Trailing-edge Flap Deflection Research Articles

Numerical simulation of ice accretion on helicopter rotor blades with trailing edge flap

PurposeThis paper aims to describe a numerical simulation method of ice accretion on BO105 helicopter blades for predicting the effects of trailing edge flap deflection on ice accretion.Design/methodology/approachA numerical simulation method of ice accretion is established based on Myers model. Next, the shape and location of ice accretion of NACA0012 airfoil are calculated, and a comparison between calculated results and experimental data is made to validate the method. This method is used to investigate the effect of trailing edge flap deflection on ice accretion of a rotor blade.FindingsThe numerical method is feasible and effective to study the ice accretion on helicopter rotor blades. The downward deflection of the trailing edge flap affects the shape of the ice.Practical implicationsThis method can be further used to predict the ice accretion in actual flights of the helicopters with multielement airfoils.Originality/valueThe numerical simulation method here can lay a foundation of the research about helicopter flight performance in icing condition through predicting the shape and location of ice accretion on rotor blades.

Interference mechanism of trailing edge flap shedding vortices with rotor wake and aerodynamic characteristics

A robust Reynolds-Averaged Navier-Stokes (RANS) based solver is established to predict the complex unsteady aerodynamic characteristics of the Active Flap Control (AFC) rotor. The complex motion with multiple degrees of freedom of the Trailing Edge Flap (TEF) is analyzed by employing an inverse nested overset grid method. Simulation of non-rotational and rotational modes of blade motion are carried out to investigate the formation and development of TEF shedding vortex with high-frequency deflection of TEF. Moreover, the mechanism of TEF deflection interference with blade tip vortex and overall rotor aerodynamics is also explored. In non-rotational mode, two bundles of vortices form at the gap ends of TEF and the main blade and merge into a single TEF vortex. Dynamic deflection of the TEF significantly interferes with the blade tip vortex. The position of the blade tip vortex consistently changes, and its frequency is directly related to the frequency of TEF deflection. In rotational mode, the tip vortex forms a helical structure. The end vortices at the gap sides co-swirl and subsequently merge into the concentrated beam of tip vortices, causing fluctuations in the vorticity and axial position of the tip vortex under the rotor. This research concludes with the investigation on suppression of Blade Vortex Interaction (BVI), showing an increase in miss distance and reduction in the vorticity of tip vortex through TEF phase control at a particular control frequency. Through this mechanism, a designed TEF deflection law increases the miss distance by 34.7% and reduces vorticity by 11.9% at the target position, demonstrating the effectiveness of AFC in mitigating BVI.

Inviscid modeling of unsteady morphing airfoils using a discrete-vortex method

Abstract A low-order physics-based model to simulate the unsteady flow response to airfoils undergoing large-amplitude variations of the camber is presented in this paper. Potential-flow theory adapted for unsteady airfoils and numerical methods using discrete-vortex elements are combined to obtain rapid predictions of flow behavior and force evolution. To elude the inherent restriction of thin-airfoil theory to small flow disturbances, a time-varying chord line is proposed in this work over which to satisfy the appropriate boundary condition, enabling large deformations of the camber line to be modeled. Computational fluid dynamics simulations are performed to assess the accuracy of the low-order model for a wide range of dynamic trailing-edge flap deflections. By allowing the chord line to rotate with trailing-edge deflections, aerodynamic loads predictions are greatly enhanced as compared to the classical approach where the chord line is fixed. This is especially evident for large-amplitude deformations. Graphical abstract

Regulating rotor aerodynamics and platform motions for a semi-submersible floating wind turbine with trailing edge flaps

The coupling effects between rotor aerodynamics and platform motions are being focused for floating offshore wind turbines (FOWTs). Ignoring these effects during the controller design can deteriorate some essential loads or lead to aero-elastic instabilities. Superimposed proportional model-free adaptive control (SP-MFAC) is proposed to cope with these effects by simultaneously optimizing blade root moments, rotor speed, platform-pitch and -yaw motions. Trailing edge flap (TEF) is used as the active flow control device. Actuators’ saturation and energy consumption are considered by introducing penalty factors into control cost function. A typical semi-submersible FOWT is selected for simulation study. A comparison is made against proportional-integral-differential-based (PID-based) TEF control and baseline control in turbulent and extreme gust conditions. Numerical results reveal that SP-MFAC costs less in TEF deflection activity, but realizes better overall performance than PID-based TEF control. Further, the physical mechanisms of load reduction are analyzed with cross-wavelet transformation. The original in-phase aero-hydro-elastic mechanisms in the semi-submersible FOWT are crippled by the TEF deflection activity.

