This paper proposes a nonlinear control architecture for flexible aircraft simultaneous trajectory tracking and load alleviation. By exploiting the control redundancy, the gust and maneuver loads are alleviated without degrading the rigid-body command tracking performance. The proposed control architecture contains four cascaded control loops: position control, flight path control, attitude control, and optimal multi-objective wing control. Since the position kinematics are not influenced by model uncertainties, the nonlinear dynamic inversion control is applied. On the contrary, the flight path dynamics are perturbed by both model uncertainties and atmospheric disturbances; thus the incremental sliding mode control is adopted. Lyapunov-based analyses show that this method can simultaneously reduce the model dependency and the minimum possible gains of conventional sliding mode control methods. Moreover, the attitude dynamics are in the strict-feedback form; thus the incremental backstepping sliding mode control is applied. Furthermore, a novel load reference generator is designed to distinguish the necessary loads for performing maneuvers from the excessive loads. The load references are realized by the inner-loop optimal wing controller, while the excessive loads are naturalized by flaps without influencing the outer-loop tracking performance. The merits of the proposed control architecture are verified by trajectory tracking tasks and gust load alleviation tasks in spatial von K'arm'an turbulence fields. @en

This paper deals with the simultaneous gust and maneuver load alleviation problem of a seamless active morphing wing. The incremental nonlinear dynamic inversion with quadratic programming control allocation and virtual shape functions (denoted as INDI-QP-V) is proposed to fulfill this goal. The designed control allocator provides an optimal solution while satisfying actuator position constraints, rate constraints, and relative position constraints. Virtual shape functions ensure the smoothness of the morphing wing at every moment. In the presence of model uncertainties, external disturbances, and control allocation errors, the closed-loop stability is guaranteed in the Lyapunov sense. Wind tunnel tests demonstrate that INDI-QP-V can make the seamless wing morph actively to resist “1-cos” gusts and modify the spanwise lift distribution at the same time. The wing root shear force and bending moment have been alleviated by more than 44% despite unexpected actuator fault and nonlinear backlash. Moreover, during the experiment, all the input constraints were satisfied, the wing shape was smooth all the time, and the control law was executed in real time. Furthermore, as compared with the linear quadratic Gaussian control, the hardware implementation of INDI-QP-V is easier; the robust performance of INDI-QP-V is also superior.@en

Quad-plane is a popular type of electric vertical and takeoff/landing (eVTOL) vehicle that hybridizes a quadrotor and a fixed-wing airplane. However, the mechanical simplicity of a quad-plane also makes it vulnerable to rotor failures. When a complete rotor fails, it becomes physically impossible to stop the quad-plane from fast yaw spinning, which further induces considerable abnormal aerodynamic forces and moments on the wing. In this paper, a novel incremental adaptive sliding mode control (I-ASMC) is proposed to address these challenges. First, by exploiting sensor measurements, it simultaneously reduces the control model dependency and the minimum possible sliding mode control/observer gains. Second, finite-time convergence is guaranteed in the Lyapunov sense. Third, the control gains are automatically adapted to their minimum possible values without prior-knowledge on the uncertainty bounds. The proposed I-ASMC method is verified on a high fidelity simulation platform with computational fluid dynamic (CFD) aerodynamic models. Simulation results demonstrate that I-ASMC can drive a quad-plane with a complete loss of a single rotor to follow a trajectory. Its robustness to aerodynamic model uncertainties and rotor faults is also better than the linear quadratic regulator (LQR) and the incremental nonlinear dynamic inversion (INDI) control. In conclusion, the reduced model dependency, implementation simplicity, and improved robustness make the proposed I-ASMC promising for enhancing quad-plane safety in real life.

