Flap Deflection Angle Research Articles

A119 制御舵面を過ぎる極超音速流の挙動

Understanding of hypersonic viscous interaction phenomenon over control surfaces is of particular impor-tance for designers of hypersonic vehicles. An experimental study of the behaviour of hypersonic flow past flaps is presented in this report. Impulsive glow discharge technique and Schlieren visualization using high speed video camera are applied to examining the structure upstream of the hinge. These experiments are conducted in a hypersonic gun tunnel at Mach number 10 and at Reynolds numbers based on the chord length ranging 0.8 - 2.12 × 105. The flap deflection angles studied are in the range from 0 to 0.523 rad at zero incidence. Numerical simulations based on the Navier-Stokes equations are also performed. Time lines obtained by the dis-charge technique are reasonable and the structure evaluated by the time lines correlates well with those through Schlieren and computational results. Normalized upstream influence increases with increasing Reynolds number and increasing flap angle in the range studied.

Modeling and Analysis of Active Flap Using Coiled Bender Piezoelectric Actuators

The applicability and effectiveness of three novel piezoelectric coiled bender actuators to deflect a plain trailing edge flap on a straight wing in the presence of incompressible flow airloads is investigated in this paper through analytical modeling. This work builds upon previous research in which the three actuator configurations, the linear spiral type, the helicoidal type, and the hybrid type, have been proposed and analytically modeled, and the first two configurations have been experimentally validated. A simple analytical model for the piezoelectrically actuated flap is developed and used for this investigation. Results show that the use of coiled piezoelectric actuators for flap deflection is an effective approach. The variation of flap deflection angles with applied actuation voltage and wing angle of attack is also shown, together with performance envelopes.

遷音速領域での超音速航空機の前縁フラップによる揚抗比の改善について

Wind tunnel tests were conducted to investigate lift to drag ratio improvement by the leading-edge flap the outboard wing on an SST model at transonic regions. Force measurements and surface pressure measurements were performed for the SST model with and without outboard leading-edge flaps of 5 and 12.2 degrees deflection angles. The lift to drag ratio was improved due to a reduction in the drag component when flaps deflected, because the flow was attached to the leading edge surface of the wing. The optimum flap deflection angles to attain the maximum lift to drag ratio at a fixed lift coefficient were estimated using experimental results.

An Investigation of Flow Fields Over Multi-Element Aerofoils

This paper presents results obtained from a combined experimental and computational study of the flow field over a multi-element aerofoil with and without an advanced slat. Detailed measurements of the mean flow and turbulent quantities over a multi-element aerofoil model in a wind tunnel have been carried out using stationary and flying hot-wire (FHW) probes. The model configuration which spans the test section 600mm×600mm, is made of three parts: 1) an advanced (heel-less) slat, 2) a NACA 4412 main aerofoil and 3) a NACA 4415 flap. The chord lengths of the elements were 38, 250 and 83 mm, respectively. The results were obtained at a chord Reynolds number of 3×105 and a free Mach number of less than 0.1. The variations in the flow field are explained with reference to three distinct flow field regimes: attached flow, intermittent separated flow, and separated flow. Initial comparative results are presented for the single main aerofoil and the main aerofoil with a nondeflected flap at angles of attacks of 5, 10, and 15 deg. This is followed by the results for the three-element aerofoil with emphasis on the slat performance at angles of attack α=10, 15, 20, and 25 deg. Results are discussed both for a nondeflected flap δf=0deg and a deflected flap δf=25deg. The measurements presented are combined with other related aerofoil measurements to explain the main interaction of the slat/main aerofoil and main aerofoil/flap both for nondeflected and deflected flap conditions. These results are linked to numerically calculated variations in lift and drag coefficients with angle of attack and flap deflection angle.

Flap-Deflection Optimization for Transonic Cruise Performance Improvement of Supersonic Transport Wing

Wing flap deflection angles of a supersonic transport are optimized to improve transonic cruise performance. For this end, a numerical optimization method is adopted using a three-dimensional unstructured Euler code and a discrete adjoint code. Deflection angles often flaps; five for leading edge and five for trailing edge, are employed as design variables. The elliptic equation method is adopted for the interior grid modification during the design process. Interior grid sensitivities are neglected for efficiency. Also tested is the validity of the approximate gradient evaluation method for the present design problem and found that it is applicable for leading edge flap design in cases of no shock waves on the wing surface. The BFGS method is used to minimize the drag with constraints on the lift and upper surface Mach numbers. Two design examples are conducted; one is leading edge flap design, and the other is simultaneous design of leading edge and trailing edge flaps. The latter gave a smaller drag than the former by about two counts.

