Multiple Trailing-edge Flaps Research Articles

On the Aeroelasticity of a Cantilever Wing Equipped with the Spanwise Morphing Trailing Edge Concept

This paper studies the aeroelastic behavior of a rectangular, cantilever wing equipped with the spanwise morphing trailing edge (SMTE) concept. The SMTE consists of multiple trailing edge flaps that allow controlling the spanwise camber distribution of a wing. The flaps are attached at the wing’s trailing edge using torsional springs. The Rayleigh–Ritz method is used to develop the equations of motion of the wing-flap system. The use of shape functions allows for representing the wing as an equivalent 2D airfoil with generalized coordinates that are defined at the wingtip. Strip theory, based on Theodorsen’s unsteady aerodynamic model, is used to compute the aerodynamic loads acting on the wing. A representative Padé approximation for Theodorsen’s function is utilized to model the aerodynamic behaviors in a state-space form allowing time-domain simulation and analysis. The model is validated using a rectangular cantilever wing and the data are available in the literature. A comprehensive parametric comparison study is conducted to assess the impact of flap stiffness on the aeroelastic boundary. In addition, the potential of the SMTE to provide load alleviation and flutter suppression is assessed for a wide range of flight conditions, using a discrete (1-cosine) gust. Finally, the implementation and validation of a controller for a wing with SMTE for gust load alleviation are studied and controller parameters are tuned for a specific gust model.

Aerostructural performance improvement in an unmanned long endurance aircraft using adaptive wing concept

This paper presents an analytical research study to improve the aerostructural performance of an unmanned medium altitude long endurance aircraft using the adaptive wing concept. Aerodynamic drag and wing root loads are minimized by optimal scheduling of multiple trailing edge flaps located on the wing. A trim optimization process is developed specifically for this purpose. The aeroelastic model is based on finite element formulation for the structure and doublet lattice method for the aerodynamics. A nonlinear numerical lifting line method is used, in combination with airfoil wind tunnel data, to estimate the induced and total drags. Results are presented for the current aircraft configuration and a more flexible proposed configuration, thereby providing an uncommon perspective on the effect of flexibility on the adaptive wing. For example, the benefits of optimal flap deployment turn out to be greater for the flexible aircraft than for the rigid one. It is hoped that this work and its insights will also aid the multidisciplinary design optimization of future aircraft.

Modal Identification of Aerodynamic Models for Spanwise Lift Distribution Control

This work presents modal identification results for the unsteady three-dimensional flow around a wing featuring multiple trailing edge flaps. The unsteady spanwise lift distribution is modeled using the unsteady vortex lattice method, which yields a linear, time-invariant, high-order state-space model. The eigensystem realization algorithm is applied for model order reduction and modal identification, providing aerodynamic mode shapes and associated eigenvalues. The wing is considered rigid, so that the dynamic mode shapes associated with the spanwise lift distribution are purely aerodynamic. It is observed that the spatial-temporal discretization of the numerical model and flap dimensions inherently limit the number of modes that can be properly identified. Open-loop simulation results demonstrate the modal behavior of the aerodynamic system. Finally, aerodynamic mode shapes for different wing planforms are also presented for comparison. By expanding the system output (lift distribution) as a truncated superposition of aerodynamic mode shapes, a low-order multiple-input and multiple-output modal representation can be obtained, suitable for controller synthesis. Ultimately, this novel approach will be applied to gust alleviation, in order to keep a desired lift distribution profile using shape control techniques.