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This paper presents a study on trailing edge deflection estimation for the SmartX camber morphing wing demonstrator. This demonstrator integrates the technologies of smart sensing, smart actuation and smart controls using a six module distributed morphing concept. The morphing sequence is brought about by two actuators present at both ends of each of the morphing modules. The deflection estimation is carried out by interrogating optical fibers that are bonded on to the wing’s inner surface. A novel application is demonstrated using this method that utilizes the least amount of sensors for load monitoring purposes. The fiber optic sensor data is used to measure the deflections of the modules in the wind tunnel using a multi-modal fiber optic sensing approach and is compared to the deflections estimated by the actuators. Each module is probed by single-mode optical fibers that contain just four grating sensors and consider both bending and torsional deformations. The fiber optic method in this work combines the principles of hybrid interferometry and FBG spectral sensing. The analysis involves an initial calibration procedure outside the wind tunnel followed by experimental testing in the wind tunnel. This method is shown to experimentally achieve an accuracy of 2.8 mm deflection with an error of 9%. The error sources, including actuator dynamics, random errors, and nonlinear mechanical backlash, are identified and discussed.

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In order to further expand the flight envelope of quadrotors under actuator failures, we design a nonlinear sensor-based fault-tolerant controller to stabilize a quadrotor with failure of two opposing rotors in the high-speed flight condition (>8 m/s). The incremental nonlinear dynamic inversion approach which excels in handling model uncertainties is adopted to compensate for the significant unknown aerodynamic effects. The internal dynamics of such an underactuated system have been analyzed, and subsequently stabilized by redefining the control output. The proposed method can be generalized to control a quadrotor under single-rotor-failure and nominal conditions. For validation, flight tests have been carried out in a large-scale open jet wind tunnel. The position of a damaged quadrotor can be controlled in the presence of significant wind disturbances. A linear quadratic regulator approach from the literature has been compared to demonstrate the advantages of the proposed nonlinear method in the windy and high-speed flight condition.@en

The advancements made in aircraft control methodology and the tendency towards increasingly lighter aircraft structures open the opportunity to higher structural performance and aerodynamic efficiency. However, with the reduction of structure weight, the structure stiffness reduces typically, which makes the structure more susceptible to external dynamic loads. How the flexibility affects the dynamics of the system, in particular in closed-loop control, cannot always be determined in the early stage of the design process. This introduces uncertainties to the dynamic model and consequently leads to inaccurate performance evaluation. Our proposed approach to fully utilize the potential of the lighter aircraft structure is to actively morph it using distributed actuators commanded by a real-time multi-objective controller. In the literature, model-based feedback control methods are widely used for flexible structure motion suppression. However, the performance of model-based controllers is impaired by model uncertainties and external disturbances. Another challenge in flexible structure control is the real-time state estimation. Although accelerometers and strain gauges can be used to capture the structural vibrations, these sensors have to be installed within the structure, which increases the difficulties in maintenance. A potentially universal, model-free, and non-invasive approach is visual tracking. In combination with robust model-free control laws, this has the potential to create smart adaptive structures that are capable of vibration suppression. In this study, an adaptive model-free state estimation methodology based on visual feedback is developed and demonstrated in unison with a non-linear model-free robust control method in a closed-loop system. The experimental setup consists of high-speed GIGE cameras observing oscillations from a clamped beam subject to disturbances at 140 Hz. The task of the controller is to reject the disturbances through a shaker input under the presence of parametric model uncertainties and external disturbances. The visual tracking utilizes adaptive estimation utilizing high-speed KCF (Kernelized Correlation Filter) tracking and an AEKF (Augmented Extended Kalman Filter) with augmented time-varying mass, stiffness, and damping states. The inclusion of the augmented Kalman filter adds robustness to occlusion and model uncertainties. The nonlinear model-free control method is developed in the incremental control framework, hybridized with sliding mode control for robustness enhancement. The control effectiveness matrix used by the controller is adapted online. Furthermore, the state and state derivative feedback signals are provided by the visual system in real-time. This research is a part of the Smart-X wing project, which represents an autonomous smart morphing wing that is capable of in-flight performance optimisation of multiple objectives. It is shown in this research that the combination of adaptive visual tracking and robust control shows how a flexible uncertain structure can be transformed into a controlled adaptive smart structure. This combination of visual tracking and control shows great potential for robust and model-free stabilization and vibration suppression. Further uses of the methodology are discussed for use in tracking problems for flexible and aeroelastic structures. @en