Flow field measurements of leading-edge separation vortex formed on a delta wing with vortex flaps

Wind tunnel tests are carried out using a 70 · delta wing model with leading-edge vortex flaps. The structure of the leading-edge separation vortex over the leading-edge vortex flap is measured by use of a 5 holes pitot probe, surface pressure measurement technique and oil flow visualization technique. Separation vortices formed on a plain delta wing, on a vortex flap and inboard the vortex flap hinge line are clearly visualized. Results indicate that the flow around the vortex flaps is classified into several different cross flow patterns. The streamwise flap deflection angle is defined to discuss the vortex flap performance. The optimum lift to drag ratio is attained when the amount of the wing angle of attack is not far different from that of the streamwise flap deflection angle, as long as the vortex flap is deflected modestly.

ボルテックス・フラップ付きデルタ翼の静的ロール特性について

Static normal force and rolling moment measurements for 60° and 70° delta wings with vortex flaps have been done for different roll angles at a Reynolds number based on a centerline chord of 1.3×105. Smoke visualization tests have also been done. Results show that the 60° delta wing is statically unstable in roll at the angle of attack of 35° when vortex flaps were deflected 30° downward. This is because the chordwise position where the vortex breakdown occurs is different for the left and right wings when in roll. The 70° delta wing with vortex flaps possesses static roll stability. Static rolling moment hysteresis has been observed for the 60° delta wing with a flap deflection angle of 30° at an angle of attack of 45°.

Fundamental Aeroservoelastic Study Combining Unsteady Computational Fluid Mechanics with Adaptive Control

A computational two-dimensional aeroservoelastic study in the time domain is described. The model, which is based on exact inviscid aerodynamics, captures the large amplitude motions and the associated strong shock dynamics in the transonic regime. The aeroservoelastic system consists of a two-degree-of-freedom airfoil with a trailing-edge control surface. By the use of first-order actuator dynamics, a digital adaptive controller is applied to provide active flutter suppression. Comparisons between time responses of the open-loop and closed-loop systems show the ability of the trailing-edge control surface to suppress nonlinear transonic aeroelastic phenomena. A relation between actuator dynamics, sampling time step, and limits on the flap deflection angle to guarantee the effectiveness of the adaptive controller is illustrated by the results.

Development of Vortical Structure over Delta Wing with Leading-Edge Flap

Development of a vortical structure over a delta wing with an oscillating leading-edge flap is studied experimentally. The base delta wing has a 60-deg sweep-angle and was inclined at 25 deg angle of attack. The width of the leading-edge flap varies from the apex to the trailing edge of the wing, forming a triangular shape of 5-deg apex angle. As the leading-edge flap is deflected statically on the windward side, the flow structure strongly depends upon the angle of attack and the flap deflection angle. Significant delay of the vortex breakdown location can be attained while the time scale (or the period) of the oscillating leading-edge flap is close to one convection time scale. When the oscillating leading-edge flap has a period much shorter than the convection time scale, the effectiveness in delaying the vortex breakdown location is reduced. Meanwhile, the development of vortex breakdown experiences a certain amount of time delay (or phase shift).

Aerodynamic characteristics of strake vortex flaps on a strake-wing configuration

The effects of strake vortex flaps (SVFs), determined experimentally, on the aerodynamic characteristics of a strake-wing configuration are presented. SVFs may improve cruise performance over that of a planar strake by partially unloading the strake and generating a thrust component. The magnitude of the nose-up pitching moment may also be reduced by unloading the strakes. The use of SVFs as lateral control devices is also investigated. The results indicate that cruise performance is improved for all vortex flap deflection angles compared to planar strakes, with minimal or no concomitant penalty at high lift coefficients. Positive pitching moment is also substantially reduced. Differentially deflected strakes appear to be capable of generating significant rolling moments at high angles of attack.

Experimental studies of vortex flaps and vortex plates

Low-speed wind-tunnel tests have been made on a number of vortex flap and vortex plate configurations at the Cranfield Institute of Technology. The objectives of the experiment are to assess the benefits of these devices on the lift/drag ratio improvement of delta wings. The force and surface pressure measurements were made on a 1.15-m span, 60-deg delta wing model. Results indicate that the vortex flap deflection angle, which causes the flow to attach on the flap surface without any large separation, shows a much higher lift/drag ratio than the flap deflection angle which forms a leading-edge separation vortex over the flap surface. The performance of a vortex plate protruding from the leading edge of the datum delta wing is comparable to that of the vortex flap. However, when the vortex plate is used with the vortex flap deflected, it showed no benefit in these tests.

