Flexible Aircraft Research Articles

An Analysis Method of Static Aeroelastic Response with Nonlinear Aerodynamic Data on 3D Bodysurface Panel Element

A static aeroelastic responses analysis method based on mapping external nonlinear aerodynamic forces onto 3D body surface panel elements is proposed. The external nonlinear aerodynamic force data can be obtained by Euler solver, Navier-Stokes solver and wind tunnel tests. In this paper, the method is applied to obtain static aeroelastic responses of a flexible aircraft, and the emphasis is placed on the change of aerodynamic coefficient, aerodynamic pressure distribution and flight load. The results are compared with static aeroelastic responses calculated by high-order panel method. It is concluded that the method presented in this study is feasible, credible and efficient, and can provide an assessment of static aeroelastic response characteristics for the detailed stage of aircraft design.

Reduced-Order <i>H<sub>∞</sub> </i>Controller Design Using the <i>μ</i>-Synthesis Applied to Lateral Control of Highly Flexible Aircraft

A lateral flight reduced-order controller for a flexible aircraft is presented. It is based on the reduced-order model of the flexible aircraft and theμ-analysis and synthesis toolbox. The first step deals with the order reduction of the full-model by using the balanced-truncation model reduction method. Then the reduced-orderHcontroller is achieved by using theμ-analysis and synthesis theory and applied to the reduced-order model and full-order model of the flexible aircraft. Finally, numerical results are presented and discussed. The simulation results show the efficiency of the reduced-order controller based on the reduced-order model since the close-loop system of the full-order model meets a set of realistic specifications in the frequency and time domains.

Optimal linear quadratic model following with an application to a flexible aircraft

This paper develops general theoretical results about an input–output model-following methodology for linear systems, as an optimal control problem. A control law is obtained by minimizing a quadratic index that takes into account the matching errors and the control inputs. The control is obtained from the Lagrange multiplier method and can be interpreted as an extension of the linear quadratic regulator, with finite and infinite horizon formulations. The major contribution of the paper is the development of solutions involving plant output feedback. The method is illustrated with an application to a nonlinear flexible aircraft with nonstationary aerodynamics and nine flexible modes. Simulations compare state and output feedback solutions. In the proposed example, when taking into account unmodeled flexible dynamics and parametric uncertainties, the best results are given by the proposed output feedback.

Final Approach and Flare Control of a Flexible Aircraft in Crosswind Landings

In this paper, the flight simulation and control of a flexible aircraft in landing is presented. The discrete form of hybrid-state equations of motion in terms of quasi coordinates takes into account the coupling between elastic and rigid-body degrees of freedom and provides the framework for the controller design and simulation. The wings of the aircraft are modeled as cantilever beams undergoing bending and torsion about their elastic axis. Distributed variables are discretized in space using the Galerkin method, and unsteady aerodynamic forces are computed using strip theory and finite state induced-flow theory. A landing condition in which the elastic aircraft encounters a crosswind all the way before touchdown is considered. To design the control system, nonlinear coupled equations of motion are linearized about the trim condition and separated into two sets of decoupled equations, in which the elastic variables affect lateral-directional equations. The landing phase is divided into the final approach and flare phases and, for each phase, separate controllers are designed for longitudinal- and lateral-directional channels. An optimal-integral feedforward control scheme is implemented to accomplish the autolanding, while suppressing the wind effects on the flight path. The performance of the control system is examined through nonlinear simulation.