Robust design of multiple trailing edge flaps for helicopter vibration reduction: A multi-objective bat algorithm approach

The objective of this study is to determine an optimal trailing edge flap configuration and flap location to achieve minimum hub vibration levels and flap actuation power simultaneously. An aeroelastic analysis of a soft in-plane four-bladed rotor is performed in conjunction with optimal control. A second-order polynomial response surface based on an orthogonal array (OA) with 3-level design describes both the objectives adequately. Two new orthogonal arrays called MGB2P-OA and MGB4P-OA are proposed to generate nonlinear response surfaces with all interaction terms for two and four parameters, respectively. A multi-objective bat algorithm (MOBA) approach is used to obtain the optimal design point for the mutually conflicting objectives. MOBA is a recently developed nature-inspired metaheuristic optimization algorithm that is based on the echolocation behaviour of bats. It is found that MOBA inspired Pareto optimal trailing edge flap design reduces vibration levels by 73% and flap actuation power by 27% in comparison with the baseline design.

Drag Reduction Methodology for Adaptive Tailless Aircraft

An approach is presented for determining optimum lift distributions for adaptive tailless aircraft. In this study, wing adaptation is achieved using multiple trailing-edge flaps that are used to optimally distribute the lift of the wing such that drag is minimized. For tailless aircraft that are stable in pitch, the lack of a secondary lifting surface makes it necessary that the lift distribution on the wing also satisfies a pitching-moment constraint to ensure trim. The current work implements a numerical approach that solves for the optimal scheduling of multiple trailing-edge flaps on the wing of a tailless aircraft for various flight conditions with a pitching-moment constraint to reduce both induced and profile drag. The approach uses superposition to construct the spanwise lift distribution from basic and additional loadings, and decomposes the flap-angle distribution into mean and variation distributions. Together, these elements enable the solution of the problem using semi-analytical methods that also provide insight. The results are presented for a planar, swept, tapered wing with two airfoil-section choices to verify the theory and provide insight for trade studies.

Identification of Flap Motion Parameters for Vibration Reduction in Helicopter Rotors with Multiple Active Trailing Edge Flaps

An active control method utilizing the multiple trailing edge flap configuration for rotorcraft vibration suppression and blade loads control is presented. A comprehensive model for rotor blade with active trailing edge flaps is used to calculate the vibration characteristics, natural frequencies and mode shapes of any complex composite helicopter rotor blade. A computer program is developed to calculate the system response, rotor blade root forces and moments under aerodynamic forcing conditions. Rotor blade system response is calculated using the proposed solution method and the developed program depending on any structural and aerodynamic properties of rotor blades, structural properties of trailing edge flaps and properties of trailing edge flap actuator inputs. Rotor blade loads are determined first on a nominal rotor blade without multiple active trailing edge flaps and then the effects of the active flap motions on the existing rotor blade loads are investigated. Multiple active trailing edge flaps are controlled by using open loop controllers to identify the effects of the actuator signal output properties such as frequency, amplitude and phase on the system response. Effects of using multiple trailing edge flaps on controlling rotor blade vibrations are investigated and some design criteria are determined for the design of trailing edge flap controller that will provide actuator signal outputs to minimize the rotor blade root loads. It is calculated that using the developed active trailing edge rotor blade model, helicopter rotor blade vibrations can be reduced up to 36% of the nominal rotor blade vibrations.

Identification of Flap Motion Parameters for Vibration Reduction in Helicopter Rotors with Multiple Active Trailing Edge Flaps

An active control method utilizing the multiple trailing edge flap configuration for rotorcraft vibration suppression and blade loads control is presented. A comprehensive model for rotor blade with active trailing edge flaps is used to calculate the vibration characteristics, natural frequencies and mode shapes of any complex composite helicopter rotor blade. A computer program is developed to calculate the system response, rotor blade root forces and moments under aerodynamic forcing conditions. Rotor blade system response is calculated using the proposed solution method and the developed program depending on any structural and aerodynamic properties of rotor blades, structural properties of trailing edge flaps and properties of trailing edge flap actuator inputs. Rotor blade loads are determined first on a nominal rotor blade without multiple active trailing edge flaps and then the effects of the active flap motions on the existing rotor blade loads are investigated. Multiple active trailing edge flaps are controlled by using open loop controllers to identify the effects of the actuator signal output properties such as frequency, amplitude and phase on the system response. Effects of using multiple trailing edge flaps on controlling rotor blade vibrations are investigated and some design criteria are determined for the design of trailing edge flap controller that will provide actuator signal outputs to minimize the rotor blade root loads. It is calculated that using the developed active trailing edge rotor blade model, helicopter rotor blade vibrations can be reduced up to 36% of the nominal rotor blade vibrations.

