Search

This paper presents a practical model predictive control (MPC) framework for gust load alleviation of a flexible flying wing. Both the controller solving and state estimation are based on a reduced-order model, which features a linear parameter-varying (LPV) form, avoiding online linearization and reducing the scale of the corresponding quadratic programming problem. An improved modeling and model reduction process is used to enhance modeling efficiency and ensure that the reduced-order model can accurately capture the rigid-flexible coupled characteristics of the flexible flying wing under arbitrary gusts. By reconstructing the output of the control-oriented model to include both rigid-body motion and flexible vibrations, the rigid-flexible coupled multi-objective control is established as an MPC problem for reference tracking. The online optimization is formulated in a sparse fashion and combined with an iterative algorithm based on predicted trajectories, describing the variation of model dynamics within the prediction horizon more accurately. With a time-varying Kalman estimator for state updating, the closed-loop simulations are performed for gust alleviation performance validation. Additionally, the real-time potential of the proposed MPC framework is demonstrated through Monte Carlo simulations. • Comprehensive modeling, model order reduction, and control design for flexible aircraft. • An extended unsteady vortex lattice model to improve aerodynamic modeling efficiency. • A reduced-order model-based control and state estimation framework suitable for engineering practice. [ABSTRACT FROM AUTHOR]

This paper makes use of sloshing reduced-order models to investigate the effects of sloshing dynamics on aeroelastic stability and response of flying wing structure. More specifically, a linear frequency-domain operator derived by an equivalent mechanical model is used to model lateral (linear) sloshing dynamics whereas data-driven neural-networks are used to model the vertical (nonlinear) sloshing dynamics. These models are integrated into a formulation that accounts for both the rigid and flexible behavior of aircraft. A time domain representation of the unsteady aerodynamics is achieved by rational function approximation of the fully unsteady aerodynamics obtained via the doublet lattice method. The case study consists of the so called Body Freedom Flutter research model in two different configurations with one or two tanks partially filled with liquid with a mass comprising 25% of the aircraft structure. The results show that linear sloshing dynamics are able to change the stability margin of the aircraft in addition to having non-negligible effects on rigid body dynamics. On the other hand, vertical sloshing acts as a nonlinear damper and eventually provides limit cycle oscillations after flutter onset. [ABSTRACT FROM AUTHOR]

The evolution of large transport aircraft is characterized by longer fuselages and larger wingspans, while efforts to decrease the structural weight reduce the structural stiffness. Both effects lead to more flexible aircraft structures with significant aeroelastic coupling between flight mechanics and structural dynamics, especially at high speed, high altitude cruise. The lesser frequency separation between rigid body and flexible modes of flexible aircraft results in a stronger interaction between the flight control system and its structural modes, with higher flexibility effects on aircraft dynamics. Therefore, the design of a flight control law based on the assumption that the aircraft dynamics are rigid is no longer valid for the flexible aircraft. This paper focuses on the design of a flight control system for flexible aircraft described in terms of a rigid body mode and four flexible body modes and whose parameters are assumed to be varying. In this paper, a conditional integral based sliding mode control (SMC) is used for robust tracking control of the pitch angle of the flexible aircraft. The performance of the proposed nonlinear flight control system has been shown through the numerical simulations of the flexible aircraft. Good transient and steady-state performance of a control system are also ensured without suffering from the drawback of control chattering in SMC. [ABSTRACT FROM AUTHOR]

This article presents the PIO identification for a conceptual aircraft subjected to aeroelastic effects. The flight data was obtained in an extensive flight simulator campaign, using the DLR AVES full-motion simulator, performed by four test pilots, as part of Embraer's/DLR's DinAFlex project. The aircraft model inserted in the simulator considered three cases, varying in structural flexibility and model fidelity (quasi-static corrections or the presence of aeroelastic modes). The first stage of the analysis presented the PIO identification from the pilots' perspective, considering their PIOR assessment. The subjective nature inherent to the PIOR scale was explored by comparing the flight configurations' assessment, using a proposed metric called pilot-to-pilot discrepancy, along with possible causes of its existence. The second stage presented the ROVER analysis, a quantitative approach that identified PIO events in the data and correlated them with the PIOR results, using another proposed metric called ROVER-to-pilot discrepancy. The discrepancy metrics provided a deep understanding of how the pilots deal with aeroelastic effects and evidenced that the PIOR assessment was impaired as these effects increased, presenting an issue to flight testing and certification of new, more flexible, aircraft, where pilot's assessment is of great importance. • A PIO identification analysis for a flexible aircraft using a moving base flight simulator is presented. • Discrepancies between the pilots' assessment using the PIOR increased, as the aircraft flexibility was increased. • Discrepancies between the ROVER algorithm and pilots' assessment also increased, as the aircraft flexibility was increased. • These metrics provide an understanding of how the pilots are dealing with PIO assessment with aeroelastic effects. [ABSTRACT FROM AUTHOR]

