Roeland De Breuker

Morphing Wing Flight Control Via Postbuckled Precompressed Piezoelectric Actuators

The design, modeling, and testing of a morphing wing for flight control of an uninhabited aerial vehicle is detailed. The design employed a new type of piezoelectric flight control mechanism which relied on axial precompression to magnify control deflections and forces simultaneously. This postbuckled precompressed bending actuator was oriented in the plane of the 12% thick wing and mounted between the end of a tapered D-spar at the 40% chord and a trailing-edge stiffener at the 98% chord. Axial precompression was generated in the piezoelectric elements by an elastic skin which covered the outside of the wing and served as the aerodynamic surface over the aft 70 % of the wing chord. A two-dimensional semi-analytical model based on the Rayleigh-Ritz method of assumed modes was used to predict the static and dynamic trailing-edge deflections as a function of the applied voltage and aerodynamic loading. It was shown that static trailing-edge deflections of ±3.1 deg could be attained statically and dynamically through 34 Hz, with excellent correlation between theory and experiment. Wind tunnel and flight tests showed that the postbuckled precompressed morphing wing increased roll control authority on a 1.4 meter span uninhabited aerial vehicle while reducing weight, slop, part-count, and power consumption.

Post-buckled precompressed elements: a new class of control actuators for morphing wing UAVs

This paper describes how post-buckled precompressed (PBP) piezoelectric bender actuators are employed in a deformable wing structure to manipulate its camber distribution and thereby induce roll control on a subscale UAV. By applying axial compression to piezoelectric bimorph bender actuators, significantly higher deflections can be achieved than for conventional piezoelectric bender actuators. Classical laminated plate theory is shown to capture the behavior of the unloaded elements. A Newtonian deflection model employing nonlinear structural relations is demonstrated to predict the behavior of the PBP elements accurately. A proof of concept 100 mm (3.94 �� ) span wing employing two outboard PBP actuator sets and a highly compliant latex skin was fabricated. Bench tests showed that, with a wing chord of 145 mm (5.8 �� ) and an axial compression of 70.7 gmf mm −1 , deflection levels increased by more than a factor of 2 to 15.25 ◦ peak-to-peak, with a corner frequency of 34 Hz (an order of magnitude higher than conventional subscale servoactuators). A 1.4 m span subscale UAV was equipped with two PBP morphing panels at the outboard stations, each measuring 230 mm

A Generic Morphing Wing Analysis and Design Framework

A generic framework for morphing wing aeroelastic analysis and design is presented. The wing is discretised into an arbitrary number of wing segments. Two types of actuation mechanisms are identified: inter-rib mechanisms operating across a wing segment and intra-rib mechanisms acting between two adjacent wing segments. Virtually, any shape can be obtained by distributing four morphing modes over the entire morphing wing. Three are an intra-rib mechanism and one is an inter-rib mechanism. The intra-rib modes are wing shear, twist and extension, and the inter-rib mode is wing folding. The wing is modeled using a close coupling between a non-linear beam formulation and Weissinger aerodynamics. The framework is intended to aid quick preliminary design of morphing wings to trade-off contradictory requirements in a flight mission. The morphing wing can be optimized for discrete points in the flight mission, and for the entire flight mission. The framework can be used to predict aerodynamic performance, load distribution, aeroelastic deformations, and the required actuation forces and moments and corresponding actuation energy. Therefore, the performance gains of wing morphing can be weighed against the energy costs and weight penalties due to the presence of the actuators. The functionality of the framework is demonstrated by making use of a folding and sweeping wing test case.

Aeroelastic Control Using Distributed Floating Flaps Activated by Piezoelectric Tabs

In this paper, a novel aeroservoelastic effector configuration that is actuated by piezoelectric tabs is presented. The effector exploits trailing-edge tabs installed on free-floating flaps (FFFs). These flaps are used to prevent flutter from occurring and to alleviate loads originating from external excitations such as gusts. A vertical tailplane wind-tunnel model with two free-floating rudders and a flutter control mechanism were designed, and the aeroelastic stability and response characteristics have been modeled numerically. The controller uses the tailplane tip acceleration as a sensor and sends control signals to the piezoelectrically actuated tabs. Wind-tunnel experiments were performed to demonstrate the feasibility of the technology. It was demonstrated experimentally that the flutter speed associated with the free rudders could be increased by 80%. The same controller, applied to the external rudder, was used to alleviate the aeroelastic response of the tailplane to the excitation of the other ...

Flutter of Partially Rigid Cantilevered Two-Dimensional Plates in Axial Flow

The introduction of adaptive materials for active camber line shape control favors flexible wing designs, thus making adaptive wings more susceptible to instability phenomena. Motivated by a new wing concept for micro aerial vehicle applications, the aeroelastic stability of partially rigid cantilevered plates in an axial flow is investigated. The plate is modeled as a beam having a rigid and a flexible part. The beam is modeled using the classical Euler-Bernoulli bending theory, the unsteady aerodynamic pressure is modeled using Theodorsens theory, and the Rayleigh-Ritz method is used to obtain a discrete model. Stability analysis is carried out in Laplaces domain. The results indicate that a partially rigid cantilevered plate in an axial flow does not show static aeroelastic divergence but exhibits dynamic aeroelastic instability. The flutter velocity at which this instability occurs is dependent on the ratio of the flexible length to the total length of the plate and the mass ratio. Adding a rigid part ahead of a flexible plate can have a stabilizing or destabilizing effect on the aeroelastic behavior of the flexible plate, depending on the mass ratio. The phase-angle difference between the upstream and downstream part of the two-dimensional plate is shown to be dependent on the mass ratio and flexible length fraction. Jumps in the flutter diagram occur because of changes in flutter mode, and the flexible length fraction also affects these jump phenomena. The jumps are shown to be the result of eigenvalue branch collision. There are multiple critical flow conditions for mass ratios around the jumps. Therefore, a new practical flutter boundary definition is introduced to remove the overlap between flutter modes near the jumps. The flutter diagram calculated according to the new definition shows better agreement with published time domain simulations.