Ideal Lift Distributions and Flap Angles for Adaptive Wings

An approach is presented for determining the optimum flap angles and spanwise loading to suit a given flight condition. Multiple trailing-edge flaps along the span of an adaptive wing are set to either reduce drag in rectilinear flight conditions or to limit the wing bending moment at maneuvering conditions. For reducing drag, the flaps are adjusted to minimize induced drag, while simultaneously enabling the wing sections to operate within their respective low-drag ranges. For limiting wing bending moment, the flaps are used to relieve the loading near the wing tips. An important element of the approach is the decomposition of the flap angles into a distribution that can be used to control the spanwise loading for induced-drag control and a constant flap that can used for profile-drag control. The problem is linearized using the concept of basic and additional lift distributions, which enables the use of standard constrained-minimization formulations. The results for flap-angle distributions for different flight conditions are presented for a planar and a nonplanar wing. Postdesign analysis and aircraft-performance simulations are used to validate the optimum flap-angle distributions determined using the current approach.

Using the Complete Authority of Multiple Active Trailing-edge Flaps for Helicopter Vibration Control

In the current work, the application and advantages of multiple active trailing-edge flaps for helicopter vibration control in forward flight are investigated. Small active flaps of this type can be actuated using smart materials. A comprehensive aeroelastic analysis is used to model a helicopter with multiple trailing-edge flaps. The objective function containing the hub vibratory loads and flap control effort is minimized by an optimal controller. The control effort is differentially weighted for different flaps, such that all trailing-edge flaps are exploited to their full authority and limited to ±2 degrees peak deflection. Numerical results are obtained for one, two and three flaps and their optimal locations identified. A reduction of 76% in the hub vibration level is achieved at an advance ratio (nondimensional forward speed) of 0.3 using three flaps. For the four-bladed, soft-inplane, hingeless rotor considered in this work, it is observed that the extent of hub vibration reduction is directly related to the effect of trailing-edge flap motion on the contribution of the second flapwise bending mode to blade motion. It is also shown that multiple flaps can effectively modify the participation of second flapwise bending mode at a much lower control power than is needed with single and dual flap configurations. Furthermore, multiple flaps are shown to allow redundancy in operation and are considerably more effective than single flaps in low speed transition flight regimes because of their capability in distributing the airloads.

Dynamic hysteresis of piezoceramic stack actuators used in helicopter vibration control:experiments and simulations

This study investigates the effect of piezoceramic actuator hysteresis on helicoptervibration control using dual trailing-edge flaps. Piezoceramic stack actuators arepromising candidates for trailing-edge flap actuation in full-scale helicopters. However,they are inherently nonlinear in response to an applied electric field and exhibithysteretic behavior between the applied electric field and displacement. In this study,bench-top tests are conducted on a commercially available piezoceramic stackactuator, and its dynamic hysteretic behavior is studied. A Preisach-type dynamichysteresis model is used to describe the hysteresis in the stack actuator. The unknowncoefficients in the model are obtained through identification from experimental data.An aeroelastic model of the helicopter with multiple trailing-edge flaps is thenused to predict the hub vibration levels under different flight conditions. Theoptimal actuator control input for hub vibration suppression in the presence ofactuator hysteresis is considerably different from the case of an ideal-linear actuator.Numerical results show the importance of modeling actuator hysteresis in helicoptervibration control using trailing-edge flaps. Ignoring or inaccurate modeling ofhysteresis in the piezoceramic actuator can lead to inaccurate phasing of thetrailing-edge flap motion which directly affects the performance of the vibration controlsystem.