Investigation of a Realistic Flap Modeling Using a Combination of Chimera Method and Grid Deformation on a Wing Fuselage Configuration

AbstractFlap deflections of an aircraft wing for active load alleviation within CFD simulations are realized using pure grid deformation due to time saving and low modeling complexity. In this case, spanwise gaps are neglected, which are present in reality during a flap deflection. Another possibility to realize the deflections is the combination of pure grid deformation and Chimera method, which allows the modeling of the gap between flap and wing or consecutive flaps. The overall aim of this work is the analysis of the aerodynamic effects caused by the different modeling approaches realizing leading and trailing edge flap deflections. The comparison of the modeling methods is investigated on the DLR LEISA configuration, which is a generic wing‐fuselage configuration. For active gust load alleviation, the leading edge flaps are deflected downward and the trailing edge flaps are deflected upward. Due to the downward deflection of the leading edge flaps, vortices are formed using the combined Chimera method as a result of the gap consideration. These vortices lead to a local drag increase resulting in a difference between both modeling methods in the spanwise as well as global drag coefficient. With the pure grid deformation these vortices do not occur. Due to the upward trailing edge deflection, the combined Chimera method leads to a pressure compensation via the effective gap enlargement, which is not present in the pure grid deformation. Overall, the combined Chimera method offers a good possibility to model the induced drag as well as the pressure compensation at a large flap deflection.

Study the effect of trailing edge flap deflection on horizontal axis wind turbine performance using computational investigation

Study the effect of trailing edge flap deflection on horizontal axis wind turbine performance using computational investigation

Near Stall Unsteady Flow Responses to Morphing Flap Deflections

The unsteady flow characteristics and responses of an NACA 0012 airfoil fitted with a bio-inspired morphing trailing edge flap (TEF) at near-stall angles of attack (AoA) undergoing downward deflections are investigated at a Reynolds number of 0.62 × 106 near stall. An unsteady geometric parametrization and a dynamic meshing scheme are used to drive the morphing motion. The objective is to determine the susceptibility of near-stall flow to a morphing actuation and the viability of rapid downward flap deflection as a control mechanism, including its effect on transient forces and flow field unsteadiness. The dynamic flow responses to downward deflections are studied for a range of morphing frequencies (at a fixed large amplitude), using a high-fidelity, hybrid RANS-LES model. The time histories of the lift and drag coefficient responses exhibit a proportional relationship between the morphing frequency and the slope of response at which these quantities evolve. Interestingly, an overshoot in the drag coefficient is captured, even in quasi-static conditions, however this is not seen in the lift coefficient. Qualitative analysis confirms that an airfoil in near stall conditions is receptive to morphing TEF deflections, and that some similarities triggering the stall exist between downward morphing TEFs and rapid ramp-up type pitching motions.

Wind Tunnel Test of the SMART Active Flap Rotor

Boeing and a team from Air Force, NASA, Army, DARPA, MIT, UCLA, and U. of Maryland have successfully completed a wind-tunnel test of the smart material actuated rotor technology (SMART) rotor in the 40- by 80-foot wind-tunnel of the National Full-Scale Aerodynamic Complex at NASA Ames Research Center. The Boeing SMART rotor is a full-scale, five-bladed bearingless MD 900 helicopter rotor modified with a piezoelectric-actuated trailing edge flap on each blade. The eleven-week test program evaluated the forward flight characteristics of the active-flap rotor at speeds up to 155 knots, gathered data to validate state-of-the-art codes for rotor aero-acoustic analysis, and quantified the effects of open and closed loop active flap control on rotor loads, noise, and performance. The test demonstrated on-blade smart material control of flaps on a full-scale rotor for the first time in a wind tunnel. The effectiveness of the active flap control on noise and vibration was conclusively demonstrated. Results showed significant reductions up to 6dB in blade-vortex-interaction and in-plane noise, as well as reductions in vibratory hub loads up to 80%. Trailing-edge flap deflections were controlled within 0.1 degrees of the commanded value. The impact of the active flap on control power, rotor smoothing, and performance was also demonstrated. Finally, the reliability of the flap actuation system was successfully proven in more than 60 hours of wind-tunnel testing.