Optimal Control of Aeroelastic Systems using Synthetic Jet Actuators

Synthetic jet actuators (SJAs) are among the most promising devices when it comes to adaptive ∞ow control. These zero net mass ∞ux actuators periodically blow out and bring back ∞uid through an oriflce into a cavity located at an airfoil surface. They can not only add momentum to the boundary layer to improve its stability and delay separation or transition, but they also add momentum to the ∞ow, which leads to a rearrangement of the stream lines around the airfoil, with a global efiect of modifying the aerodynamic forces. Reduced Order Models (ROMs) based on Theodorsen theory have been developed to describe the efiect of such actuators on the global aerodynamic forces and moments acting on an airfoil. These analytical expressions allow for a rapid calculation of the static and dynamic aeroelastic response of the lifting body; as such, they are ideally suited to be incorporated into a feedback control loop. This paper aims to contribute to the optimal control of aeroelastic systems using SJAs by investigating the applicability of a linear quadratic regulator (LQR).

Post-buckled precompressed (PBP) subsonic micro flight control actuators and surfaces

This paper describes a new class of flight control actuators using Post-Buckled Precompressed (PBP) piezoelectric elements to provide much improved actuator performance. These PBP actuator elements are modeled using basic large deflection Euler-beam estimations accounting for laminated plate effects. The deflection estimations are then coupled to a high rotation kinematic model which translates PBP beam bending to stabilator deflections. A test article using PZT-5H piezoceramic sheets built into an active bender element was fitted with an elastic band which induced much improved deflection levels. Statically the bender element was capable of producing unloaded end rotations on the order of ±2.6°. With axial compression, the end deflections were shown to increase nearly 4-fold. The PBP element was then fitted with a graphite-epoxy aeroshell which was designed to pitch around a tubular stainless steel main spar. Quasi-static bench testing showed excellent correlation between theory and experiment through ±25° of pitch deflection. Finally, wind tunnel testing was conducted at airspeeds up to 120kts (62m/s, 202ft/s). Testing showed that deflections up through ±20° could be maintained at even the highest flight speed. The stabilator showed no flutter or divergence tendencies at all flight speeds. At higher deflection levels, it was shown that a slight degradation deflection was induced by nose-down pitching moments generated by separated flow conditions induced by extremely high angles of attack.

Flutter of Partially Rigid Cantilevered Plates in Axial Flow

The introduction of adaptive materials for active camber line shape change makes adaptive wings more susceptible to non conventional instability phenomena. Motivated by a new wing concept for micro aerial vehicle applications the aeroelastic stability of partially rigid cantilevered plates in an axial ∞ow is investigated. The plate is modeled as a beam having a rigid and a ∞exible part. The beam is modeled using classical Euler Bernoulli bending theory, the unsteady aerodynamic pressure is modeled using Theodorsen’s theory and the Rayleigh Ritz method is used to obtain a discrete model. Stability analysis is carried out in Laplace’s domain. The results indicate that a partially cantilevered plate in an axial ∞ow does not show static aeroelastic divergence but exhibits dynamic aeroelastic instability. The ∞utter velocity at which this occurs is dependent on the mass ratio and on the ratio of the ∞exible length to the total length of the plate. At certain mass ratios, jumps in the ∞utter speed occur due to changes in ∞utter mode. The ∞exible length fraction has a signiflcant efiect on the ∞utter speed and the jump phenomena.

Model Validation and Simulated Fatigue Load Alleviation of SNL Smart Rotor Experiment.

In this paper an individual flap controller (IFC) design for smart rotors is presented and compared to the model identification of the Sandia National Labs Smart Rotor experiment. The controller design has been carried out using an in-house aeroservoelastic software - DU_SWAT. The root bending moment response due to a step input of the flap deflection has been linearized. Based on this linearization a Coleman transform combined with a proportionality controller has been used to eliminate the cyclic components in the root bending moment. It was shown that IFC can reduce the 1P mode similar to individual pitch control, while only modest flap deflections of below 5

Energy-based aeroelastic analysis of a morphing wing

Aircraft are often confronted with distinct circumstances during different parts of their mission. Ideally the aircraft should fly optimally in terms of aerodynamic performance and other criteria in each one of these mission requirements. This requires in principle as many different aircraft configurations as there are flight conditions, so therefore a morphing aircraft would be the ideal solution. A morphing aircraft is a flying vehicle that i) changes its state substantially, ii) provides superior system capability and iii) uses a design that integrates innovative technologies. It is important for such aircraft that the gains due to the adaptability to the flight condition are not nullified by the energy consumption to carry out the morphing manoeuvre. Therefore an aeroelastic numerical tool that takes into account the morphing energy is needed to analyse the net gain of the morphing. The code couples three-dimensional beam finite elements model in a co-rotational framework to a lifting-line aerodynamic code. The morphing energy is calculated by summing actuation moments, applied at the beam nodes, multiplied by the required angular rotations of the beam elements. The code is validated with NASTRAN Aeroelasticity Module and found to be in agreement. Finally the applicability of the code is tested for a sweep morphing manoeuvre and it has been demonstrated that sweep morphing can improve the aerodynamic performance of an aircraft and that the inclusion of aeroelastic effects is important.