© Universal Technology Corporation, 2018.There are many challenges related to the design and operation of flexible aircraft. Flight control law design for improving handling qualities is one of them because the major problem in the design of controllers then concerns aeroservoelastic stability. To deal with this difficulty, methodologies for flight control law design considering the aeroelastic dynamics of the aircraft are being pursued. In this paper, an output-feedback-based stability augmentation system is proposed and designed to improve the handling qualities of a flexible aircraft and to increase the structural damping of some of its aeroelastic modes. The design is based on the projective control technique, which allows preserving in the closed-loop system the eigenstructure of certain modes of interest whose dynamic characteristics stem from an optimal state feedback solution. The X-HALE flexible aircraft is considered. Numerical studies by means of nonlinear simulations allow testing and confirming the effectiveness of the proposed controller.

© Universal Technology Corporation, 2018.The development of modern, more efficient transport aircraft and High-Altitude-Long-Endurance (HALE) aircraft requires solutions that involve lightweight structures and high aspect ratio wings for maximum aerodynamic efficiency. However, this may result in flexible wings, such that the coupling between rigid-body and structural modes deteriorates stability and handling qualities of the aircraft. These can be improved by the continuous operation of control systems with the use of structural information, such as wing deformed shape. Strain gages can be used for such an application. However these sensors may have input interference, such as temperature variation, that can cause false strain responses other than the expected measurements due to flight loads. This work presents the analyzes of strain gage measurements due to thermal loads in the X-HALE aircraft wing. Three thermal correction methods were proposed: functions of temperature, dummy gage, and high-pass filtering. Each was tested on indoor and outdoor experiment setup, which uses a lamp and natural temperature variation throughout a full day as heat sources. All methods were effective in removing thermal drift of indoor experiment data. However, only high-pass filtering method was successful in the outdoor experiment. Indoor thermal test results done on the full 4m X-HALE aircraft, with the addition of impact excitations on the aircraft’s wing tip, showed that the high-pass filtering was able to remove thermal drift while maintaining all responses in the frequency range of interest. Therefore, this method is suitable to the present application.

© Universal Technology Corporation, 2018.Biodynamic feedthrough is a phenomenon in which structural vibrations on the cockpit are fedback to the pilot, who transmits those vibrations on the aircraft through involuntary inceptor displacements. In the case these commands lead to instabilities, the phenomenon is then called Pilot-Assisted (or Augmented) Oscillations (PAO). This is a complex phenomenon that depends on three basic elements. First element is the aeroelastic modes of the aircraft, especially their modal shapes (since only specific modal shapes interfere in this phenomenon) and frequencies. Second one is the inceptor’s system characteristics, such as natural frequency and damping. The last element is the human pilot dynamics, which may be modelled as a passive spring-mass-damper system. This paper aims to explore how these three elements affect the development of the PAO phenomenon for more flexible aircraft. The investigations herein are accomplished for a virtual, flexible aircraft by the analysis of pilot-in-the-loop simulations of a high gain maneuver, varying airframe elasticity levels, inceptor system parameters and pilot model characteristics.

© Universal Technology Corporation, 2018.Formulations for the flight dynamics of flexible aircraft have been commonly applied to aircraft having all six rigid-body degrees of freedom unconstrained. However, during takeoff and landing, the aircraft motion becomes constrained by the ground. An intricate dynamical problem arises when the flexible aircraft has multiple underwing pod-mounted landing gears. During liftoff, the most outboard landing gears become airborne earlier than the most inboard ones. The equations of motion used in a simulation model must then include time-varying ground constraint forces. Aerodynamically, the ground effect becomes relevant and can be modeled using the method of images. Wheel-to-ground rolling resistance also needs to be considered. This paper aims at the derivation of equations of motion that allow simulating the takeoff of such complex aircraft configurations. The developed formulation is important because it is readily applicable also to the landing phase, allowing the calculation of landing loads and the design of takeoff and landing control systems with less uncertainties. As spinoffs, the stability analysis of flexible aircraft in ground effect and evaluation of takeoff performance are also made possible.

