Search

This paper presents a computational study of unsteady transonic effects on a pitching supercritical laminar airfoil using high-order implicit large-eddy simulation. A CAST10-2 airfoil operating at a chord-based Reynolds number of [Formula: see text], a transonic freestream Mach number of [Formula: see text], and a small nominal angle of attack of [Formula: see text] deg was subjected to forced pitching oscillations about the quarter-chord with a small amplitude of [Formula: see text] deg and reduced frequencies of [Formula: see text], 0.23, and 0.285. Complex interactions between flow separation, the motion of multiple shocks, and excursions of transitional flow were elucidated and compared against inviscid simulations. The boundary layer was found to be highly sensitive to pitch frequency, due in large part to the effects of frequency on shock motion. [Formula: see text] was found to be indirectly affected by the impact of shock motion on the boundary-layer state. Aerodynamic damping diminished with increasing oscillation frequency until becoming negative at [Formula: see text], indicating potential for aeroelastic instability.

The role of aspect ratio on the dynamic stall process of an unswept finite wing is investigated using high-fidelity large-eddy simulations. Three aspect ratios ([Formula: see text], 8, and 16) are explored for wings (NACA 0012 cross section) at chord Reynolds number [Formula: see text] and freestream Mach number [Formula: see text]. The wings pitch sinusoidally from initial incidence of 4° to a maximum angle of attack of 22° with reduced frequency [Formula: see text] over one pitching cycle. The three-dimensional unsteady flowfields show similarity among the three wings through laminar separation bubble formation/bursting. The flow topology during dynamic stall exhibits distinctly different evolutions at the higher aspect ratio relative to the lower, baseline aspect ratio. Rather than evolving into a [Formula: see text] vortex ([Formula: see text]), the higher-aspect-ratio wings show dramatic three-dimensional deformation of the vortex tube that resembles cellular structures. The vortical structure eventually interacts with the trailing-edge vortex, which contrasts with the lower aspect ratio. Examination of the unsteady loads shows an increase in lift slope, average loads, peak loads, and earlier stall with aspect ratio.

High-fidelity large-eddy simulations are performed to study the unsteady boundary layer sensitivity of a natural-laminar-flow airfoil (NLF-0414) undergoing dynamic stall in comparison to two tradit...

Numerical calculations were carried out in order to investigate the delay of transition to turbulence on a wing section by means of local dynamic surface deformation. Physically, the deformation ma...

This paper presents high-fidelity implicit large-eddy simulations of counterclockwise-oriented parallel vortical gust interactions with a NACA0012 airfoil operating at a transitional chord-based Re...

Large-eddy simulations of the onset of dynamic stall for a NACA 0012 airfoil, pitching at a constant rate, are performed for a chord Reynolds number of 1.0×106. Using four simulations, spanning M∞=...

A novel high-frequency flow control strategy for mitigation of dynamic stall is demonstrated for the case of a finite wing using high-fidelity wall-resolved implicit large-eddy simulations. A NACA ...

An investigation of spanwise end effects on deep dynamic stall of a nominally two-dimensional pitching wing section is performed using high-fidelity large-eddy simulations (LES). This study was mot...

A three-dimensional transitional shock boundary layer interaction (SWBLI) over a finite-span flexible panel is investigated by performing direct numerical simulations (DNS). The laminar inflow is at Mach 2, on which an oblique shock of turn angle 5 . 6 2 ∘ is imposed. The shock impinges at the mid-chord length of the panel, giving rise to flow separation due to the induced adverse pressure gradient. The flow separation leads to streamline curvature in the laminar boundary layer, initiating a centrifugal instability similar to Gortler instability; where the nominally two-dimensional and steady SWBLI becomes three-dimensional and unsteady at a critical Reynolds number. The flow transition is characterized by the presence of counter-rotating streamwise-oriented Gortler-like vortices, engendering spanwise undulations for increasing Gortler number. The transitional SWBLI is analyzed over both the rigid and flexible panels by performing proper orthogonal decomposition (POD) in order to examine the onset of transition as well as the modal response of the flexible panel. The presence of flexible panel promotes flow transition, resulting in lower values of the critical Reynolds and Gortler numbers. Furthermore, at a post-critical Reynolds number, the interaction results in increased turbulence levels and exchange of energy at the low and high wavenumber modes.

