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Purpose This study aims to predict dynamic responses of aileron and spoiler control surfaces in subsonic flight via the use of surrogate models. The prepared reduced order models prove useful when quick estimations for a large number of variations are required. Design/methodology/approach The linear frequency domain (LFD) method was used for the simulation study. Each surrogate contained a database of 100 control surface dynamic responses over a spectrum of 200 harmonics computed with LFD. To interpolate new results, the DLR surrogate modelling toolbox, SMARTy, was used. The database’s samples were prepared in a Halton sequence, making interpolation reliable. The surrogate’s parameter space was the Mach number, Reynold’s number, angle of attack, control surface deflection angle and the control surface chord length. Findings The LFD method proved effective for the mentioned purpose: the surrogates were accurate, up to 15% of relative error, in reproducing dynamic responses of aileron and spoiler deflections at low speed, within the limitations of flow field linearity, as well as surrogate prediction capability. The restrictions of the surrogate, and the reasoning thereof, are also presented in detail in the study. Future load alleviation studies are a potential of the findings here. Originality/value LFD is an innovative technique for load prediction and alleviation studies. This paper provides a reference for engineers wishing to use the method for the two mentioned control surfaces, or the like.

This paper describes a frequency-domain method for the identification of multi-input control equivalent turbulence input models. Such models can be identified from flight data gathered in specific ...

This paper investigates multi-objective optimization of coordinated patrolling flight of multiple unmanned aerial vehicles in the vicinity of terrain, while respecting their performance parameters....

Whirl flutter of a tiltrotor aircraft is a complex aeroelastic phenomenon and it can result in catastrophic consequences. The deflection of an aileron mounted on a wing has the potential to solve this fatal problem. Whirl flutter suppression using an actively controlled aileron is studied in this study. Firstly, a semi-span aeroelastic model is established for the whirl flutter problem using the Hamilton principle. This model is composed of three parts: a rigid rotor, a rigid nacelle and a flexible wing, and the effect of the aileron deflection on the aeroelastic responses is also taken into consideration through a quasi-steady aerodynamic model. In addition, the accuracy of this aeroelastic model is validated with the results of two different wind-tunnel tests. Then, an LQR controller is developed to control the dynamic deflection of the aileron, and a full-dimensional state observer is built to estimate the state of the time-invariant system of a tiltrotor aircraft. Finally, simulations are carried out using the aeroelastic model and the LQR controller at different flight conditions to study the influence of the aileron deflection on whirl flutter. The simulation results demonstrate that the flutter boundary speed can be improved by 18.1% with the active deflection of the aileron, compared with the uncontrolled condition.

Purpose The transonic buffet is a complex aerodynamics phenomenon that imposes severe constraints on the design of high-speed vehicles, including for aircraft and space launchers. The origin of buffet is still debated in the literature, and the control of this phenomenon remains difficult. This paper aims to propose an original scenario to explain the origin of buffet, which in turn opens promising perspectives for its alleviation and attenuation. Design/methodology/approach This work relies on the use of numerical simulations, with the idea to reproduce the buffet phenomenon in a transonic aileron designed for small space launchers. Two numerical approaches are tested: unsteady Reynolds averaged Navier–Stokes (URANS) and large-eddy simulation (LES). The numerical predictions are first validated against available experimental data, before to be analysed in detail to identify the origin of buffet on the studied configuration. A complementary numerical study is then conducted to assess the possibility to delay the onset of buffet. Findings The buffet control strategy is based on wall cooling. By adequately choosing the wall temperature, this work shows that it is feasible to delay the emergence of buffet. More precisely, this paper highlights the crucial role of the subsonic flow inside the boundary layer, showing the existence of upstream travelling pressure waves that are responsible for the flow coupling between both sides of the airfoil, at the origin of the buffet phenomenon. Originality/value This paper proposes a new scenario to explain the origin of buffet, based on the use of a Fanno and Rayleigh flow analogies. This approach is used to design a control solution based on a modification of the wall temperature, showing very promising results.