© Universal Technology Corporation, 2018.An analysis of the flight dynamics of flexible aircraft having high-aspect-ratio wings shows that, even if the elastic deformations are not large enough to incur structural-dynamic geometrical nonlinearities, important aerodynamic effects can arise that are geometrically nonlinear in essence. For instance, the dihedral effect of the deformed wing is usually significantly different from that of the undeformed one, leading to unacceptable inaccuracies when analyzing the aircraft response to side gusts. Classical approaches to the aeroelastic modeling of flexible aircraft, using geometrically-linear finite elements, the vortex- or doublet-lattice methods, and linear or surface spline interpolation techniques, are unable to represent aerodynamic geometrical nonlinearities, because all the aeroelastic model matrices are calculated a priori for the undeformed aircraft and remain unchanged in the analysis. However, numerical experimentation indicates that, if small deformations occur, then the aerodynamic geometrical nonlinearities can be approximately modeled with an on-line update of the spline matrices to take into account the instantaneous deformed normal directions, without the need to perform the much costlier aerodynamic mesh deformation. This paper aims at the derivation of the equations for the modified generalized aerodynamic forces and at the validation of the proposed method in both static and dynamic conditions.

© Universal Technology Corporation, 2018.Previous works evidenced stability problems associated with flight control law design for flexible aircraft. In this paper, a fuzzy-based gain-scheduling approach is proposed to adequate closed-loop response. An interactive method that aims performance improvement while enforcing global stability considering fuzzy gain-scheduling was proposed. The application of the technique was demonstrated for the flexible X-HALE aircraft nonlinear model and compared to classical interpolation-based gain-scheduling techniques. Results revealed that fuzzy-based gain-scheduling is promising for flexible aircraft control.

© 2019 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.Control law design for flexible aircraft with coupled structural and rigid-body dynamics is currently a challenge. This paper aims to improve the dynamical characteristics of a flexible aircraft by proposing a performance index for stability augmentation system design. The theoretical formulation will be presented, comparisons of numerical results between open-loop and closed-loop systems will be made and, finally, experimental data obtained via flight tests will be shown.

© 2019 by Daniel Drewiacki, Flávio José Silvestre, and Antônio Bernardo Guimarães Neto.The advent of fly-by-wire technology brought many advantages to aircraft design, but also increased the number of occurrences of the undesirable pilot-induced oscillation phenomena. To analyze such phenomena, handling qualities criteria have been established, mostly in the frequency domain. In the time domain, dynamic models to represent pilot behavior have been proposed for pilot-in-the-loop analysis and simulations. Both cases assume the aircraft as a rigid body. For more flexible aircraft, this approximation is no longer valid, and aeroelastic dynamics may interact with both flight dynamics and pilot body dynamics, possibly leading to the occurrence of other issues, such as biodynamic feedthrough. In this paper, the importance of considering the influence of airframe flexibility in the study of the pilot-induced oscillation phenomena is addressed by the application of handling qualities criteria using a flexible aircraft dynamic model.

© 2017, The Author(s) 2017.The main concern related to the flutter phenomenon is predicting and avoiding it. This paper describes the application of a flexural-torsional flutter testbed for acceleration reduction by applying active and passive model-based control. The model consists of the 2D typical section, with aerodynamic loads estimated by an unsteady time-domain formulation based on Wagner’s function. The active control architecture consists of a stability augmentation system with output feedback and gain scheduling via the linear-quadratic regulator theory and actuation by servomechanism. The passive control employs a shape-memory alloy to provide additional torsional stiffness. Experimental results show considerable reduction of oscillations at a relative low cost for both active and passive control strategies, and that the use of shape memory alloys in aeroelastic stability problems is promising.

© 2018 Elsevier Masson SASAircraft point-mass equations of motion have been largely adopted to calculate optimal trajectories with local aerodynamic models, i.e. valid in a restricted domain. However, some optimal maneuvers may need aerodynamic models valid for a broader range of flight conditions. For this purpose, global aerodynamic models are attractive but their nonlinear structure can preclude obtaining optimal trajectories by an indirect method together with the point-mass equations of motion. To solve this impasse without resorting to direct methods the authors propose a new set of aircraft equations of motion. When compared to the point-mass equations, the proposed set permits the inclusion of the angular velocity in the evaluation of aerodynamic forces, making them more accurate. Another advantage of the proposed model over the point-mass one is that it allows a qualitative estimate of the control surface deflections after the trajectory is obtained, which enables to discard solutions with infeasible deflections. To verify consistency, the proposed equations of motion are compared by simulation to the point-mass and to the rigid-body equations. The use of the proposed set of equations is demonstrated by three optimizations of a 360∘ roll problem.