A numerical study has been conducted to investigate the unsteady nature of afterbody vortices shed from an upswept region of a simple fuselage configuration. The upsweep angle considered is ϕ=28 d...

An investigation of the effect of moderate sweep on three-dimensional (3D) dynamic stall of a pitching finite-aspect-ratio wing is performed by means of large-eddy simulations. The wings have an as...

This work explores the aerodynamics of an airfoil oscillating in a uniform-shear flow at chord Reynolds number of approximately 1.2×104 using complementary two-dimensional Navier–Stokes computation...

Direct numerical simulations were carried out in order to reproduce the generation and control of transition on a flat plate by means of local dynamic surface modification. The configuration and fl...

A recently proposed high-frequency control concept for dynamic-stall mitigation is demonstrated using large-eddy simulation of a pitching NACA 0012 airfoil at a chord Reynolds number of Rec=106. At...

Effect of airfoil thickness on onset of dynamic stall is investigated using large eddy simulations at chord-based Reynolds number of 200 000. Four symmetric NACA airfoils of thickness-to-chord ratios of 9 %, 12 %, 15 % and 18 % are studied. The three-dimensional Navier–Stokes solver, FDL3DI is used with a sixth-order compact finite difference scheme for spatial discretization, second-order implicit time integration and discriminating filters to remove unresolved wavenumbers. A constant-rate pitch-up manoeuver is studied with the pitching axis located at the airfoil quarter chord. Simulations are performed in two steps. In the first step, the airfoil is kept static at a prescribed angle of attack ($=4^{\circ }$). In the second step, a ramp function is used to smoothly increase the pitch rate from zero to the selected value and then the pitch rate is held constant until the angle of attack goes past the lift-stall point. The solver is verified against experiments for flow over a static NACA 0012 airfoil. Static simulation results of all airfoil geometries are also compared against XFOIL predictions with a generally favourable agreement. FDL3DI predicts two-stage transition for thin airfoils (9 % and 12 %), which is not observed in the XFOIL results. The dynamic simulations show that the onset of dynamic stall is marked by the bursting of the laminar separation bubble (LSB) in all the cases. However, for the thickest airfoil tested, the reverse flow region spreads over most of the airfoil and reaches the LSB location immediately before the LSB bursts and dynamic stall begins, suggesting that the stall could be triggered by the separated turbulent boundary layer. The results suggest that the boundary between different classifications of dynamic stall, particularly leading edge stall versus trailing edge stall, is blurred. The dynamic-stall onset mechanism changes gradually from one to the other with a gradual change in some parameters, in this case, airfoil thickness.

An investigation of dynamic stall on a pitching NACA 0012 aspect ratio 4 wing is performed by means of high-fidelity wall-resolved large-eddy simulations. The flow parameters are freestream Mach nu...

Direct numerical simulations (DNSs) were carried out in order reproduce the generation and control of transition on a flat plate by means of local dynamic surface modification. The configurations and flow conditions duplicate those of previous numerical investigations, and are similar to an experimental arrangement, which employed piezoelectrically driven actuators to impart small amplitude local deformation of the plate surface. In those studies, one actuator was located in the upstream plate region, and oscillated at the most unstable frequency of 250 Hz in order to generate small disturbances, which amplified Tollmien–Schlichting instabilities. A second actuator placed downstream, was then oscillated at the same frequency, but with appropriate amplitudes in order to mitigate disturbance growth and delay the evolution of transition. Prior simulations employed an empirical process to determine optimal values of the control parameters. In the current effort, this process is replaced with a closed-loop control law. Numerical solutions are obtained to the two-dimensional and three-dimensional compressible Navier–Stokes equations, utilizing a high-fidelity numerical scheme and an implicit time-marching approach. Local surface modification of the plate is enforced via grid deformation. Results of the simulations are presented, and features of the flowfields are described. Comparisons are made between results obtained with the two control methods, and effectiveness of the closed-loop approach is evaluated.