For mitigating the chattering effect in the sliding mode control (SMC), many adaption mechanisms have been proposed to reduce the switching gains. However, less attention is paid to the control structure, which influences the resulting uncertainty term and determines the minimum possible gains. This paper compares three control structures for inducing higher-order sliding modes in finite time: nonlinear dynamic inversion (NDI) based SMC, higher-order sliding mode control (HOSMC) with artificially increased relative degree, and the recently proposed incremental nonlinear dynamic inversion (INDI) based SMC. The latter two control structures have reduced model dependency as compared to NDI-SMC. Moreover, their nominal control increments are found to be approximately equivalent if the sampling interval is sufficiently small and if their gains satisfy certain conditions. Under the same circumstances, the norm value of the resulting uncertainty using INDI-SMC is several orders of magnitude smaller than those using other control structures. For maintaining the sliding modes, the minimum possible gains required by HOSMC approximately equal those needed by INDI-SMC divided by the sampling interval. Nevertheless, these two approaches have comparable chattering degrees, which are effectively reduced as compared to NDI-SMC. The analytical results are verified by numerical simulations.

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This paper proposes a novel control framework that combines the recently reformulated incremental nonlinear dynamic inversion with (higher-order) sliding-mode controllers/observers, for generic multi-input/multi-output nonlinear systems, named incremental sliding-mode control. As compared to the widely used approach that designs (higher-order) sliding-mode controllers/observers based on nonlinear dynamic inversion, the proposed incremental framework can further reduce the uncertainties while requiring less model knowledge. Because the uncertainties are reduced in the incremental framework, theoretical analyses demonstrate that the incremental sliding-mode control can passively resist a wider range of perturbations with reduced minimum possible control/observer gains. These merits are validated via numerical simulations for aircraft command tracking problems, in the presence of sudden actuator faults and structural damage.

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As a sensor-based control method, incremental nonlinear dynamic inversion (INDI) has been applied to various aerospace systems and has shown desirable robust performance against aerodynamic model uncertainties. However, its previous derivation based on the time scale separation principle has some limitations. There is also a need for stability and robustness analysis for INDI. Therefore, this paper reformulates the INDI control law without using the time scale separation principle and generalizes it for systems with arbitrary relative degree, with consideration of the internal dynamics. The stability of the closed-loop system in the presence of external disturbances is analyzed using Lyapunov methods and nonlinear system perturbation theory. Moreover, the robustness of the closed-loop system against regular and singular perturbations is analyzed. Finally, this reformulated INDI control law is verified by a Monte Carlo simulation for an aircraft command tracking problem in the presence of external disturbances and model uncertainties.@en

This paper proposes incremental nonsingular terminal sliding mode control for a class of multi-input and multi-output nonlinear systems considering model uncertainties, external disturbances, and sudden actuator faults. This method is free from singularity because it does not involve any negative fractional power. The convergence time in both reaching and sliding phases are proved to be finite. Moreover, by fully exploiting sensor measurements, the proposed incremental control method simultaneously reduces model dependency and the uncertainty remaining in the closed-loop system. The reduction of model dependency simplifies the implementation process and reduces the computational load, while the reduction of uncertainty decreases the minimum possible sliding mode control gains, which is beneficial to chattering reduction. These merits are verified by a quadrotor trajectory tracking problem. Simulation results demonstrate that the proposed method has better robustness against model uncertainties, gusts, and actuator faults than the model-based nonsingular terminal sliding mode control in the literature.

