Post-stall Conditions Research Articles

Zonal Detached-Eddy Simulation of an Airfoil in Poststall Condition

This paper presents a zonal detached-eddy simulation of a leading-edge stall airfoil in poststall conditions. Different detached-eddy-simulation-type methodologies are first evaluated and discussed and the paper presents in detail the results obtained with the mode 1 of the zonal detached-eddy simulation that is based on a prescribed switch between the Reynolds-averaged Navier–Stokes and the large-eddy simulation regions. The results obtained with the selected approach are presented and compared to detailed experimental results including unsteady pressure measurements and unsteady velocity flowfield measurements acquired using time-resolved particle image velocimetry and laser Doppler velocimetry. Important features of the flow are correctly captured by the computation and good agreements are observed with experiments on pressure distribution, separation location and aerodynamic forces. The unsteady properties of the calculation are also compared with unsteady measurements showing that the frequency indicating that the vortex shedding is correctly predicted in the computation. Detailed analysis of the mixing-layer and the wake properties highlights some deficiency of the computations that are fully analyzed. Especially, the grid resolution in the early stages of the mixing layer appears to be of primary importance for the present application. The detailed comparisons of the zonal detached-eddy simulation result with the experimental data allow general requirements to be drawn for further improvements.

Plunging Wake Analysis of an Airfoil Equipped With a Gurney Flap

Experimental measurements were conducted on a plunging Eppler 361 Gurney flapped airfoil to study wake structure and dynamic stall phenomenon in the wake. The heights of Gurney flap were 2.6 and 3.3% chord. All oscillation data were taken at the plunging amplitude of 6 cm and Reynolds number 1.5 × 105. Special attention was focused on the temporal progressions of the plunging wake for the range of initial angle of attack (0° and 12°) in prior and post-stall flow conditions. The velocity in the wake was measured by hot-wire anemometry. Surface pressure measurements, as supplementary data, were also carried out to look into the link between the boundary-layer flow and the shedding vortical flow. It was found that the hysteresis is detected between the plunging wake in the upstroke and down-stroke. The shape and width of the wake hysteresis loops strongly depend on the initial angle of attack and vertical positions of the sensor. At prior static-stall angle of attack, positive camber effects of flapped airfoil shifted wake profiles downward and more velocity deficits were detected. In the post-stall conditions the hysteresis loop widths for lower-than-centerline vertical positions were remarkably more than counterpart upper positions. The energetic dynamic stall vortex shedding was found to be main responsible for large hysteresis and velocity deficit at these positions. Furthermore, the extent and strength of the stalled wake or flow separation were found to increase for the flapped case which results in a significant increase in the hysteresis loop widths.

Low Reynolds number airfoil aerodynamic loads determination via line integral of velocity obtained with particle image velocimetry

The small magnitude lift forces generated by both a NACA 0012 airfoil and a thin flat plate at Re = 29,000 and 54,000 were determined through the line integral of velocity, obtained with particle image velocimetry, via the application of the Kutta–Joukowsky theorem. Surface pressure measurements of the NACA0012 airfoil were also obtained to validate the lift coefficient C l. The bound circulation was found to be insensitive to the size and aspect ratio of the rectangular integration loop for pre-stall angles. The present C l data were also found to agree very well with the surface pressure-determined lift coefficient for pre-stall conditions. A large variation in C l with the loop size and aspect ratio for post-stall conditions was, however, observed. Nevertheless, the present flat-plate C l data were also found to collectively agree with the published force-balance measurements at small angles of attack, despite the large disparity exhibited among the various published data at high angles. Finally, the ensemble-averaged wake velocity profiles were also used to compute the drag coefficient and, subsequently, the lift-to-drag ratio.

Eddy-Preserving Limiter for Unsteady Subsonic Flows

A slope-limiting algorithm for second-order monotone upstream-centered schemes for conservation laws is introduced to reduce the dissipation of vortices in flow simulations. The algorithm is based on the reconstruction of velocity components along the principle axes of the vortex and the augmentation of the central gradients’ weight for the interpolation of velocity components on the swirl plane of the vortex. The performance of the scheme in different vortical flow problems is investigated. The proposed limiting algorithm has been able to considerably reduce the dissipation of vortices provided that the spatial and temporal discretization of the problem have been fine enough to resolve the length and time scales of the corresponding vortical motion. In particular, the scheme has significantly outperformedthe conventionalvan Albadalimiter to resolvethesecond peak inthe lift coefficient spectrainthe case of a NACA0021 airfoil at a poststall condition.

波状前縁付き矩形翼周りの流れ解析

The objective of the present study is to investigate the flow mechanism of delaying stall by use of a wavy leading edge on rectangular wing. Flows around rectangular wings with wavy leading edge were simulated by large eddy simulation with the Dynamic Smagorinsky Model (DSM). The Reynolds number based on the chord length of the wing and the uniform-flow velocity was 1:38X105. The predicted lift and drag coefficients for poststall condition agreed qualitatively with experimental data by using DSM, moreover those for prestall condition were in very good agreement with the experimental data. The mechanism of delaying stall was revealed by the flow simulation. The predicted flows around the wavy leading edge indicated existence of longitudinal vortices and the flows around wavy leading edge remained attached at poststall angle of attack. The longtiudinal vortices were carried to wing tips and the flow separation was suppressed in the area from the protuberances to the wing tips.

