 2.497 % when deflected 8º, by 2.110% when deflected 15º, and reduced the maximum lift coefficient by 2.783% when deflected 25º. A spoiler of 15% chord length produced smaller maximum lift coefficient gains; 0.393% when deflected 8º, by 1.760% when deflected 15º, and reduced the maximum lift coefficient by 4.475% when deflected 25º. Deflecting the spoiler increased the stall angle between 37.658% and 87.544% when compared with the clean configuration. The drag coefficient of spoiler configurations was lower than the clean configuration at angles of attack above 18º. The combination of the increased lift and reduced drag at angles of attack above 18º created by the spoiler configurations resulted in a higher aerodynamic efficiency than the clean configuration case. A 10% chord length spoiler deflected at 8º produced the highest aerodynamic efficiency gains. At low angles of attack, the computational study produced consistently higher lift coefficients compared with the wind tunnel experiment. The lift-slope was consistent with the wind tunnel experiment lift-slope. The spoiler airfoil stall behaviour was inconsistent with the results from the wind tunnel experiment. The drag coefficient results were consistent with the wind tunnel experiment at low angles of attack. However, the spoiler equipped airfoils did not reduce drag at high angles of attack. Therefore, the computational model was not valid for the spoiler configurations at high angles of attack.