Numerical investigations on aerodynamic forces of deformable foils in hovering motions

In this paper, the aerodynamic forces of deformable foils for hovering flight are numerically investigated by a two-dimensional finite-volume arbitrary Lagrangian Eulerian Navier-Stokes solver. The effects of deformation on the lift force generation mechanisms of deformable wings in hovering flight are studied by comparison and analysis of deformable and rigid wing results. The prescribed deformation of the wings changes their morphing during hovering motion in both camber and angle of incidence. The effects of deflection amplitude, deflection phase, and rotation location on the aerodynamic performances of the foils, as well as the associated flow structures, are investigated in details, respectively. Results obtained show that foil morphing changes both Leading Edge Vortex (LEV) and Trailing Edge Vortex (TEV) generation and development processes. Consequently, the lift force generation mechanisms of deformable wings differ from those of rigid foil models. For the full deformation foil model studied, the effect of foil deformation enhances its lift force during both wake capture and delayed stall. There is an optimized camber amplitude, which was found to be 0.1*chord among those cases simulated. Partial deformation in the foil does not enhance its lift force due to unfavorable foil camber. TEV is significantly changed by the local angle of attack due to the foil deformation. On the other hand, Trailing Edge Flap (TEF) deflection in the hinge connected two-rigid-plate model directly affects the strength of both the LEV and TEV, thus influencing the entire vortex shedding process. It was found that lift enhancement can reach up to 33.5% just by the TEF deflection alone.

Effects of a dynamic trailing-edge flap on the aerodynamic performance and flow structures in hovering flight

To examine the effects of wing morphing on unsteady aerodynamics, deformable flapping plates are numerically studied in a low-Reynolds-number flow. Simulations are carried out using an in-house immersed-boundary-method-based direct numerical simulation (DNS) solver. In current work, chord-wise camber is modeled by a hinge connecting two rigid components. The leading portion is driven by a biological hovering motion along a horizontal stroke plane. The hinged trailing-edge flap (TEF) is controlled by a prescribed harmonic deflection motion. The effects of TEF deflection amplitude, deflection phase difference, hinge location, and Reynolds number on the aerodynamic performance and flow structures are investigated. The results show that the unsteady aerodynamic performance of deformable flapping plates is dominated by the TEF deflection phase difference, which directly affects the strength of the leading-edge vortex (LEV) and thus influences the entire vortex shedding process. The overall lift enhancement can reach up to 26% by tailoring the deflection amplitude and deflection phase difference. It is also found that the role of the dynamic TEF played in the flapping flight is consistent over a range of hinge locations and Reynolds numbers. Results from a low aspect-ratio (AR=2) deformable plate show the same trend as those of 2-D cases despite the effect of the three-dimensionality.

デルタ形状ボルテックスジェネレーターによる後縁フラップ上の離流れ制御について

Low-speed wind tunnel tests were performed to investigate aerodynamic effects induced by delta-shaped vortex generators installed on the upper surface of a trailing-edge flap. The vortex generator was applied to control the flow field around the trailing-edge flap where the flow was separated from the hinge line of the trailing-edge flap. The test results showed that the vortex generators increase the lift coefficient and the lift-to-drag ratio when the trailing-edge flap deflects to 30degrees. The leading-edge separation vortices formed from the delta-shaped vortex generators suppressed the flow separation on the trailing-edge flap. The height of the vortex generator geometrically changes depending on the trailing-edge flap deflection angles. Therefore, when the trailing-edge flap is not deflected such as cruise flight condition where the flow field is dominated by the attached flow, the vortex generator does not impose a penalty on the aerodynamic performance.

Aerodynamic response of an airfoil section undergoing pitch motion and trailing edge flap deflection: a comparison of simulation methods

AbstractThe study presents and compares aerodynamic simulations for an airfoil section with an adaptive trailing edge flap, which deflects following a smooth deformation shape. The simulations are carried out with three substantially different methods: a Reynolds‐averaged Navier–Stokes solver, a viscous–inviscid interaction method and an engineering dynamic stall model suitable for implementation in aeroelastic codes based on blade element momentum theory. The aerodynamic integral forces and pitching moment coefficients are first determined in steady conditions, at angles of attack spanning from attached flow to separated conditions and accounting for the effects of flap deflection; the steady results from the Navier–Stokes solver and the viscous–inviscid interaction method are used as input data for the simpler dynamic stall model. The paper characterizes then the dynamics of the unsteady forces and moments generated by the airfoil undergoing harmonic pitching motions and harmonic flap deflections. The unsteady aerodynamic coefficients exhibit significant variations over the corresponding steady‐state values. The dynamic characteristics of the unsteady response are predicted with an excellent agreement among the investigated methods at attached flow conditions, both for airfoil pitching and flap deflection. At high angles of attack, where flow separation is encountered, the methods still depict similar overall dynamics, but larger discrepancies are reported, especially for the simpler engineering method. Copyright © 2014 John Wiley & Sons, Ltd.