© 2018, American Institute of Aeronautics and Astronautics Inc, AIAA. All right reserved.The interest in the class of unmanned vehicle known as High Altitude Long Endurance aircraft has been growing in the latest years. In this paper, stability problems associated with flight control law design for flexible aircraft are evidenced. A fuzzy-based gain-scheduling approach is proposed to adequate closed-loop response. The application of the technique was demonstrated for the flexible X-HALE aircraft and compared to classical interpolation-based gain-scheduling techniques. Nonlinear simulation results revealed that fuzzy-based gain-scheduling is promising for flexible aircraft control.

© 2018, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved.PIO (Pilot-Induced Oscillations) phenomena have been widely explored over the past recent years and led to the development of mathematical models to describe human pilot behavior as well as frequency domain analysis criteria. However, both these pilot models and prediction criteria have been developed neglecting the influence of airframe flexibility. Since aircraft are becoming more flexible, this influence has to be considered during the development of fly-by-wire systems. In this paper, we address the effects of aeroelastic dynamics in aircraft handling qualities analysis using an adequate representation of the aircraft flight dynamics as a flexible body and mathematical pilot models.

© 31st Congress of the International Council of the Aeronautical Sciences, ICAS 2018. All rights reserved.The advantages of full-state feedback adaptive control in dealing with uncertainties of a flexible aircraft model are demonstrated with the use of model reference adaptive control. Design of an output-feedback system with an observer, based on the separation principle, is attempted. Differences in both stability and performance characteristics of the full-state feedback and the output-feedback closed-loop systems demonstrate that further investigation is needed to design adaptive output-feedback controllers.

© 31st Congress of the International Council of the Aeronautical Sciences, ICAS 2018. All rights reserved.Flight simulation of flexible aircraft is computationally expensive. This paper presents an approach based on reduced-order models for computational cost reduction of the aeroelastic equations. The model order reduction technique, the X-HALE aircraft and a low computational cost model are described. Finally, simulation time histories are presented comparing time responses and computational costs. Numerical results show good agreement between full-order and reduced-order models, with the latter presenting a significant reduction in computational time.

© 31st Congress of the International Council of the Aeronautical Sciences, ICAS 2018. All rights reserved.This paper presents a new concept on pilot model that allows the study of handling qualities and PIO (Pilot-Induced Oscillations) of elastic aircraft since it computes not only the voluntary pilot inputs in order to accomplish a given task, but also the involuntary commands due to the biodynamic feedthrough (BDFT) or Pilot-Assisted Oscillations (PAO) phenomena. Frequency-domain criteria used for PIO prediction are analyzed and pilot simulations are also presented.

© 31st Congress of the International Council of the Aeronautical Sciences, ICAS 2018. All rights reserved.To stabilize a very flexible wing with a control system it is necessary to feedback wing measurements such as accelerations, deformations or displacements. In a previous work, displacement sensors were positioned at the wing tips and then the gain matrix of the feedback control was optimized. In this way, the performance of the control is restricted to the predetermined position of the sensors. Hence, in this paper we apply another methodology to simultaneously optimize the sensors positioning and the gain matrix. The results are compared in order to evaluate the benefits of this methodology in the considered problem.

© 31st Congress of the International Council of the Aeronautical Sciences, ICAS 2018. All rights reserved.This article presents the development of a real-time flight load measurement system convertible to real-time flight geometry measurement. It uses already known full bridge strains gages in a different manufacturing possibility: inside composite layers before resin cure. The manufacturing process is validated using a composite bar prototype. The prototype is submitted to a four points structural test in order to validate the measurements and influences of the manufacturing process. The whole process is applied to X-HALE aircraft disclosing a composite built-in measurement system for real-time flight load and flight geometry.