Nonlinear Gust Response Analysis of Free Flexible Aircraft

Gust response analysis plays a very important role in large aircraft design. This paper presents a methodology for calculating the flight dynamic characteristics and gust response of free flexible aircraft. A multidisciplinary coupled numerical tool is developed to simulate detailed aircraft models undergoing arbitrary free flight motion in the time domain, by Computational Fluid Dynamics (CFD), Computational Structure Dynamics (CSD) and Computational Flight Mechanics (CFM) coupling. To achieve this objective, a structured, time-accurate flow- solver is coupled with a computational module solving the flight mechanics equations of motion and a structural mechanics code determining the structural deformations. A novel method to determine the trim state of flexible aircraft is also stated. First, the field velocity approach is validated, after the trim state is attained, gust responses for the one-minus-cosine gust profile are analyzed for the longitudinal motion of a slender-wing aircraft configuration with and without the consideration of structural deformation. RBA), while the second aims mainly to the analysis of elastic aircraft experiencing relatively small deformations. The idea to not consider cross-coupling effects between these two disciplines has been commonly justified by the large frequency separation of the characteristic motion which is typical of conventional structures. Nowadays, the focus on weight minimization for aircraft, leads toward more and more flexible vehicles. The resulting underlying structures may not exhibit the usual wide frequency separation among the rigid body degrees of freedom and the remaining elastic modes. So that the approach mentioned above can lead to mistakes/errors in analyses of flight performance, flying qualities, and control systems design. In these cases, an integrated analysis of flight mechanics and aeroelasticity is necessary from the very early stages of preliminary design (6).

Robust Gust Alleviation and Stabilization of Very Flexible Aircraft

Robust linear control, combined with model-reduction methodologies, is investigated for gust rejection on large, very flexible aircraft using trailing-edge control surfaces. Controllers are designed on linearizations of the nonlinear aeroelastic equations, and the closed-loop response of both the linearized and nonlinear system to discrete gust distributions is compared to the open-loop dynamics. Results show that an controller performs well on a relatively large linearized system, with 9% load alleviation in root bending moments for the critical gust length. When applied to the nonlinear model of the same vehicle, the robust controller gives a good performance in response to short gusts, including the critical length, with even better load reductions than the linear case. However, this performance gap decreases as the gust length is increased. It is also shown how standard model-reduction techniques can provide metrics for the selection of a minimum size of the aeroelastic system. Finally, it is shown that both the full and reduced controllers are capable of stabilizing a large-amplitude phugoid mode on the nonlinear model of the vehicle.

Frequency and Time Domain Recursive Parameter Estimation For a Flexible Aircraft

Aerodynamic parameter estimation is an integral part of aerospace system design and life cycle process. Recent advances in computational power have allowed the use of online parameter estimation techniques in various applications. In this paper, two online parameter estimation methods are considered for the identification of aerodynamic derivatives of a flexible aircraft: Equation Error Method and Kalman Filtering Method. The performance of the two algorithms is compared in terms of accuracy, computational efficiency and convergence. Simulation results indicate that Kalman filtering method has good convergence property and yields parameter estimates with better accuracy compared to the equation error method.

Dynamic Landing Response Analysis of a Flexible Tailless Delta Wing Aircraft

Modern tailless delta wing aircraft is built for high performance with light weight leading to structural flexibility. During landing, high vertical sink rate of aircraft causes peak transient loads in and around the landing gear attachment locations. These reactions forms the load cases for local sizing of airframe and its responses determine the overall integrity. This paper describes the co-simulation approach to carry out the dynamic landing analysis on a flexible tailless delta wing aircraft and examines the responses at certain critical locations such as the tip of wings, fin and nose for nominal landing conditions. The reactions are obtained from the Landing gear supplier through MSC Adams multi-body dynamic simulation of full aircraft with non-linear stiffness and damping models of Nose and Main landing gear systems. The landing response analysis is carried out on a flexible aircraft using industry standard MSC Nastran as transient dynamic problem with base enforced motion. It is observed that the...

Dynamics and Control of a Flexible Flapping Wing Aircraft

To investigate the aerodynamic performance of a flexible flapping wing aircraft, a flapping-wing system were design and an experiment were set up to measure the unsteady aerodynamic forces of the flapping motion. The thrust formula and resistance formula described aerodynamic forces. The lift and thrust of this mechanism were measured for different angles of attack and wind velocities. Results indicate that the thrust increases with the flapping frequency and the lift increase with the wind velocity, while the lift coefficient decreases while the velocity increases. It is realized that the wing’s transformation which imitated birds leads less resistance when flapping upward which impacts the aerodynamic lift generation and the bionic winglet leads to a change in the leading edge vortex and span-wise flow structures, which decrease the airflow’s backward pull. Models were introduced which were used in the design process and show its aerodynamic performance. The flexible flapping wing vehicle is still an open research area.