In this paper, we review the concept of Lie brackets and how it can be exploited in generating motion in unactuated directions through nonlinear interactions between two or more control inputs. Applying this technique to the airplane flight dynamics near stall, a new rolling mechanism is discovered through nonlinear interactions between the elevator and the aileron control inputs. This mechanism, referred to as the Lie Bracket Roll Augmentation (LIBRA) mechanism, possesses a significantly higher roll control authority near stall compared to the conventional roll mechanism using ailerons only; it produces more than an order-of-magnitude stronger roll motion over the first second. The main contribution of this paper is to study the nonlinear flight physics that lead to this superior performance of the LIBRA mechanism. In fact, the LIBRA performance in free flight (six DOF) is double that in a confined environment of two-DOF roll-pitch dynamics. The natural feedback from the airplane motion (roll, yaw, and sideslip) into the LIBRA mechanism boosts its performance through interesting nonlinear interplay between roll and yaw, while exploiting some of the changes in the airplane characteristics near stall.

[ES] El objetivo principal del presente trabajo de final de grado es analizar la influencia de las superficies de control primarias (alerón, timón de profundidad y timón de dirección) en una aeronave de tipo comercial, empleando para ello el Boeing 737-800. El análisis realizado se basa en un estudio en CFD (Computational Fluid Dynamics) por medio del uso del programa STAR-CCM+. Para ello, se estudian diversas configuraciones de dichas superficies de control en la aeronave, diferenciándose cuatro casos claros: Deflexión de alerón, deflexión de timón de profundidad, deflexión de alerón y timón de profundidad, y deflexión de alerón y timón de dirección. Para cada uno de ellos, se analizan dos deflexiones de cada superficie de control distintas a dos números de Mach. El documento se encuentra estructurado de tal forma que permita, primeramente, entender el escenario general del tema a tratar, así como los objetivos del proyecto y la motivación para su realización. Después, se lleva a cabo una explicación del marco teórico que engloba los aspectos propios de la mecánica del vuelo longitudinal y lateral-direccional, así como los relacionados con CFD. Posteriormente, se realiza una exposición y análisis de los resultados obtenidos por medio del estudio, para finalizar obteniendo las conclusiones más relevantes y elaborando un presupuesto que refleje los costes asociados al uso de los Software empleados y las horas invertidas por parte del realizador del estudio y la tutora del mismo., [EN] The main objective of this final degree project is to analyze the influence of the primary control surfaces (aileron, elevator and rudder) in a commercial aircraft, in this case the Boeing 737-800. The analysis performed is based on a CFD (Computational Fluid Dynamics) study using the STAR-CCM+ program. To do this, various configurations of said control surfaces in the aircraft are studied, differentiating four clear cases: Aileron deflection, elevator deflection, aileron and elevator deflection, and aileron and rudder deflection. For each of them, two different deflections of each control surface at two Mach numbers are analyzed. The document is structured in such a way that it allows, firstly, to understand the general scenario of the topic to be discussed, as well as the objectives of the project and the motivation for its realization. Afterwards, an explanation of the theoretical framework is carried out, which encompasses the aspects of the mechanics of longitudinal and lateral-directional flight, as well as those related to CFD. Subsequently, a presentation and analysis of the results obtained through the study is carried out, to conclude by obtaining the most relevant conclusions and preparing a budget that reflects the costs associated with the use of the Software used and the hours invested by the director of the study and its tutor.

Most modern aircraft employ discrete ailerons for roll control. The induced drag, rolling moment, and yawing moment for an aircraft depend in part on the location and size of the ailerons. In the present study, lifting-line theory is used to formulate theoretical relationships between aileron design and the resulting forces and moments. The theory predicts that the optimum aileron geometry is independent of prescribed lift and rolling moment. A numerical potential flow algorithm is used to evaluate the optimum size and location of ailerons for a wide range of planforms with varying aspect ratio and taper ratio. Results show that the optimum aileron design to minimise induced drag always extends to the wing tip. Impacts to induced drag and yawing moment are also considered, and results can be used to inform initial design and placement of ailerons on future aircraft. Results of this optimisation study are also compared to theoretical optimum results that could be obtained from morphing-wing technology. Results of this comparison can be used to evaluate the potential benefits of using morphing-wing technology rather than traditional discrete ailerons.