On power and actuation requirement in swashplateless primary control of helicopters using trailing-edge flaps

AbstractThis paper examines three specific aspects pertaining to the trailing-edge flap (TEF) enabled swashplateless primary control of a helicopter. The study is based on a utility helicopter very similar to the UH-60A Blackhawk helicopter, with rotor torsion frequency reduced to 2·1/rev, and 20% chord TEFs extending from 70-90% span. The questions addressed in the paper are the power penalty due to aerodynamic drag associated with TEF control, the pitch-index required to limit the range of TEF deflections over variations in aircraft gross-weight and airspeed, and the influence of rotor RPM variation on swashplateless primary control. Results indicate that the power penalty associated with TEF enabled primary control at high speeds is in the range of 6-7%, due to increased drag on the advancing side in the region of the TEFs and at the blade tips. At low to moderate speeds the increase in power is 2-4% on average, more dependent on the pitch-index, and due to drag increase over most of the azimuth in the region of the TEFs. A variation in pitch-index from 16° for lower speeds and gross weights, to 20-22° for higher speeds and gross weights, would reduce the steady level flight TEF defection requirements to under ±3°, leaving sufficient control margin. Increase in rotor RPM does not increase directly increase thrust (as with a stiff-torsion rotor) but reduces the rotating torsion frequency, and together with the increased dynamic pressures increases the sensitivity to TEF control. At low to moderate speeds a 9% increase in RPM reduces the maximum TEF deflections required by about 1°, but is accompanied by a large increase in rotor power. Conversely, a 9% RPM reduction decreases rotor power required, but the TEF defections required increase by 1–1·5°.

Frequency-Weighted Model Predictive Control of Trailing Edge Flaps on a Wind Turbine Blade

This paper presents the load reduction achieved with trailing edge flaps during a full-scale test on a Vestas V27 wind turbine. The trailing edge flap controller is a frequency-weighted linear model predictive control (MPC) where the quadratic cost consists of costs on the zero-phase filtered flapwise blade root moment and trailing edge flap deflection. Frequency-weighted MPC is chosen for its ability to handle constraints on the trailing edge flaps deflection, and to target at loads with given frequencies only. The controller is first tested in servo-aeroelastic simulations, before being implemented on a Vestas V27 wind turbine. Consistent load reduction is achieved during the full-scale test. An average of 13.8% flapwise blade root fatigue load reduction is measured.

Pitching Airfoil with Combined Gurney Flap and Unsteady Trailing-Edge Flap Deflection

b = semispan c = airfoil chord Cl = lift coefficient Cm = pitching moment coefficient Cp = surface pressure coefficient f = oscillation frequency h = Gurney flap height td = flap actuation duration ts = flap actuation start time u1 = freestream velocity = angle of attack ds = dynamic-stall angle f = final angle i = initial angle ss = static stall angle = f i, pitch amplitude t = total airfoil pitch time max = maximum flap deflection = fc=u1, reduced frequency

Reducing Trailing Edge Flap Deflection Requirements in Primary Control with a Movable Horizontal Tail

Reducing Trailing Edge Flap Deflection Requirements in Primary Control with a Movable Horizontal Tail

PIV study of flow around unsteady airfoil with dynamic trailing-edge flap deflection

The flow around an oscillating NACA 0015 airfoil with prescheduled trailing-edge flap motion control was investigated by using particle image velocimetry (PIV). Aerodynamic load coefficients, obtained via surface pressure measurements, were also acquired to supplement the PIV results. The results demonstrate that upward flap deflections led to an improved negative peak pitching moment coefficient C m,peak, mainly as a consequence of the increased suction pressure on the lower surface of the flap. The behavior of the leading-edge vortex (LEV) was largely unaffected. Its strength was, however, reduced slightly compared to that of the uncontrolled airfoil. No trailing-edge vortex was observed. For downward flap deflection, the strength of the LEV was found to be slightly increased. A favorable increase in C l,max, as a consequence of downward flap-induced positive camber effects, accompanied by a detrimental increase in the nose-down C m,peak, due to the large pressure increase on the lower surface of the flap, was also observed.