Cruising Autopilot for a Flexible Aircraft with an Internal Loop of Model Following

AbstractA cruising autopilot is developed for a nonlinear flexible aircraft based on the model-following method. The autopilot controller, called the outer loop, is designed for a rigid body approximation of the flexible aircraft. The model-following control, named the inner loop, approximates the input-output behavior of the flexible aircraft to a rigid body. The autopilot is applied to the flexible aircraft with a model-following inner loop. This inner loop controls the flexible modes and the outer loop, which are the rigid body dynamics. The model-following control is a suboptimal linear quadratic formulation with output feedback. This technique can stabilize the flexible modes and keep their amplitudes acceptable. The autopilot is a speed and altitude tracker with a wing leveler. It is designed via a robust H∞ with static output feedback. An aircraft model with nine flexible modes is used to evaluate the method. It is a conceptual medium size jet, like an Embraer 190/190 or Boeing 737–200/300. Simulat...

Assessment of Wake-Tail Interference Effects on the Dynamics of Flexible Aircraft

This work investigates the effect of aerodynamic interference in the coupled nonlinear aeroelasticity and flight mechanics of flexible lightweight aircraft at low speeds. For that purpose, a geometrically exact composite-beam formulation is used to model the vehicle flexible-body dynamics by means of an intuitive and easily linearizable representation based on the displacement and Cartesian rotation vectors. The aerodynamics are modeled using the unsteady vortex-lattice method, which captures the instantaneous shape of the lifting surfaces and the free inviscid wake, including large deformations and interference effects. This results in a framework for simulation of high aspect ratio planes that provides a medium-fidelity representation of flexible-aircraft dynamics with a modest computational cost. Previous independent studies on the structural-dynamics and aerodynamics modules are complemented here with the integrated simulation methodology, including vehicle trim, and linear and nonlinear time-domain solutions. A numerical investigation is next presented on a simple wing-fuselage-tail configuration, assessing the interference effects between wing wake and horizontal tail, and the downwash due to the proximity of the wake is shown to play a significant role in the longitudinal dynamics of the vehicle. Finally, a brief discussion of direct wake-tail encounters is included to show the limitations of the approach.

Generation of a reduced-order LPV/LFT model from a set of large-scale MIMO LTI flexible aircraft models

In the civilian aviation industry, the aeroelastic behavior of an aircraft is often modeled at frozen flight and mass configurations using high fidelity numerical tools. Unfortunately, the resulting large-scale models cannot be handled in such form by modern analysis and control techniques, which generally require the considered models to be written as low-order Linear Fractional Representations (LFR). In this context, a methodology is described to derive a reduced-order Linear Parameter Varying (LPV) model from a reference set of large-scale Multiple Input Multiple Output (MIMO) Linear Time Invariant (LTI) models describing a given system at frozen configurations. The proposed approach is in two steps. The reference models are first reduced using recent advances in Krylov methods, leading to a set of low-order state-space representations with consistent state vectors. An LPV model is then obtained by polynomial approximation and converted into an LFR of reasonable size. A special effort is made to avoid data overfitting by using as simple as possible approximation formulas. The method is applied to a long-range commercial aircraft model developed in an industrial context: a set of large-scale flexible models linearized at different mass configurations is converted into a single low-order LPV model. More generally, any kind of purely numerical models for which the analytical structure is unknown can be considered.

Aeroelastic trim and flight loads analysis of flexible aircraft with large deformations

A method for static aeroelastic trim analysis and flight loads computation of a flexible aircraft with large deformations has been presented in this paper, which considers the geometric nonlinearity of the structure and the nonplanar effects of aerodynamics. A nonplanar vortex lattice method is used to compute the nonplanar aerodynamics. The nonlinear finite element method is introduced to consider the structural geometric nonlinearity. Moreover, the surface spline method is used for structure/aerodynamics coupling. Finally, by combining the equilibrium equations of rigid motions of the deformed aircraft, the nonlinear trim problem of the flexible aircraft is solved by iterative method. For instance, the longitudinal trim analysis of a flexible aircraft with large-aspect-ratio wings is carried out by both the nonlinear method presented and the linear method of MSC Flightloads. Results obtained by these two methods are compared, and it is indicated that the results agree with each other when the deformation is small. However, because the linear method of static aeroelastic analysis does not consider the nonplanar aerodynamic effects or structural geometric nonlinearity, it is not applicable as the deformations increase. Whereas the nonlinear method presented could solve the trim problem accurately, even the deformations are large, which makes the nonlinear method suitable for rapid and efficient analysis in engineering practice. It could be used not only in the preliminary stage but also in the detail stage of aircraft design.