Experiments on a 60° delta wing with vortex flaps and vortex plates

AbstractLow speed wind tunnel tests have been made to investigate the flow around a leading edge vortex flap at the maximum LID condition. Tests were also made to measure the performance of a vortex plate. The force measurements and flow visualisation tests were conducted on a 0·53 m span 60° delta wing model. Results indicate that the lift to drag ratio is a maximum for any given flap deflection angle that the flow comes smoothly onto the deflected vortex flap without forming a large leading edge separation vortex on the flap surface. The benefit of the vortex plate is seen in the drag results which are smaller than those for the datum wing. This benefit is due to leading edge suction acting on the forward facing region between the delta wing and the vortex plate.

スラット・フラップ翼まわりの流れの計算

In this paper we show a numerical procedure for computing the incompressible viscous flow about an airfoil with a slat and a flap. The full Navier-Stokes equations are discretized in the curvilinear coordinates, using the cell method. The generalized Quick method is incorporated to suppress the instability from the convective terms at higher Reynolds numbers. The Multi-Domain technique, which divides the whole computational domain into several subdomains, is used to calculate this flow problem with a complicated geometry. Each solution is connected at the domain boundary by a versatile interpolation method. This procedure has been applied to calculate the flow about a NACA4412 airfoil at high Reynolds number for various attack angles, which resulted in a good agreement with experiment. Furthermore, the flows around an airfoil with a slat and a flap were successfully calculated. The flow patterns around the airfoil and inside the narrow slat and flap gap, and the effect of a jet through the gap are made clear. These results also show reasonable aerodynamic coefficients up to the stall region for two cases of flap deflection angle: 0°and 40°.

Calculation of flow over multielement airfoils at high lift

An interactive boundary-layer procedure has been used to calculate the flow around three two-element airfoil arrangements. The procedure is known to be accurate, is seen as the foundation of a generally applicable calculation method, and is used here in comparatively simple form. The calculated results are in close agreement with measurements for angles of attack up to around 10 deg, with flap-deflection angles of up to 20 deg. The range of accuracy of the predictions can be extended by the incorporation of the wake, and this will be required to deal with high angles of attack, high flap deflection angles, and airfoil elements with smaller slot gaps than those considered here.

Aerodynamics of Inverted Leading-Edge Flaps on Delta Wings

Subsonic wind-tunnel tests were conducted to determine the aerodynamic effects of leading-edge flaps deflected upward from 60- and 75-deg sweep delta wings. Leading-edge flaps of various sizes and shapes were tested at a range of flap deflection angles. It was found that inverted flaps cause a strong vortex lift at low-tomoderate angles of attack and give large increases in CL at those angles. Examination of pitching moment data reveals that the lift increases due to inverted flap use are not necessarily accompanied by the large changes in pitching moment which are associated with trailing-edge flap deployment. With a properly shaped leading-edge flap, a negative flap deflection can give substantial increases in CL with no change in longitudinal stability.

Laminar Separation, Transition, and Turbulent Reattachment near the Leading Edge of Airfoils

The laminar separation, transition, and turbulent reattachment near the leading edge of a cylindrical noseconstant thickness airfoil model were investigated using a low-turbulence, low-speed smoke wind tunnel. The locations of separation, transition, and reattachment were obtained from smoke flow photographs and surface oil flow techniques for chord Reynolds numbers from about 150,000 to 470,000. These visual data combined with static pressure distributions delineate the effects of angle of attack, flap deflection angle, and chord Reynolds number on the separation bubble characteristics. The data concerning the length of the laminar and turbulent portions of the bubble agree with the empirical prediction methods for short bubbles.