전투기 형상의 외부장착물이 꼬리날개에 미치는 영향에 대한 수치적 연구

본 연구에서는 삼차원 비정렬 비점성 유동 해석 코드를 이용하여 전투기 형상의 외부 장착물이 꼬리 날개에 미치는 공력 간섭효과에 대한 연구를 수행하였다. 수치적 기법으로는 격자점 중심(vertex-centered)에 기초한 유한체적법과 시간 적분을 위한 내재적인 point Gauss-Seidel 반복 계산법을 사용하였다. 해석 코드의 검증을 위해 전투기 형상에 대한 정상 유동 계산을 수행하였으며 결과를 실험 데이터와 비교하였다. 외부 장착물의 후류(wake)를 정확히 포착하기 위해 예상되는 후류 영역에 대해 국부적인 격자 조밀화를 수행하였으며 천음속 영역에서의 비정상 유동 해석을 통해 외부 장착물에서 발생하는 후류가 수평꼬리날개에 미치는 간섭효과를 확인하였다. 공력 간섭효과를 감소시키기 위한 대안으로는 뒷전 플랩 꺽임각(trailing edge flap deflection)을 고려하였으며 이에 대한 정량적인 감소효과를 제시하였다. A three-dimensional inviscid flow solver has been developed based on unstructured meshes for the investigation of the effect of the external stores on the tail surfaces of a generic fighter aircraft. The numerical method is based on a vertex-centered finite-volume discretization and an implicit point Gauss-Seidel time integration. The calculations were made for a steady flow and the computed results were compared with experimental data to validate the flow solver. An unsteady time-accurate computation of the generic fighter aircraft with external stores at transonic flight conditions showed that the external stores cause unsteady loading on the horizontal tail surface due to the mutual interference between their wake and the horizontal tail surface. It was shown that downward deflection of the trailing edge flap significantly reduces the undesirable interference effect.

Transonic Free-To-Roll Analysis of the F-35 (Joint Strike Fighter) Aircraft

The free-to-roll technique is used as a tool for predicting areas of uncommanded lateral motions. Recently, the NASA/Navy/Air Force Abrupt Wing Stall Program extended the use of this technique to the transonic speed regime. Using this technique, this paper evaluates various wing configurations on the Joint Strike Fighter (F-35) aircraft. The configurations investigated include leading- and trailing-edge flap deflections, leading-edge flap gap seals, and vortex generators. These tests were conducted in the NASA Langley 16-Foot Transonic Tunnel. The analysis used a modification of a figure of merit developed during the Abrupt Wing Stall Program to discern configuration effects. The results showed how the figure of merit can be used to schedule wing flap deflections to avoid areas of uncommanded lateral motion. The analysis also used both static and dynamic wind-tunnel data to provide insight into the uncommanded lateral behavior. The dynamic data were extracted from the time history data using parameter identification techniques. In general, sealing the gap between the inboard and outboard leading-edge flaps on the Navy version of the F-35 eliminated uncommanded lateral activity or delayed the activity to a higher angle of attack.

Swashplateless Helicopter Rotor System with Trailing-Edge Flaps for Flight and Vibration Controls

The objective of this study is to demonstrate the concept of active trailing-edge flaps as primary flight control and vibration reduction devices for a typical full-scale helicopter. A comprehensive rotorcraft analysis based on UMARC was developed to analyze the swashplateless rotor. A parametric study of various key design variables involved in the trailing-edge flap design was carried out. An optimal design of a trailing-edge flap system that provides effective control authority within the complete range of advance ratios as well as minimum actuation requirements was achieved. Trailing-edge flaps demonstrated the capability of performing both primary flight control and active vibration control functions. At a high forward speed (advance ratio of 0.32), the 4/rev vertical force and roll and pitch moments at hub are successfully eliminated (by 90%), and the 4/rev in-plane hub forces are reduced by more than 40%. The half peak-to-peak value of the trailing-edge flap deflection for primary flight control is 7.1 deg, and an additional 4.7 deg is required for active vibration control.