Flexible high-altitude long endurance (HALE) aircraft face challenges such as the delay effect caused by aeroelasticity and low efficiency during high-altitude cruising. This paper addresses a new active control scheme that replaces traditional aerodynamic control surfaces with longitudinal and lateral moving masses and adopts an observer-based adaptive dynamic programming (ADP) control strategy, aiming to solve the attitude control of HALE aircraft with flexible wings. Firstly, considering the account of unsteady aerodynamics, attitude angular velocity, and moving masses, complete dynamic model is established for a flexible aircraft, in order to more accurately describe the state of the HALE aircraft. This model includes two moving masses to replace traditional ailerons and elevators. It is difficult to design attitude control strategies considering the delay effects and unknown dynamic disturbances of HALE aircraft. To this end, the actor-critic structure of the ADP approach is designed to achieve the near-optimal control strategy of the system, which eliminates the need for prior knowledge of model parameters. A fixed time disturbance observer (FTDO) is proposed to estimate the future tracking error information and unknown disturbances to improve the control performance of the attitude. Finally, the stability of the system is proved via the Lyapunov technique and a comprehensive simulation of HALE aircraft is provided. The simulation results show that the proposed control algorithm can enable the aircraft attitude angles to quickly response and maintain stability under gust wind. Compared with a single ADP control strategy, this paper's algorithm has advantages in response speed and error. [ABSTRACT FROM AUTHOR]

Modern aircraft such as missile and rocket, due to the large slenderness ratio of slender body vehicles, the influence of elastic deformation and vibration on navigation, guidance, and engine modules in simulation can not be ignored. For the problems of slow calculation speed and incapability of real-time simulation for time-domain simulation, by analyzing the time proportion of each calculation step under different computing scale, the dynamic parallel construction of octree is used to represent the aerodynamic parameter table under the environment of single and multi GPU. Meanwhile, an innovative parallel algorithm of element stiffness matrix based on finite element model is designed in GPU architecture. Accordingly, the optimized performance is enhanced through the adaptive hardware resources and rational use of shared memory. Furthermore, A multi-threaded asynchronous framework based on task queue and thread pool is proposed to realize the parallel task calculation with different granularities. The numerical result shows that the acceleration ratio of about 20 times in the single GPU condition can be obtained, and the acceleration ratio of at least 30 times can be obtained by the parallel computing of dual GPUs, enabling the real-time simulation of the flexible aircraft with 1200 elements within 20ms.

This paper makes use of sloshing reduced-order models to investigate the effects of sloshing dynamics on aeroelastic stability and response of flying wing structure. More specifically, a linear frequency-domain operator derived by an equivalent mechanical model is used to model lateral (linear) sloshing dynamics whereas data-driven neural-networks are used to model the vertical (nonlinear) sloshing dynamics. These models are integrated into a formulation that accounts for both the rigid and flexible behavior of aircraft. A time domain representation of the unsteady aerodynamics is achieved by rational function approximation of the fully unsteady aerodynamics obtained via the doublet lattice method. The case study consists of the so called Body Freedom Flutter research model in two different configurations with one or two tanks partially filled with liquid with a mass comprising 25% of the aircraft structure. The results show that linear sloshing dynamics are able to change the stability margin of the aircraft in addition to having non-negligible effects on rigid body dynamics. On the other hand, vertical sloshing acts as a nonlinear damper and eventually provides limit cycle oscillations after flutter onset.

The new aeronautical structures have become more flexible and light weight to reduce energy consumption, emissions, and noise, and to operate at high altitudes for long periods in the air, such as those required in the NASA Helios project. The increased structural flexibility of these aircraft has reignited concerns related to aeroelastic instabilities, such as flutter. Improving the techniques and methods used in aircraft certification flights is an important concern of the aeronautical community, because current standards and procedures do not provide recommendations and guidelines for aircraft with a high degree of flexibility. The techniques traditionally used in commercial airplanes cannot be used in this new aircraft concept, because they have a high degree of non-linearity in their flight dynamics. Current research indicates an increasing awareness about the importance of vision in the monitoring of UAV structural health. This work presents a new methodology to measure natural frequencies of aeronautical structures using a computer vision system. We also discusses new approaches to sense and acquire vibration data on aeroelastic certification flights test. These new approaches aim to reduce both the time required to identify the aeroelastic phenomenon and the size of the hardware that must be boarded on the aircraft, thus minimizing the risks and costs of the vibration tests. The advance of computer vision systems enables the use of cameras as a motion tracker sensor with millimeter precision and accuracy. Non-contact sensors are suitable for flutter analysis because they do not interfere with the dynamics of the aircraft. Therefore, this new methodology is able to process the obtained images and provide the user with the data about movements in a ready to use vector, at a reasonable cost. Using the data provided by this methodology the natural frequencies of the first bending modes were identified. This new methodology can be a user-friendly tool to support the Brazilian National Civil Aviation Agency - ANAC program called iBR 2020, which aims to certify small aircraft. [ABSTRACT FROM AUTHOR]