Accepted: 2021-02-28, 資料番号: PA2120065000

The implementation of interconnect gain from aileron to rudder surface on the majority of the aircraftis to decrease sideslip which is generated because of adverse yaw with the movement of control stick in lateral axis and also enhances the turning rate performance.The Aileron to Rudder Interconnect (ARI)involves significant part to decouple the Dutch roll oscillations from roll rate response to aileron command. ARI is feed-forward gain whichis susceptible to aircraft system uncertainty. Incorrect ARI gain can lead to side slip buildup which can cause aircraft to depart in case of fault scenarios. Four systematic ARI design methods are proposed. One of the proposed methods which use the norm of ARI transfer function at roll damping frequency is suitable for online reconfiguration of control law. The reconfiguration of ARI gain is illustratedwith the simulation responses of fault scenario case of aileron surface damage.

This paper is concerned with small UAV (unmanned aerial vehicle) control algorithm development during landing. The UAV landing problem for small UAVs (less than 20 kg) remains actual despite the great amount of books, papers, and monographs devoted to this topic. In this paper, a model with V-shaped keel is under consideration. Mechanization of such a vehicle consists of flaperons and ailerons. To describe a flight, a system of nonlinear differential equations is developed. The automatic control both for longitudinal and lateral motion as well as for thrust is designed. The stability analysis of the controlled system is conducted by plotting stability regions in the space of the feedback coefficients. The results of a numerical modeling of a flight with external disturbances are given.

The paper is devoted to the problem of safe flight in vortex hazard conditions. A vortex wake is always produced behind flying aircraft. This vortex wake is invisible to other aircraft following it. Therefore, getting into a vortex wake from a preceding aircraft often becomes unexpected and quite often leads to flight incidents. This is confirmed by the statistics of flight accidents. With the introduction of aircraft weighing more than 500 tons, the problem of vortex safety has only sharpened. The fact is that the old standards that define safe intervals between flying aircraft are still active. These standards are given in this paper. It is also shown that even if these standards are observed, the flight of a medium-class aircraft at the same level as a heavy-class aircraft is unsafe. To study the effect of the vortex wake from the preceding aircraft, a special computational software complex based on the discrete vortex method was developed. This complex has passed the necessary testing and state registration. A number of measures were carried out to validate and verify the developed complex, confirming the operability of the programs included in it and the reliability of the results obtained. On the basis of this computational and software complex studies on the effect of the A380 aircraft vortex wake on MS-21 class aircraft were performed. The studies were undertaken for various altitudes and speeds of the A380 aircraft and for its various flight configurations. As a criterion for the MS-21 aircraft safe flight, the roll moment that occurs due to falling into the vortex wave, in particular, its coefficient was chosen. If the coefficient of the MS-21 aircraft roll moment when it hits the vortex wake exceeds that from the ailerons, such a flight is considered dangerous. The distances behind the A380 that are unsafe for the MS-21 aircraft flight are given.