Structural and Aeroelastic Design of a Joined-Wing UAV

The Italian Aerospace Research Center (CIRA) is currently designing an unmanned aerial research system that is lightweight and has high-structural flexibility, code named the High Altitude Performance Demonstrator (HAPD).The project is framed within the Italian Aerospace Research Program, under the Unmanned Aerial Vehicle (UAV) Chapter. This unmanned aerial system is mainly aimed at developing and validating advanced modeling methodologies for flexible aircrafts. A compendium of the system is provided in this paper, together with a deeper discussion of how CIRA developed the structural and aeroelastic design of HAPD. Some experimental tests performed to validate the main concepts are also presented. The vehicle has an unconventional joined-wing configuration that mitigates undesired extreme flexibility, but that results in a more complicated design. First, aeroelasticity has been taken into account from the preliminary stages of design because flexibility significantly affects aircraft behavior. Second, the HAPD structure is redundant with regard to constraints (because of its joined wing), thus making the internal forces dependent on the stiffness distribution. For these reasons, the availability of an integrated methodology that can support the structural design is mandatory. The output of such a methodology consists of primary structure stiffness distributions (fuselage, wings, and vertical tail), compatibly with the absence of any aeroelastic instability, and structural failure under operative loads.

Dynamics and Performance of Tailless Micro Aerial Vehicle with Flexible Articulated Wings

The purpose of this paper is to analyze and discuss the performance and stability of a tailless micro aerial vehicle with flexible articulated wings. The dihedral angles can be varied symmetrically on both wings to control the aircraft speed independently of the angle of attack and flight-path angle, while an asymmetric dihedral setting can be used to control yaw in the absence of a vertical tail.Anonlinear aero-elastic model is derived, and it is used to study the steady-state performance and flight stability of the micro aerial vehicle. The concept of the effective dihedral is introduced, which allows for a unified treatment of rigid and flexible wing aircraft. It also identifies the amount of elasticity that is necessary to obtain tangible performance benefits over a rigid wing. The feasibility of using axial tension to stiffen the wing is discussed, and, at least in the context of a linear model, it is shown that adding axial tension is effective but undesirable. The turning performance of an micro aerial vehicle with flexible wings is compared to an otherwise identical micro aerial vehicle with rigid wings. The wing dihedral alone can be varied asymmetrically to perform rapid turns and regulate sideslip. The maximum attainable turn rate for a given elevator setting, however, does not increase unless antisymmetric wing twisting is employed.

Aeroservoelastic Design Optimization of a Flexible Wing

In this paper a multidisciplinary design optimization framework is developed that integrates control system design with aerostructural design. The equations of motion are derived for a flexible aircraft and used to perform aeroservoelastic analysis. The objective of this framework is to go beyond the current limits of aircraft performance through simultaneous design optimization of aerodynamic shape, structural sizing and control system. The control system uses load alleviation to reduce the critical structural loads. Time-domain analysis of the aircraft performing an altitude change maneuver and encountering an atmospheric gust is included in the design process. The optimal trade-off between aerodynamics, structures and control system is found by maximizing the endurance subjected to stress and maneuverability constraints. Two cases — with and without load alleviation system — are considered. Due to the proposed MDO framework, the inclusion of load alleviation system in design leads to a significant increase in endurance performance.