Turbulence and gusts cause variations in the aerodynamic forces and moments applied to the structure of aircraft, resulting in passenger discomfort and dynamic loads on the structure that it must be designed to support. By designing Gust Load Alleviation (GLA) systems, two objectives can be achieved: first, realizing higher passenger comfort; and second, reducing the dynamic structural loads, which allows the design of lighter structures. In this paper, a methodology for designing combined feedback/feedforward GLA systems is proposed. The methodology relies on the availability of a wind profile ahead of the aircraft measured by a Doppler LIDAR sensor, and is based on $H_{\infty}$ -optimal control techniques and a discrete-time preview-control problem formulation. Moreover, to allow design trade-offs between those two objectives (to achieve design flexibility) as well as to allow specification of robustness criteria, a variant of the problem using multi-channel $H_{\infty}$ -optimal control techniques is introduced. The methodology developed in this paper is intended to be applied to large aircraft, e.g. transport aircraft or business jets. The simulation results show the effectiveness of the proposed design methodology in accounting for the measured wind profile to achieve the two mentioned objectives, while ensuring both design flexibility and controller robustness and optimality. [ABSTRACT FROM AUTHOR]

This article aims to indicate the differences between rigid and flexible wing aircraft flying (FQ) and handling (HQ) qualities. The Simulation Framework for Flexible Aircraft was used to provide a generic cockpit environment and a piloted mathematical model of a bare airframe generic high aspect ratio wing aircraft (GA) model. Three highly qualified test pilots participated in the piloted simulation trials campaign and flew the GA model with both rigid and flexible wing configurations. The results showed a negligible difference for the longitudinal HQs between rigid and flexible wing aircraft. However, significant changes were indicated for the lateral/directional HQs of the flexible wing aircraft. A wing ratcheting phenomenon manifested itself during the roll mode tests, the spiral mode exhibited neutral stability and the Dutch roll mode shape changed from a horizontal to a vertical ellipse. The slalom task flight tests, performed to assess the FQs of the aircraft, revealed the degradation of both the longitudinal and lateral/directional FQs.

This article aims to indicate the differences between rigid and flexible wing aircraft flying (FQ) and handling (HQ) qualities. The Simulation Framework for Flexible Aircraft was used to provide a generic cockpit environment and a piloted mathematical model of a bare airframe generic high aspect ratio wing aircraft (GA) model. Three highly qualified test pilots participated in the piloted simulation trials campaign and flew the GA model with both rigid and flexible wing configurations. The results showed a negligible difference for the longitudinal HQs between rigid and flexible wing aircraft. However, significant changes were indicated for the lateral/directional HQs of the flexible wing aircraft. A wing ratcheting phenomenon manifested itself during the roll mode tests, the spiral mode exhibited neutral stability and the Dutch roll mode shape changed from a horizontal to a vertical ellipse. The slalom task flight tests, performed to assess the FQs of the aircraft, revealed the degradation of both the longitudinal and lateral/directional FQs. [ABSTRACT FROM AUTHOR]

This paper presents an improved process for the modeling, model order reduction (MOR), and control design of a flexible flying wing unmanned aerial vehicle based on the linear parameter-varying (LPV) technology. By applying the Lagrangian formulation and doublet lattice method, the aeroservoelastic (ASE) model of the flexible aircraft is first derived and then rewritten as an airspeed-dependent LPV system, which consists of rigid-body states, structural modes, aerodynamic states, etc. This leads to a high-order LPV model, and the stability of the system varies from stable to unstable when the speed of the aircraft exceeds the critical flutter speed. To facilitate the control design, an oblique projection-based MOR method is then adopted to reduce the order of the plant. Different from previous studies, the MOR method is improved to systematically generate a reduced-order LPV model with consistent states for both the stable and unstable dynamics parts. Finally, an output-feedback LPV controller is designed to simultaneously achieve the body freedom flutter suppression and longitudinal attitude control. Numerical simulation results show the effectiveness of the improved LPV-based ASE modeling, order reduction, and control design process for the flexible aircraft. [ABSTRACT FROM AUTHOR]