OS9-10 翼周りの流れと騒音の受動的制御による揚力・抗力への影響(OS9 熱流動,循環型社会における動力エネルギー技術)

In this paper we attempted to reduce the aerodynamic noise generated from a two-dimensional NACA0012 airfoil at a Reynolds number of 3×10^5 in post-stall condition by passive flow and noise control on airfoil surface with various materials. The experimental results showed that the Pile Fabrics reduced the aerodynamic noise with 4.4dB in post-stall condition, but in spite of the remarkably thin pile fabrics degraded lift and drag performance of the airfoil. In case of the control by relatively thin Urethane, even if airfoil surface was largely covered, there had no reduction of aerodynamic noise but little degradation of lift and drag performance. In case of the control by Net, there had least degradation of lift and drag performance among the materials and a little reduction of aerodynamic noise in post-stall condition about 1.8dB.

A study of a turbulent boundary layer in stalled-airfoil-type flow conditions

The experimental study of the turbulent boundary layer under external flow conditions similar to those found on the suction side of airfoils in trailing-edge post-stall conditions has been performed. Detailed boundary layer measurements were carried out with a PIV system and a two-sensor wall probe. They cover the region downstream of the suction peak where the boundary layer is subjected to a very strong adverse pressure gradient and has suffered from an abrupt transition from strong favorable to strong adverse pressure gradients. The experiments show that in spite of these severe conditions, the boundary layer is surprisingly able to recover a state of near-equilibrium before separating. In this near-equilibrium zone, the mean velocity defect and all the measured Reynolds stresses are self-similar (in the outer region) with respect to the outer scales δ and U e δ*/δ. The mean momentum balance indicates that for the upper half of the outer region, the advection terms dominate all the stress-gradient terms in the zone prior to separation. A large portion of the outer region has therefore become essentially an inertial flow zone where an approach toward equilibrium is expected.

G302 失速した2次元翼から発生する空力音の制御(GS-3 音響・騒音,一般セッション)

G302 失速した2次元翼から発生する空力音の制御(GS-3 音響・騒音,一般セッション)

Computation of Viscous Flow for a Boeing 777 Aircraft in Landing Configuration

A series of Navier-Stokes simulations of a complete Boeing 777-200 aircraft configured for landing is obtained using a structured overset grid process and the OVERFLOW CFD code. At approach conditions, the computed forces for the 777 computation are within 1.5% of experimental data for lift, and within 4% for drag. The computed lift is lower than the experiment at maximum-lift conditions, but shows closer agreement at post-stall conditions. The effect of sealing a span-wise gap between leading edge elements, and adding a chine onto the nacelle is computed at a high angle of attack. These additions make a significant difference in the flow over the wing near these elements. Detailed comparisons between computed and experimental surface pressures are shown. Good agreement is demonstrated at lower angles of attack, including a prediction of separated flow on the outboard flap.

Extensions to Maskell’s theory for blockage effects on bluff bodies in a closed wind tunnel

AbstractExtensions to Maskell’s original correction method, developed over several years, are consolidated and designated ‘Maskell III’. The procedures were applied to dedicated tests on a family of flat-plate wing models in a small, low-speed wind tunnel at NRC. Test conditions included angles of attack from -10° to 110° and models of up to 16% of tunnel area. Off-centre tests were included with model-to-wall distances down to 0.72 chords. Corrected lift and drag data correlated well between models of different sizes and at different offsets from the tunnel centreline. Comparisons are made with corrections using the pressure-signature and two-variable methods, emphasising post-stall conditions. These showed that the ‘Maskell III’ procedures, which require minimal input, correlated as well as the other methods for most model sizes and positions in the tunnel.

Analysis of Turbulent Separated Flows for the NREL Airfoil Using Anisotropic Two-Equation Models at Higher Angles of Attack

Numerical simulations of the turbulent flow fields at stall conditions are presented for the NREL (National Renewable Energy Laboratory) S809 airfoil. The flow is modeled as compressible, viscous, steady/unsteady and turbulent. Four two-equation turbulence models (isotropic k–ε and q–ω models, anisotropic k–ε and q–ω models), are applied to close the Reynolds-averaged Navier-Stokes equations, respectively. The governing equations are integrated in time by a new LU-type implicit scheme. To accurately model the convection terms in the mean-flow and turbulence model equations, a modified fourth-order high resolution MUSCL TVD scheme is incorporated. The large-scale separated flow fields and their losses at the stall and post-stall conditions are analyzed for the NREL S809 airfoil at various angles of attack (α) from 0 to 70 degrees. The numerical results show excellent to fairly good agreement with the experimental data. The feasibility of the present numerical method and the influence of the four turbulence models are also investigated.