Identification of Aerodynamic Derivatives of a Flexible Aircraft

THE effects of flexibility on the flight dynamics of large transport aircraft have been shown to be quite significant, especially as the frequencies of the elastic modes become lower and approach those of the rigid body modes [1]. The handling characteristics of such vehicles are altered significantly from those of a rigid vehicle, and the design of the flight-control system may become drastically more complex. Therefore, mathematical modeling of an aeroelastic vehicle for dynamic analysis and control system design is a major issue in flight vehicle dynamics. Consequently, the need to model accurately the dynamics of such vehicles is becoming inevitable. Parameter estimation from flight data, as applied to aircraft in the linear flight regime, is currently being used on a routine basis with the assumption that the rigid body model is valid [2]. Elastic degrees of freedom are, therefore, absent from the aircraft derivative model used in the estimation algorithm. However, the newly introduced highly maneuverable aircraft with a high degree of flexibility poses a new challenge to search for appropriate aeroelastic models for inclusion in parameter estimation algorithms [3]. A simplified integrated modeling approach to account for the aeroelastic effects in aircraft dynamics was suggested in [4], and its use for a highly elastic aircraft was demonstrated. The aeroelastic model of [4] was simplified in [5] and the resulting model, with a reduced number of unknown terms, was used in the identification of the aeroelastic aircraft. In another attempt, parameter estimation from simulated data of a flexible aircraft was carried out by defining the aerodynamic derivatives in a linear form with the quasi-steady aeroelastic effects represented through the flex factor [6].

Optimal sensors placement and spillover suppression

A new approach to optimal placement of sensors (OSP) in mechanical structures is presented. In contrast to existing methods, the presented procedure enables a designer to seek for a trade-off between the presence of desirable modes in captured measurements and the elimination of influence of those mode shapes that are not of interest in a given situation. An efficient numerical algorithm is presented, developed from an existing routine based on the Fischer information matrix analysis. We consider two requirements in the optimal sensor placement procedure. On top of the classical EFI approach, the sensors configuration should also minimize spillover of unwanted higher modes. We use the information approach to OSP, based on the effective independent method (EFI), and modify the underlying criterion to meet both of our requirements—to maximize useful signals and minimize spillover of unwanted modes at the same time. Performance of our approach is demonstrated by means of examples, and a flexible Blended Wing Body (BWB) aircraft case study related to a running European-level FP7 research project ‘ACFA 2020—Active Control for Flexible Aircraft’.

Design of a 450-passenger blended wing body aircraft for active control investigations

ACFA 2020 (active control for flexible aircraft) is a collaborative research project funded by the European Commission under the seventh research framework programme. The project deals with innovative active control concepts for 2020 aircraft configurations like the blended wing body (BWB) aircraft. The main objectives of ACFA are the design of a new ultra efficient 450 passenger BWB type aircraft as well as the provision of robust adaptive multi- channel control architectures for such aircraft. The objective of the newly designed controllers is an ambitious improvement of ride comfort and handling qualities, as well as load reduction on BWB type aircraft. Based on the attained load reduction, the structure of the 450-passenger aircraft can be resized with the goal of an ambitious weight saving for further improvement of fuel efficiency. Active control requirements influence the design process of control surfaces and overall aircraft design, respectively. Hence, the conventional aircraft design process had to be adapted to the new requirements. The aircraft design framework described in this article has proven its efficiency in the development of the ACFA BWB aircraft. Within a time period of 1 year, an airframe has been developed under the constraints of multiple domain requirements. This article presents the process and the results of the design activities of the BWB aircraft, which form the basis for a detailed concept analysis as well as the investigation of multiple-input multiple-output control architectures.

A New Method for Modeling of Fully Flexible Aircrafts

The purpose of this study is to produce a modeling capability for integrated flight dynamics of flexible aircraft that can better predict some of the complex behaviors in flight due to multi-physics coupling. Based on the studying of the exiting modeling approaches, the author put forward a new modeling method, and developed a new formulation integrating nonlinear rigid-body flight mechanics and linear aeroelastic dynamics for fully elastic aircrafts using Lagrangian mechanics. The new equations of motion overcome the disadvantages of the exiting methods, and include automatically all six rigid-body degrees of freedom and elastic information, the seamless integration is achieved by using the same reference frame and the same variables to describe the aircraft motions and the forces acting on it, including the aerodynamic forces. The formulation is modular in nature, in the sense that the structural model, the aerodynamic theory, and the controls method can be replaced by any other ones to better suit different types of aircraft.