This paper focuses on the development of an explicit finite difference numerical method for approximating the solution of the inhomogeneous fourth-order Euler–Bernoulli beam bending equation with velocity-dependent damping and second moment of area, mass and elastic modulus distribution varying with distance along the beam. We verify the method by comparing its predictions with an exact analytical solution of the homogeneous equation, we use the generalised Richardson extrapolation to show that the method is grid convergent and we extend the application of the Lax–Richtmyer stability criteria to higher-order schemes to ensure that it is numerically stable. Finally, we present three sets of computational experiments. The first set simulates the behaviour of the un-loaded beam and is validated against the analytic solution. The second set simulates the time-dependent dynamic behaviour of a damped beam of varying stiffness and mass distributions under arbitrary externally applied loading in an aeroelastic analysis setting by approximating the inhomogeneous equation using the finite difference method derived here. We compare the third set of simulations of the steady-state deflection with the results of static beam bending experiments conducted at Cranfield University. Overall, we developed an accurate, stable and convergent numerical framework for solving the inhomogeneous Euler–Bernoulli equation over a wide range of boundary conditions. Aircraft manufacturers are starting to consider configurations with increased wing aspect ratios and reduced structural weight which lead to more slender and flexible designs. Aeroelastic analysis now plays a central role in the design process. Efficient computational tools for the prediction of the deformation of wings under external loads are in demand and this has motivated the work carried out in this paper.

Lightweight flexible aircraft suffers from unwanted oscillatory vibrations during aircraft manoeuvres. A recently developed distributed-delay signal (DZV) shaper is therefore proposed to be applied as a feedforward controller to alleviate the manoeuvre loads, as an alternative to traditional structural filters used routinely in this context. Structural filters are essentially linear low-pass filters with bandwidth below the significant flexible modes, applied to control signals generated either by the pilot's direct input or by the flight control system. It has been showed that if instead a properly tuned signal shaper is used, better performance can be achieved: first, the target modes are significantly attenuated while the responsiveness of the aircraft is less compromised and secondly, the oscillatory nature of the vibrations are reduced. The high fidelity simulation results on a full scaled dynamic model of a highly flexible blended wing–body (BWB) aircraft show that in comparison to traditional structural filters, signal shapers significantly reduce the wing root loading (forces and moments) which provides potential structural benefits. [ABSTRACT FROM AUTHOR]

The increase in aircraft flexibility has a great impact on aircraft flight quality. The traditional evaluation index of flight quality for rigid aircraft cannot meet the flight quality evaluation for the flexible aircraft. In this paper, the rigid and flexible air craft linearized flight dynamics models are established respectively. The frequency-domain and time-domain response are compared to analyze the performance of rigid and flexible aircraft. Aiming at the deficiency of the existing flight quality evaluation index for flexible aircraft, the existing control anticipation parameter (CAP) evaluation index is modified and the structural modal frequency is introduced into the evaluation index formula. The CAP values of the rigid aircraft and the flexible aircraft are calculated respectively. The results show that the CAP value of the high speed and large flexible aircraft is larger than that of the rigid air-craft, which might affect the flight quality level. [ABSTRACT FROM AUTHOR]

The flight dynamics and handling qualities of any flexible aircraft can be analysed within the Cranfield Aircraft Accelerated Loads Model (CA2LM) framework. The modelling techniques and methods used to develop the framework are presented. The aerodynamic surfaces were modelled using the Modified Strip Theory (MST) and a state-space representation to model unsteady aerodynamics. With a modal approach, the structural flexibility and each mode’s influence on the structure deflections are analysed. To supplement the general overview of the framework equations of motion, models of atmosphere, gravity, fuselage and engines are introduced. The AX-1 general transport aircraft model is analysed as an example of the CA2LM framework capabilities. The results showed that, according to the Gibson Dropback criterion, the aircraft with no control system lacks the stability and its longitudinal handling qualities are unsatisfactory. Finally, the steps for future developments of the CA2LM framework are listed within conclusions.