© 31st Congress of the International Council of the Aeronautical Sciences, ICAS 2018. All rights reserved.The assumption of small deformations in the formulation of the flight dynamics of flexible aircraft can be very convenient as it allows the use of a reduced number of modes of vibration to represent the structural dynamics with little loss of accuracy. However, depending on the level of structural flexibility, deformations may become large enough to violate this simplifying assumption, making geometrically-nonlinear formulations necessary. Delimiting the range of validity of small deformations is then indispensable for the flight-mechanics engineer in state-of-the-art aircraft design. In this paper, small- and large-deformation formulations are compared in equilibrium conditions and in time-marching simulations for aircraft with different levels of structural flexibility. The importance of geometrical nonlinearities and the range of validity of small deformations are assessed.

© 2017 International Forum on Aeroelasticity and Structural Dynamics (IFASD). All Rights Reserved.Futuristic aircraft designs and novel aircraft such as High Altitude Long Endurance (HALE) involve a higher level of structural flexibility than in conventional aircraft. Even at present, the trends in the aviation industry are to increase wing length (to reduce induced drag) and maximize use of composites, which lead to increased structural flexibility. This necessitates a rethink of conventional (linear) aeroelastic analysis, since the increased flexibility results in coupling between the flight dynamic and aeroelastic dynamics, and consequently, limit-cycle oscillations of the structure. In this paper, a new three-dimensional low-order model for unsteady aerodynamics that accounts for large oscillation amplitudes and nonplanar wakes is developed. An experiment with a cantilevered flat plate at low Reynolds number is set up and used to validate the low-order model, as well as to study post-flutter limit-cycle oscillations. Results from the low-order model are promising, but show that aerodynamic nonlinearities such as flow separation and leading-edge vortex shedding must also be modeled in order to predict all possible limit-cycle oscillations of the aeroelastic system.

© 2017 International Forum on Aeroelasticity and Structural Dynamics (IFASD). All rights reserved.A simple and self-contained methodology to assess the validity of the assumption of small deformations in linear structural-dynamic models was recently proposed. The advantages of the methodology lie in the fact that it does not depend on the availability of higher-fidelity, nonlinear models: it is rather based on the selection of two different structural nodes where the structural motion is to be one at a time completely constrained, typically, a node near the center of mass and another in the region of maximum structural displacements with respect to mean axes. If the two displacement vectors calculated in each case can be transformed between themselves with linear rigid-body modes of the structure, then it is still in the regime of small deformations. In the present paper, in order to demonstrate the value of this methodology, it is applied to the X-HALE aircraft in its four-, six- and eight-meter-span configurations, and the results obtained with the assumption of small deformations are compared with a higher-fidelity model that comprises large structural deformations.

© 2017 International Forum on Aeroelasticity and Structural Dynamics (IFASD). All rights reserved.Highly-flexible aircrafts present high-aspect-ratio wings that introduce nonlinearities into the flight dynamics and make more complex models and control methods necessary. In this paper, a loop separation concept is applied to the X-HALE aircraft, giving rise to a control system comprising an inner-loop capable of controlling the shape of the aircraft while keeping the plant stable, and an outer-loop capable of controlling velocity, altitude, bank angle and sideslip angle. The outer-loop has a decoupled structure for longitudinal and for the lateral-directional axes. The matrices of the compensators were obtained using non-smooth optimization. Gain-scheduling techniques are implemented to bypass stability problems arising from changes in the plant with flight speed. Nonlinear simulations show promising results for implementation on the real aircraft.

© 2017 International Forum on Aeroelasticity and Structural Dynamics (IFASD). All Rights Reserved.Real–time simulation is a valuable tool for aircraft development. However, flexible aircraft models are computationally expensive and can be prohibitive for fast simulations. This paper investigates the applicability of a nonlinear model order reduction technique to a moderate flexible aircraft for real–time simulation purposes. Details of the implementation and test cases are presented. Results show that the reduced–order model can be simulated in real–time and produces better results than linearized model and rigid–body model with aeroelastic correction and we conclude that the technique is promising for real–time simulations.