We describe the Intelligent Autopilot System (IAS), a fully autonomous autopilot capable of piloting large jets such as airliners by learning from experienced human pilots using Artificial Neural Networks. The IAS is capable of autonomously executing the required piloting tasks and handling the different flight phases to fly an aircraft from one airport to another including takeoff, climb, cruise, navigate, descent, approach, and land in simulation. In addition, the IAS is capable of autonomously landing large jets in the presence of extreme weather conditions including severe crosswind, gust, wind shear, and turbulence. The IAS is a potential solution to the limitations and robustness problems of modern autopilots such as the inability to execute complete flights, the inability to handle extreme weather conditions especially during approach and landing where the aircraft’s speed is relatively low, and the uncertainty factor is high, and the pilots shortage problem compared to the increasing aircraft demand. In this paper, we present the work done by collaborating with the aviation industry to provide training data for the IAS to learn from. The training data is used by Artificial Neural Networks to generate control models automatically. The control models imitate the skills of the human pilot when executing all the piloting tasks required to pilot an aircraft between two airports. In addition, we introduce new ANNs trained to control the aircraft’s elevators, elevators’ trim, throttle, flaps, and new ailerons and rudder ANNs to counter the effects of extreme weather conditions and land safely. Experiments show that small datasets containing single demonstrations are sufficient to train the IAS and achieve excellent performance by using clearly separable and traceable neural network modules which eliminate the black-box problem of large Artificial Intelligence methods such as Deep Learning. In addition, experiments show that the IAS can handle landing in extreme weather conditions beyond the capabilities of modern autopilots and even experienced human pilots. The proposed IAS is a novel approach towards achieving full control autonomy of large jets using ANN models that match the skills and abilities of experienced human pilots and beyond.

A nonlinear controllability analysis for fixed-wing aircraft has been performed in an earlier effort by the authors, which revealed a novel roll mechanism due to a nonlinear interaction between elevator and aileron control inputs. In this effort, we perform a detailed investigation of this novel roll mechanism, called Lie Bracket Roll Augmentation (LIBRA). First, we show the nonlinear flight physics associated with the LIBRA mechanism. Second, using the Fliess functional expansion, we perform a theoretical study of the effectiveness (degree of controllability) of the LIBRA mechanism in comparison to the conventional mechanism (using ailerons only). Third, to simulate the airplane response to a LIBRA input, we cast the problem of executing the LIBRA mechanism as a nonholonomic motion planning problem. In such a problem, a Lie bracket input is applied to generate motion along an unactuated direction. A Lie bracket input represents a nonlinear interaction between two (or more) control inputs to steer the system along a direction that is not directly actuated by any of these inputs (or their linear combinations). In this language, the LIBRA mechanism is simply a Lie bracket interaction between the elevator and aileron control inputs. We modify existing nonholonomic motion planning algorithms for systems with drift to be more feasible for flight control applications with bounded controls. We show that the LIBRA novel roll mechanism is superior compared to the conventional one during stall, where the aileron sensitivity degrades. In particular, the novel roll mechanism can provide an order of magnitude enhancement in the rolling capability over the conventional roll mechanism near stall.

Aileron to Rudder Interconnect (ARI) gain is implemented on most fighter aircraft, primarily to reduce the side slip produced due to adverse yaw from pilot lateral control stick input and to improve the turn rate response. A systematic and non-iterative design procedure for ARI gain is proposed herein based on the evaluation of a transfer function magnitude at the aircraft roll mode frequency. The simplicity of the proposed method makes it useful for real-time flight control law reconfiguration in situations where the aileron control authority is diminished due to damage. This is demonstrated by a simulation example considering an aileron surface damage scenario.

This work aims to model transonic airfoil–gust interaction and the gust response on transonic aileron-buzz problems using high-fidelity computational fluid dynamics (CFD) and the Long Short Term Me...

For a mechatronic system, the lack of robustness is mainly due to an ignorance of variability, which is unavoidable in every designed system at all stages of system development and during its life cycle. In fact, the inherent imperfections of manufacturing processes and operational variations such as material wear or multiphysical effects (thermal, vibrations, etc.) involve parametric variations, which can degrade the proper system function. To ensure a high level of quality and to improve design robustness, the deviations between actual and target definition should be restricted by specified tolerances. Moreover, it is important to understand the relationship between parameter deviations and system behavior. Thus, in this article, a variational approach is proposed to identify performance sensitivity to parameter deviations in a mechatronic system and to determine influential parameters that affect significantly the proper functioning of the system and its behavior. Furthermore, this approach is used to specify the admissible parameter deviations for which the given specifications can be guaranteed. The complete approach is applied to an example in the aeronautic field: an electromechanical actuator driving an aircraft primary flight control surface (aileron).