Two-dimensional turbulent viscous flow simulation past airfoils at fixed incidence

The present work discusses the computation of the time-mean, turbulent, two-dimensional incompressible viscous flow past an airfoil at fixed incidence. A new physically consistent method is presented for the reconstruction of velocity fluxes which arise from discrete equations for the mass and momentum balance. This closure method for fluxes makes possible the use of a cell-centred grid in which velocity and pressure unknowns share the same location, while circumventing the occurrence of spurious pressure modes. The influence of several turbulent models is investigated. The models involve either an algebraic eddy viscosity or determine the eddy viscosity from transport equations. The method is tested on the following airfoils in pre-stall or post-stall conditions: NACA4412, AS-240 B, GAW-(1) airfoil, for which documented experimental data are available.

The X-31 aircraft: Advances in aircraft agility and performance

The X-31 enhanced fighter maneuverability (EFM) demonstrator has pioneered agile flight in the post-stall flight regime and explored integrated multi-axis thrust vectoring across a broad flight envelope. Its maneuvering achievements include sustained flight up to 70 degrees angle of attack, velocity vector rolls in deep post-stall conditions, and post-stall turns from high entry to exit speeds with ultra low turning/transitional conditions. The concept of post-stall maneuverability was extensively studied in simulations preceding initiation of the X-31 program. These simulations provided a baseline for tactical utility demonstrations and vehicle design requirements. Post-stall maneuverability was not achieved without encountering and mitigating the effects of highly unsteady, asymmetric, vortex-dominated flow-fields associated with post-stall flight. Anomalies in vehicle response to control inputs were observed at high angles of attack, as were differences between simulator and actual flight parameters due to a misrepresentation of the effects of these complex flowfields. Some preliminary force and moment data for the X-31 configuration during dynamic maneuvers are provided to highlight the complex nature of the flowfield. The X-31 aircraft's enabling capabilities, including multi-axis thrust vectoring and integrated flight/propulsion control also provided performance enhancements across the entire flight envelope. In what were known as ‘quasi-tailless’ experiments, conventional aerodynamic control surfaces were used to reduce or eliminate the stabilizing influence of the vertical stabilizer, while the vehicle's multi-axis thrust vectoring capability was used for restabilization. Properly exploited, these technologies can lead to the reduction or elimination of traditional aerodynamic control surfaces, which provides profound improvements in vehicle range, weight, payload, and low observability. This review focuses on some of the principal aerodynamic issues encountered during the course of the X-31 program, the resulting lessons learned, and the derived results concerning the complex aerodynamic effects associated with exploiting this aircraft's unique capabilities. Suggested research and technology development issues which came to light during the X-31 program are also discussed.

Impulsive start of a symmetric airfoil at high angle of attack

The fluid dynamic phenomena following the impulsive start of a NACA 0015 airfoil were studied by using a time accurate solution of the incompressible laminar Navier-Stokes equations. Angle of attack was set at 10 deg to simulate steady-state poststall conditions at a Reynolds number of 1.2 x 10(exp 4). The calculation revealed that large initial lift values can be obtained, immediately following the impulsive start, when a trapped vortex develops above the airfoil. Before the buildup of this trapped vortex and immediately after the airfoil was set into motion, the fluid is attached to the airfoil's surface and flows around the trailing edge, demonstrating the delay in the buildup of the classical Kutta condition. The transient of this effect is quite short and is followed by an attached How event that leads to the trapped vortex that has a longer duration. The just described initial phenomenon eventually transits into a fully developed separated flow pattern identifiable by an alternating, periodic vortex shedding.

Buffet excitation of wings at low speeds

Wind-tunnel experiments were undertaken to investigate the unsteady excitation arising from separated flow at low speeds. A series of untapered wing models with a common section chosen to be insensitive to Reynolds number effects was used in the investigation. The models had sweep angles of — 20, 0, and 30 deg, and aspect ratios ranging from 6 to 16. The excitation was measured indirectly using wing-root strain gauges and presented in terms of the buffet excitation parameter. The experimental data obtained in this investigation indicate that this parameter attains a common limit under poststall conditions irrespective of wing planform; physical reasons for the limit are discussed. Poststall conditions refer to angles of attack in excess of that at which the lift is maximized. Buffet excitation can attain larger values at lower incidences, and also at high incidences for lower reduced frequencies, which correspond more closely to rigid-body modes.

A rapid technique for measuring and visualizing the extent of separated flow

A method is described for rapidly measuring and visualizing the extent of separated flow suitable for a wind tunnel environment. The method utilizes a continuously swinging total pressure probe. This technique permits acquiring and presenting graphical records of separated regions in a fraction of the time required by other methods. Typical results indicate the presence of highly complicated three-dimensional separated regions for a typical general aviation twin-engine aircraft at post-stall conditions.