© 2017 International Forum on Aeroelasticity and Structural Dynamics (IFASD). All Rights Reserved.Formulations for the flight dynamics of flexible aircraft have been commonly applied to aircraft free to fly in the three-dimensional space, having all six rigid-body degrees of freedom. For risk reduction in the future flight operations of the X-HALE testbed at ITA, however, wind-tunnel tests of the remotely-piloted, four-meter-span configuration of the aircraft were performed. In the wind tunnel, the rigid-body translations were completely constrained, but the same was not valid for the rigid-body rotations, which could be conveniently left free or not with a proper selection of the connection between the aircraft and the wind-tunnel mount. In the present paper, in order to computationally assess the response and stability characteristics of the aircraft in the wind tunnel, we derive equations of motion for a constrained flexible aircraft with up to three rigid-body rotational degrees of freedom, mounted on an also flexible wind-tunnel strut. The numerical model has its value confirmed by the wind-tunnel tests in the predicted and observed roll-control reversal for anti-symmetrical deflections of the all-moving tails, and absence of reversal for aileron deflections.

© 2017 International Forum on Aeroelasticity and Structural Dynamics (IFASD). All rights reserved.In this paper, the application of adaptive control is demonstrated for highly flexible aircraft in a dynamic gain scheduling approach. For that, an adaptive control law is designed for the stability augmentation system, taking as reference model the closed-loop dynamics provided by a baseline controller. Global stability is demonstrated by the Lyapunov direct method. Closed-loop nonlinear simulations are performed. The performance of the control system is evaluated considering the aircraft in high loading operations, specifically in accelerated turning maneuvers. Overall, this control strategy proved to overcome instabilities arising when linear control is applied, while conferring the aircraft adequate time response.

Copyright © 2016 by Flavio Silvestre. Published by the American Institute of Aeronautics and Astronautics, Inc., with permission.In this paper, the control law design for flexible aircraft is discussed. First, the traditional procedure of decoupling rigid-body and aeroelastic dynamics with low-pass and notch filters is addressed, with focus on controller performance as well as the resulting stability margin issues. A procedure based on a unified formulation of the flexible aircraft dynamics for flight control law design is proposed. In this procedure, the aeroservoelastic dynamics is assessed in the loop, and the offline filtering process is avoided. The formulation is applied to the virtual aircraft generic narrow-body airliner, with improvements in closed-loop performance and stability margins.

Copyright © 2016 by Daniel Drewiacki, Flavio Jose Silvestre, Antonio Bernardo Guimaraes Neto. Published by the American Institute of Aeronautics and Astronautics, Inc.Frequency domain handling qualities criteria were developed in order to predict and prevent Pilot Induced Oscillation (PIO) occurrences. However, these criteria have been developed neglecting the influence of airframe flexibility. Since aircraft are becoming more flexible, this influence has to be considered during the development of y-by-wire systems. In this paper, we address the effects of aeroelastic dynamics in aircraft handling qualities.

© Copyright 2016 by the American Institute of Aeronautics and Astronautics, Inc.An inertially-coupled formulation for the flight dynamics of flexible aircraft undergoing small deformations is developed. The availability of a structural-dynamic finite element model of the aircraft is presupposed. With all the coupled dynamics taken into account, an arbitrary choice of the body reference frame can be made. This frame is also allowed to be noncoincident with the frame of reference used to calculate the aerodynamic loads. In the equations of motion, the inertial coupling terms are linearized in the elastic displacements around a calculated equilibrium condition. Appropriate modes of vibration are then used in the calculation of the dynamic deformation of the structure. A simple quasi-steady incremental aerodynamic model based on the vortex-lattice method is used. The formulation is tested in the flight simulation of an idealized forward-swept-wing aircraft model. Numerical results show that, under small deformations, different body axes lead to the same overall motion of the aircraft with respect to an inertial frame. The stiffness level at which geometrically nonlinear formulations would become necessary is also determined.

We study in this paper the use of a unified formulation of the flexible aircraft dynamics for flight control law design, applied to the virtual aircraft Generic Narrow-Body Airliner4,5 (GNBA). In this procedure, the aeroservoelastic dynamics can be assessed in the loop, and the offline filtering process can be avoided. The controller performance as well as the stability margin issues associated to notch and low-pass filtering are addressed.