Aileron is a control surface that functions as a regulator of roll motion. The movements of the ailerons are opposite to the left and right sides. Previous studies have shown that graphs of hinge moment coefficient (C hm ) values increases with increasing angle of attack. This study is to determine the aerodynamic characteristics of aileron by combining the surface area of the vane into the aileron by varying the aileron’s deflection. The calculation is performed using a numerical method in two dimensions (2D) commercial CFD simulation software. The results of the study concluded that the hinge moment coefficient for modified airfoil at δA = -20°, 0°, and 20° was -0.071, 0.078, and 0.177, respectively. These values are smaller when compared to C hm value in basic aileron that was -0.094, 0.095, and 0.201, respectively.

Variable-camber technology is considered an effective way to adaptively improve the aerodynamic performance of aircraft under various flight conditions. This paper studies the aerodynamic characteristics of the trailing-edge variable-camber technology by means of Computational Fluid Dynamics (CFD) and a drag decomposition method. Trailing-edge variable-camber technology can be simply realized by the continuous deflection of the flaps and ailerons of a wing. A supercritical airfoil is used to study the two-dimensional effect of variable-camber technology, and a wide-body airplane model is used to validate the three-dimensional improvement in the wing's airfoil made by variable-camber technology. An optimization strategy for airfoil that incorporates variable-camber technology is proposed. The optimization results demonstrate that the proposed method can obtain better results than the traditional segregated shape optimization.

In this paper we consider Model Predictive Control (MPC) design for roll angle control for a Fixed-Wing Unmanned Aerial Vehicle (UAV) with Multiple Segmented Control Surfaces. The challenge of roll angle control for a Fixed-Wing UAV consists of switching between inner and outer aileron-pairs with hard constraints due to safety, energy saving and switching actuators. The novelty consists of formulating a hybrid-control problem as a switched linear constrained MPC-QP problem and switched state observer design for Fixed-Wing UAV. A fast novel QP-solver based on the active-set QP-solver Hildreth is developed to meet the real-time implementation target, which is to stay below the sampling time of Ts = 10 ms. The designed MPC controllers are simulated using Matlab. Simulations and the CPU-time from the improved QP-solvers show MPC to be a very good choice for real-time roll angle control for Fixed-Wing UAVs., In this paper we consider model predictive control (MPC) design for roll angle control for a fixed-wing unmanned aerial vehicle (UAV) with multiple segmented control surfaces. The challenge of roll angle control for a fixed-wing UAV consists of switching between inner and outer aileron pairs with hard constraints due to safety, energy saving and switching actuators. The novelty consists of formulating a hybrid control problem as a switched linear constrained MPC-QP problem and switched state observer design for fixed-wing UAV. A fast novel QP-solver based on the active-set QP-solver Hildreth is developed to meet the real-time implementation sampling time of Ts = 10 ms. The designed MPC controllers are simulated using Matlab. Simulations and the CPU-time from the improved QP-solvers show MPC to be a very good solution for real-time roll angle control of fixed-wing UAVs.

Purpose A standard automatic flight control system – autopilot – will become required equipment of the future aircraft, operating in the common sky. For a specific group of aircraft, they are too expensive and too energy-consuming solutions. This paper aims to present the concept of an automatic flight control system that overcomes those limitations. Design/methodology/approach The proposed automatic flight control system uses the trim tabs in all prime flight controlling surfaces: elevator, ailerons and rudder, for stabilizing and controlling the steady flights of an aircraft. Findings The results of an aeroplane flight controlled with the use of trim tabs simulation tests and remarks have been presented and discussed. The simulation was conducted in real-time hardware in the loop environment. The stabilization of the flight was achieved in performed test scenarios. Originality/value The possibility to control an aircraft with coordinated deflections of the trimming surfaces is a beneficial alternate to those currently used and can be recommended for use in the next-generation aircraft.