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Fault-tolerant flight control has the potential of improving the aircraft survivability in real life. This paper proposes an Incremental Backstepping Sliding Mode Control (IBSMC) framework for multi-input/output nonlinear strict-feedback systems considering model uncertainties, sudden faults, and external disturbances. This approach is a hybridization of the Sliding Mode Control (SMC) and a reformulated Incremental Backstepping (IBS). By virtue of the benefits contributed by both SMC and IBS, theoretical analyses prove IBSMC has less model dependency and enhanced robustness as compared to backstepping and backstepping hybridized with SMC (BSMC). When applied to an aircraft fault-tolerant control problem, numerical simulations demonstrate IBSMC can passively tolerate a wider range of model uncertainties, sudden actuator faults, and sudden structural damages as compared to backstepping and BSMC, using smooth control inputs with lower gains. @en

This paper proposes an Incremental Sliding Mode Control driven by Sliding Mode Disturbance Observers (INDI-SMC/SMDO), with application to a quadrotor fault tolerant control problem. By designing the SMC/SMDO based on the control structure of the sensor-based Incremental Nonlinear Dynamic Inversion (INDI), instead of the model-based Nonlinear Dynamic Inversion (NDI) in the literature, the model dependency of the controller and the uncertainties in the closed-loop system are simultaneously reduced. This allows INDI-SMC/SMDO to passively resist a wider variety of faults and external disturbances using continuous control inputs with lower control and observer gains. When applied to a quadrotor, both numerical simulations and real-world flight tests demonstrate that INDI based SMC/SMDO has better performance and robustness over NDI based SMC/SMDO, in the presence of model uncertainties, wind disturbances, and sudden actuator faults. Moreover, the implementation process is simplified because of the reduced model dependency and smaller uncertainty variations of INDI-SMC/SMDO. Therefore, the proposed control method can be easily implemented to improve the performance and survivability of quadrotors in real life.

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The swift growth of air traffic volume stresses the importance of flight safety enhancement. Statistical data shows that fly-by-wire technology with automatic flight control systems can effectively reduce the fatal accident rate of loss of control in-flight. Although the dynamics of an aircraft are nonlinear and time-varying, it is common practice to design flight control laws based on local linear time-invariant (LTI) dynamic models, and apply gain-scheduling method. Here, the flight envelope is divided into many smaller operating regimes, and LTI model-based controllers are designed and tuned for each of them. However, this approach is cumbersome and cannot guarantee flight stability and performance in-between operational points. In view of the challenges encountered by LTI model-based control, nonlinear control methods have attracted attention from the flight control community. Nonlinear dynamic inversion (NDI) and backstepping (BS) are two frequently used nonlinear control methods in flight control. These two approaches cancel the nonlinearities in the closed loop using a nonlinear model of the system. However, mismatches between the model and real dynamics inevitability exist, especially when an aircraft encounters atmospheric disturbances and when sudden actuator faults or even structural damages occur. To enhance the robustness of model-based nonlinear control methods to model mismatches, a commonly adopted approach is to augment them with online model identification. This process, however, is computational intensive and requires sufficient excitation, which can make an impaired aircraft fly out of the diminished safe flight envelope. In consideration of these challenges, the main goal of this thesis is: To design a stability-guaranteed nonlinear flight control framework with reduced model dependency and enhanced robustness. @en

This paper designs an incremental nonlinear dynamic inversion control law for free-flying flexible aircraft, which can regulate rigid-body motions, alleviate gust loads, reduce the wing root bending moment, and suppress elastic modes. By fully exploring the sensor measurements, the model dependency of the proposed control law can be reduced while maintaining desirable robustness, which simplifies the implementation process and reduces the onboard computational load. The elastic states are observed online from accelerometer measurements, with a Padé approximation to model the pure time delay. Theoretical analyses based on the Lyapunov methods and the nonlinear system perturbation theory show that the proposed control has inherent robustness to model uncertainties, external disturbances, and sudden actuator faults. These merits are demonstrated by time-domain simulations in various spatial turbulence and gust fields, as well as by a Monte Carlo study.@en

In this paper, an Incremental Nonlinear Dynamic Inversion (INDI) controller is developed for the flexible aircraft gust load alleviation (GLA) problem. First, a flexible aircraft model captures both inertia and aerodynamic coupling effects between flight dynamics and structural vibration dynamics is presented. Then an INDI GLA controller is designed for this aircraft model based on sensor measurements and the Kalman filter online estimation. Besides, the fifth order Padé approximation is used to model the pure time delay in the state estimation. Furthermore, simulations of the flexible aircraft flying through various spatial turbulence and gust fields demonstrate the effectiveness of the proposed controller on rigid-body motion regulation, vertical load alleviation, wing root bending moment reduction and elastic modes suppression. Additionally, numerical perturbation tests and a Monte-Carlo study show the robustness of the proposed controller to aerodynamic model uncertainties.@en