The use of correction factors to improve the accuracy of the aerodynamic influence coefficient (AIC) matrices produced by the vortex-lattice and the doublet-lattice methods has been an engineering practice in the field of aeroelasticity. In order to account for either viscous or transonic flow effects not considered in the linearized formulation of such methods, the most frequent correction techniques have been to pre-multiply or to post-multiply the AIC matrices by diagonal matrices comprising semi-empirical weighting factors. This paper proposes a different correction approach: the control-point-placement method (CPPM), based on the idea of displacing the control point of each panel - the point where the boundary condition of flow tangency must be satisfied. Both the vortex- and the doublet-lattice methods have been developed with the singularities placed at the quarter-chord line of the panels and the control points at three quarters of their mean chords. With the calculation of modified control point positions, the CPPM intrinsically changes the mutual aerodynamic influence between the panels and allows the lifting surface methods to predict steady-state pressure distributions that match or approximate with minimum error those derived from wind tunnel measurements or higher-fidelity CFD solutions. Different approaches to extend the aerodynamic correction for application at non-zero reduced frequencies in the doublet-lattice method are then studied. Results are presented that are in acceptable agreement with benchmark wind tunnel data and comparisons are made between the proposed methodology and the traditional diagonal matrix corrections. © 2014 Published by Elsevier Masson SAS.

This chapter presents the investigations about the characteristics of the cathode sheath generated in an argon low-pressure hollow cathode discharge (HCD). The theoretical model for the cathode sheath is considered to be non-collisional. Secondary electrons emitted from the cathode surface are taken into account in the discharge model and their influence on the theoretical sheath potential profile is investigated. The plasma parameters and the floating potential profile along the discharge axis were inferred from the current-voltage characteristics of a single Langmuir probe positioned at the inter-cathode space of the HCD. For a low pressure HCD, typical values of the electron density and electron temperature are ne ≈ 1016 m-3 and Te = 4 eV, respectively. By using the probe data, the floating potential profile was determined to verify the position of the plasma-cathode sheath interface and to promote a qualitative discussion between theoretical and experimental results. © 2013 by Nova Science Publishers, Inc. All rights reserved.

Aeronautical engineering has faced a significant and continued development over the last decades towards the design of lighter, more maneuverable and more multidisciplinarily optimized aircraft, leading to more flexible vehicles. In this context, the fields of aeroelasticity and aeroservoelasticity play a very important and increasing role. Neglecting such flexibility effects on the flight dynamics and control system analysis and design may be an invalid premise, depending on how intense might be the coupling between the rigid and the flexible degrees of freedom. Traditional modeling approaches have often neglected the effects of inertial coupling in the treatment of the dynamics of the deformable aircraft, allowing great simplifications of the equations of motion. Most authors have indeed considered the body axes to be mean axes, what requires some care regarding the enforcement of the correct constraints and the expression of the aerodynamic force components along these axes directions. Looking for circumventing those limitations, while keeping the hypothesis of small local deformations, this work presents an integrated modeling methodology for the flight dynamics of deformable aircraft which takes into account all the coupled dynamics and is based on attached body axes. The formulation is developed for direct use with a finite-element model of the aircraft structure, with known distributed or lumped mass properties. The nonlinear inertial coupling terms are linearized with respect to the linear elastic displacements around an equilibrium condition. This condition is determined with the full nonlinear dynamics, considering displacement and load-transferal between the aerodynamic model and the finite-element model. Inertia-relieved constrained modes of vibration are then used as shape functions in the calculation of the dynamic deformation of the structure, thus not canceling the inertial coupling terms as would happen in the case of free-free normal modes. The proposed formulation is implemented and tested for simulating the flight of a generic narrow-body airliner (GNBA) model which has been developed for the purpose of these studies. The aerodynamic forces and moments are treated as the superposition of two contributions: the expected rigid-body ones and the incremental ones due to the structural deformation. The incremental aerodynamic forces and moments are modeled by the doubletlattice method (DLM). Rational-function approximation (RFA) together with the method of least squares for complex variables to determine the coefficients of the RFA and inverse Laplace transforms are employed to represent the reduced-frequency-domain forces in the time domain, leading to an augmented state-space system in which the aerodynamic lag phenomenon is taken into account.