Effect of Spanwise Blowing on the Aerodynamic Characteristics of the F-5E

A 1/lO-scaIe F-5E was tested to investigate the spanwise blowing concept to provide improved aerodynamic characteristics with primary emphasis on high angle-of-attack performance and stability. Test data were obtained in the Northrop 7 X 10-ft low-speed facility at a freestream Mach number of 0.18 for a range of model angle of attack, sideslip, jet momentum coefficient, and leading- and trailing-edge flap deflection angles. Spanwise blowing on the 32 deg-swept wing from the wing/leading-edge extension (LEX) junction at a nozzle sweep angle of 55 deg resulted in LEX and wing leading-edge vortex enhancement and large vortex-induced lift increments and drag polar improvements at the higher angles of attack. Spanwise blowing improved the lateral/directional characteristics by delaying wing stall and maintaining vertical tail effectiveness to higher angles of attack. Deflecting the leading- and trailing-edge flaps down to 24 deg and 20 deg, respectively, delayed to higher model attitudes the more beneficial effects of blowing. Blowing reduces the specific excess power available for maneuvering the F-5E.

Hybrid Upper Surface Blown Flap Propulsive-Lift Concept for the QSRLP

The hybrid upper surface blowing concept consists of wing-mounted turbofan engines with a major portion of the fan exhaust directed over the wing upper surface to provide high levels of propulsive lift, but with a portion of the fan airflow directed over selected portions of the airframe to provide boundary-layer control. NASAsponsored preliminary design studies identified the hybrid upper surface blowing concept as the best propulsive lift concept to be applied to the Quiet Short-Haul Research Aircraft (QSRA) that is planned as a flight facility to conduct flight research at low noise levels, high approach lift coefficients, and steep approaches. Data from NASA in-house and NASA-sponsored small-and large-scale wind-tunnel tests of various configurations using this concept are presented. Nomenclature b = wing span BLC = boundary-layer control CL = lift coefficient, L/qS Cia = 0 = h'ft coefficient at zero angle of attack CL^ ~ 2. ' = lift coefficient at an angle of attack of -2 ° CL^ax = maximum lift coefficient C/ = rolling-moment coefficient, rolling moment/qSb Cn = yawing-moment coefficient, yawing momQnt/qSb CT = thrust coefficient, gross thrust/qS C/z = blowing momentum coefficient, gross thrust from BLC nozzle/qS L -lift Pl£IP^ = pressure ratio, total pressure at engine nacelle exit/tunnel static pressure q = freestream dynamic pressure S = wing area V = velocity W/S = wing loading a = angle of attack 6/r = deflection angle of thejrailing-edge flap

Geometry effects on STOL engine-over-the-wing acoustics with 5:1 slot nozzle

The effect of nozzle roof (kickdown) angle and nozzle location relative to the wing trailing edge on noise was determined at model scale with a 5:1 slow nozzle (equivalent diameter, 5.1 cm) for several OTW configurations. Nozzle roof angles of 10° to 40°, with and without cutback at the nozzle sides were tested. Also included, for comparison, was a 5:1 slot nozzle with external deflector. Three wing chord sizes were used: standard (33 cm), 23-standard, and 32-standard. Flap deflection angles used were 20° and 60°. The nozzle was located at 20% and 35% of chord. In addition, velocity maps were obtained at the trailing edge of the wings. With increasing wing size, representing a variety of airframe/engine installations, low-frequency noise was increasingly attenuated and shielding benefits increased directly with wing size. With increasing nozzle roof angle, the jet velocity at the trailing edge was decreased, causing a decrease in trailing-edge noise; however, the jet impingement noise on the wing increased. Cutback of the nozzle sides improved flow attachment and reduced the trailing-edge velocity and noise. The best flow attachment and least trailing-edge noise was obtained with the external deflector configuration.

Acoustic Results Obtained with an Upper-Surface-Blowing Lift-Augmentation System

The noise caused by the interaction of the exhaust jet and the wing was measured under static conditions for several versions of a small-scale STOL engine-over-the-wing configuration. Three basic nozzles were tested: a circular nozzle with deflector, a 5:1 slot nozzle with and without deflector, and a 10:1 slot nozzle. The wing included a flap lift-augmentation system with flap deflection angles of 20° for a takeoff condition and 60° for an approach condition. Far-field noise data are presented for nominal jet pressure ratios of 1.25 to 1.7 for the flyover mode. The data are discussed in terms of SPLs and normalized sound pressure spectra: comparisons with nozzle-only data are also made. The model OASPL near 90°—measured from an engine inlet—follow a 6th–7th power relationship with jet velocity compared to an 8th–9th power law for the jet alone. Only minor acoustic differences were obtained between the two flap deflection positions. The most significant noise reduction was achieved with the 10:1 slot nozzle. Implications of extending the small-scale model acoustic data to a full-scale aircraft are discussed and indicate a sizable noise attenuation may be achieved owing to wing shielding.