This paper presents a unified framework for aeroelastic tailoring of free-flying aircraft with composite wings. A continuous-time state-space model is used to describe the flow. The 3D composite wing structures is condensed into a Timoshenko beam model by means of a cross-sectional modeler. The aerodynamic and structural models are closely coupled with the six degrees of freedom flight dynamic equations of motion in the state-space formulation. This paper refers to the clamped-wing aeroelastic tailoring as classic aeroelastic tailoring. Hence, the term aeroelastic tailoring will point at the novel approach that includes free-flying aeroelastic phenomena into the optimization process. The emphasis of the present paper is to show the effects of aeroelastic tailoring on body-freedom flutter and flight dynamic stability at large. The results of this paper will be used in the further development of aeroelastic tailoring practices for composite aircraft design.@en

As a sensor-based control approach, the Incremental Nonlinear Dynamic Inversion (INDI) method has been successfully applied on various aerospace systems and shown desirable robust performance to aerodynamic model uncertainties. However, its previous derivations based on the so-called time scale separation principle is not mathematically rigorous. There also lack of stability and robustness analysis for INDI. Therefore, this paper reformulated the INDI control law without using the time scale separation principle and generalized it to not necessarily relative-degree-one problems, with consideration of the internal dynamics. Besides, the stability of the closed-loop system in the presence of external disturbances is analyzed using Lyapunov methods and nonlinear system perturbation theory. Moreover, the robustness of the closed-loop system against regular and singular perturbations is analyzed. Finally, the reformulated INDI control law and main conclusions are verified by a rigid aircraft gust load alleviation problem.@en

In order to achieve high-speed flight of a damaged quadrotor with complete loss of a single rotor, a multiloop hybrid nonlinear controller is designed. By fully making use of sensor measurements, the model dependence of this control method is reduced, which is conducive to handling disturbance from the unknown aerodynamic effects. This controller is tested on a quadrotor vehicle with one rotor completely removed in the high-speed condition. Free flights are performed in the Open Jet Facility, a large-scale wind tunnel. Over 9 m/s flight speed is reached for the damaged quadrotor in these tests. In addition, several high-speed spin-induced aerodynamic effects are discovered and analyzed.@en

This paper presents a novel active gust load alleviation approach within a multi-objective flight control framework developed by NASA for a flexible wing aircraft. The aircraft model is based on the NASA Generic Transport Model (GTM). The wing structures incorporate an aerodynamic control surface known as the Variable Camber Continuous Trailing Edge Flap (VCCTEF). Previous work already showed the ability of the VCCTEF to perform aeroelastic mode suppression, drag minimization and maneuver load alleviation in a multi-objective flight control framework. In this paper, the multi-objective flight control framework is extended to include active gust load alleviation. A Linear-Quadratic Gaussian (LQG) controller is augmented with Model Reference Adaptive Control (MRAC) to provide active gust load alleviation. Disturbance estimation is done using an Extended State Observer (ESO) to support the design of the active gust load alleviation controller. The results demonstrate the potential of active gust load alleviation within a multi-objective flight control framework for a high-aspect ratio flexible wing aircraft embodied with the VCCTEF.@en

Controllability Margin (CM) and Observability Margin (OM) for linear systems, including Linear Time Invariant (LTI) and Linear Time Varying (LTV) systems are defined from the view of singular perturbation and regular perturbation, and the corresponding assessment methods are established. Based on the CM, OM and the recently introduced Singular Perturbation Margin (SPM) and Generalized Gain Margin (GGM), the concept of Model Error Structure (MES) is provided as a structural property metric for linear systems. The results in this paper are illustrated by the examples of a flexible beam and Hill equations.

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