The flight dynamics and handling qualities of any flexible aircraft can be analysed within the Cranfield Aircraft Accelerated Loads Model (CA2LM) framework. The modelling techniques and methods used to develop the framework are presented. The aerodynamic surfaces were modelled using the Modified Strip Theory (MST) and a state-space representation to model unsteady aerodynamics. With a modal approach, the structural flexibility and each mode’s influence on the structure deflections are analysed. To supplement the general overview of the framework equations of motion, models of atmosphere, gravity, fuselage and engines are introduced. The AX-1 general transport aircraft model is analysed as an example of the CA2LM framework capabilities. The results showed that, according to the Gibson Dropback criterion, the aircraft with no control system lacks the stability and its longitudinal handling qualities are unsatisfactory. Finally, the steps for future developments of the CA2LM framework are listed within conclusions. [ABSTRACT FROM PUBLISHER]

A parametric reduced order methodology for loads estimation is described that produces a fast and accurate prediction of gust and manoeuvre loads for different flight conditions and structural parameter variations. The approach enables efficient prediction of the peak loads whilst maintaining the correlated time histories for different loads. It is then possible to determine correlated loads plots with reduced computation without losing accuracy. The effectiveness of the methodology is demonstrated by considering loads arising from families of gusts and pitching manoeuvres acting upon a numerical transport aircraft aeroservoelastic model with varying flight conditions and structural properties. [ABSTRACT FROM AUTHOR]

The objective of this paper is to investigate how the aeroelastic stability, particularly flutter, is affected by aircraft maneuvers. The authors' investigation is based on a comprehensive mathematical model of aircraft, which is achieved by seamlessly integrating all the disciplines pertinent to flight of aircraft. The aircraft is treated as an unstrained, flexible multibody system subject to unsteady aerodynamics. The bodies are fuselage, wing, and horizontal and vertical stabilizers, whose structures are modeled as beams in bending and torsion. The equations of motion are derived using Lagrange's equations in quasi-coordinates. The resulting equations are a set of nonlinear ordinary differential equations of relatively high order. The final model is used to determine flutter speeds of aircraft at steady level turn and steady climb at various altitudes. These maneuvers are especially chosen to keep the equations time invariant. The numerical results are given for a generic transport model (GTM). The stability of a GTM is affected by turn radius, climb angle, and altitude. The results for climbing flight can be extended to address stability of gliding flight. [ABSTRACT FROM AUTHOR]

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. [ABSTRACT FROM PUBLISHER]

A 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 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 or Boeing . Simulations are performed in the presence of a stochastic Dryden turbulence model combined with a 1-cosine (1-cos) shape wind gust. The model-following inner loop is validated via a comparison with an internal loop of flexible mode stabilization, designed via a standard linear quadratic regulator with output feedback. [ABSTRACT FROM AUTHOR]

The effect of flexibility on the flight dynamics of flexible aircraft has been shown to be quite significant, especially as the frequencies of the elastic modes become lower and approach those of the rigid-body modes. The handling characteristics may become complex drastically. Consequently, the need to model accurately the dynamics of such vehicles, and to develop valid simulations, is becoming particularly acute. The key theories and results aimed at developing integrated equations of motion for a generic flexible aircraft model are reviewed. First, the significance of obtaining a valid mathematical model of an aeroelastic vehicle is outlined. After that, the theories of modeling dynamic systems are elaborated, as well as the selection of the floating reference frames by which the motion of flexible aircrafts can be described. Then followed by a discussion on the existing models, their advantages and disadvantages are also summarized. Based on the investigations, a novel floating frame (instant-coordinates) was put forward; it contains the advantages of both the mean axes and the quasi-coordinates. A new formulation, integrates seamlessly all the necessary materials from the areas of flight dynamics, structural dynamics, aerodynamics and controls, was formulated through the use of Lagrange's equations in terms of the new floating frame. 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, provided certain criteria are satisfied. In the end, some comments and future directions on modeling the highly flexible aircrafts are addressed. [ABSTRACT FROM AUTHOR]

The issue of this paper is to compute an LPV/LFT aircraft model, i.e., a continuum of models, by approximately interpolating a finite gridding of very high-order models. Reduced-order models of the flexible part are computed and interpolated using a rescaled companion state-space representation. [ABSTRACT FROM AUTHOR]

This paper investigates projectively lag synchronization of a new hyperchaotic system (proposed by Yang et al. in Nonlinear Anal., Real World Appl., 10:1601, ), with certain/uncertain parameters. Based on the Lyapunov function constructing method and application of the Babarǎt lemma, projectively lag synchronization is achieved for the new hyperchaotic system by designing nonlinear controllers. Furthermore, a novel method is proposed to identify unknown parameters based on projectively lag synchronization of the new hyperchaotic system. Finally, several numerical simulations are given to verify and test the correctness and effectiveness of the novel methods we proposed. [ABSTRACT FROM AUTHOR]