Dynamic stall due to a ramp-type pitching motion is investigated on the NACA 0012 airfoil at chord Reynolds number of $Re_{c}=1.0\times 10^{6}$ through the use of wall-resolved large-eddy simulation. Emphasis is placed on the unsteady boundary-layer interactions that develop as the airfoil approaches stall. At this Reynolds number it is shown that turbulent separation moves upstream across much of the airfoil suction surface. When turbulent separation reaches the leading-edge separation bubble, a bursting event is initiated leading to a strong coherent leading-edge vortex structure. This vortex wraps up the turbulent shear layer to form a large dynamic stall vortex. The use of large-eddy simulation elucidates the roll of the laminar separation bubble in defining the onset of the dynamic stall process. Comparisons are made to identical simulations at lower Reynolds numbers of $Re_{c}=0.2\times 10^{6}$ and $0.5\times 10^{6}$. This comparison demonstrates trends in the boundary-layer mechanics that explain the sensitivity of the dynamic stall process to Reynolds number.

Oblique shockwaves may impinge on supersonic vehicles internally in an engine or externally on the outer mold line. They create a severe loading environment and may induce dynamic instabilities such as panel flutter. This study computationally explores the effect of shock-induced panel flutter response in 3D, inviscid, Mach 2 flow. Flutter behavior of a square panel is compared across several incident shock angles and inflow dynamic pressures. The presence of an impinging shockwave is found to produce panel flutter that is characteristically different than the shock-free condition. The response contains significantly larger local pressure gradients, larger spanwise variations, and higher-order modal activity in the panel. Results are also compared to two-dimensional inviscid flow over an infinite-span panel. The 3D centerline is found to compare closely with 2D simulations. However, away from the centerline, 3D effects have a significant influence on the solution. In general, stronger oblique shockwaves raise flutter amplitude and frequency, while weaker shockwaves stabilize the panel response. These latter findings indicate important considerations for both structural lifing and flow control applications.

This paper presents high-fidelity implicit large-eddy simulations of parallel vortical-gust interactions with a NACA0012 airfoil operating at a transitional Reynolds number of Rec=200,000 and an an...

Dynamic stall represents a challenge in a number of engineering applications including rotorcraft, maneuvering aircraft, gust encounters, and wind turbines. Delay of the onset of dynamic stall and ...

This article explores the evolution of the unsteady flow structure for rigid and elastically mounted NACA0012 airfoils subject to a parallel vortical gust disturbance at a Reynolds number of Re = 150 , 000 using implicit large-eddy simulation coupled with a structural dynamics model. A Taylor vortex is supplied upstream of the airfoil and shown to be successful at eliciting laminar separation flutter in the elastically mounted system under conditions which normally require an artificial disturbance. Much of the gust-induced moment responsible for flutter excitation is supplied by a two-stage flow transition process on the undersurface of the wing after the gust passes the airfoil. The gust response triggers transition of the separated lower-surface boundary layer into spanwise coherent vortices followed by a laminar separation bubble accompanied by a secondary emergence of flow transition. These events appear to be analogous in some ways to the processes that appear during fully developed flutter but occur over a shorter timescale.

This work explores self-sustained pitching oscillations of a NACA0012 airfoil operating at low-to-moderate Reynolds numbers in which the aerodynamic flow is in a transitional regime. One-degree of freedom (DOF) pitching oscillations were explored over a range of Reynolds numbers ( 7 . 7 × 1 0 4 Re c ≤ 2 . 0 × 1 0 5 ) using high-order implicit large-eddy simulation coupled with structural dynamics. Limit-cycle oscillation is observed at all Reynolds numbers tested but requires a disturbance to initiate at the highest flow speeds identifying a bifurcation in possible solution states. In all cases, aerodynamic loading is dominated by primarily two features. Negative aerodynamic damping is largely provided by suction beneath a separation bubble located behind the elastic axis. This feature induces a moment that reinforces the pitch-rate at small angles of incidence and is directly influenced by flow transition at different Reynolds numbers. Open trailing edge separation on the opposite surface transitions and reattaches immediately preceding the largest angles of incidence. This process imparts a spike in the pitching moment that opposes the pitch-rate and briefly damps oscillations. Transition of the detached shear layer occurs at smaller angles of incidence as Reynolds number is increased, attenuating oscillation amplitude as Reynolds number is increased.

The onset of unsteady separation and dynamic stall vortex formation over a constant-rate pitching airfoil is analyzed by means of high-fidelity large-eddy simulations. The flowfields are computed b...

The current work reports on the results of high-accuracy two-dimensional and high-fidelity three-dimensional Navier–Stokes simulations of an SD7003 airfoil interacting with canonical upstream flow ...