Extensive research in recent years has focused on improving the current loss-of-control prevention systems and developing new strategies for safe path planning of the impaired aircraft. Success in developing such systems requires a comprehensive perception of the influence of damage on the aircraft’s dynamic behaviour and performance, and the effect of various failure degrees on the flight envelope confinement and the remaining safe maneuvers. This paper comprehensively describes the effects of lateral control surface failure on the NASA Generic Transport Model (GTM) flight envelope, defined by a set of attainable steady-state maneuvers herein referred to as trim points. The study utilises a large database of high-fidelity maneuvering flight envelopes computed for the unimpaired case and wide ranges of the aileron and rudder failure cases at different flight conditions. Flight envelope boundary is rigorously investigated, and the key parameters confining the trim points at different boundary sections are identified. Trend analyses of the impaired flight envelopes and the corresponding limiting factors demonstrate the effect of various failure degrees on the remaining feasible trim points. Results can be employed in emergency path planning with potential uses in the development of aircraft resilient control and upset recovery systems.

Aiming at the problem of energy collection and consumption of solar-powered UAV under different maneuvers, an energy collection model combining solar-powered UAV attitude and solar radiation was established. Firstly, the ASHRAE model was used as the solar radiation estimation model in the flight environment. Secondly, the energy harvesting model of the solar-powered UAV is deduced by the relationship between the attitude of the UAV, the relative position of the sun and the earth. The experimental data was used to establish the energy consumption model of solar-powered UAV in different flight states. Finally, the flight time of the same mission path under different operating modes is obtained through simulation. Simulation results show that when the solar-powered UAV is hovering at a height of 3 000 meters and a radius of 50 meters, the flight time of aileron operation is 7.19% longer than power differential operation; when hovering at a radius of 170 meters, the flight time of the two operations is almost equal; when hovering at a radius of 600 meters and flying in a line, the flight time of the power differential operation is about 1.08% longer than that of the aileron operation.

Based on mechanism research of dynamic coupling and inertial coupling of flying wing UAV at high attack angle or stall maneuver flight, a decoupling compensator is designed to realize the longitude and lateral channels, the aileron and the rudder channel. The maneuver control law is designed based on the robust servo method, which solves the problem of rapid maneuvering of the flying wing. The maneuvering boundary protection controller effectively reduces the risk of boundary during the maneuver flying. At last, the flying tests verified the engineering feasibility of decoupling controller and robust maneuvering controller.

Nixus is the world’s first fly-by-wire (FBW) sailplane, with a custom 92-foot span, 53.3 aspect ratio wing. With the second-largest-aspect-ratio wing ever built for a crewed airplane, Nixus faces unique challenges for aeroelastic design, requiring detailed study and special considerations for safe operation. The use of an FBW system for the wing control surfaces allows the exploration of new strategies for automatic flap positioning, tailored aileron deflections, load alleviation, and, in the future, aeroelasticity control. This paper briefly describes some aspects of this sailplane’s fabrication and covers in detail the ground vibration test (GVT) and use of the GVT results in the flight envelope expansion campaign. The preparation and execution of the GVT are presented, including aspects related to the involvement of Cal Poly’s Aerospace Engineering undergraduate students in this process as part of an extracurricular activity. Results of the GVT, comparison with finite element analysis (FEA), and flutter predictions from the test-verified FEA model are provided. Additionally, the execution of flutter flight tests that used the FBW system to excite the wing is presented.

This paper details the development of a high redundancy Unmanned Aerial Vehicle (UAV) platform with a sliding mode fault tolerant control system. The high redundancy UAV platform is of a Tilt Wing design, utilising four wings, each with a propeller. Conventional fixed wing aircraft control surfaces are also implemented i.e. ailerons, elevators and rudders. The benefits of such a platform is it combines the efficient forward flight of fixed wing aircraft, with the agility and Vertical Take Off and Landing (VTOL) capability of a conventional multi-rotor aircraft. Results show good forward flight performance despite the presence of actuator faults/failures.

Aircraft flight control usually requires driving flaps and ailerons. However, the air drags increase significantly due to the corners of flaps and aileron. Especially, the gap between mother wing and flap / aileron causes a drag increase. Therefore, studies are being conducted on morphing wings that smoothly and greatly deform the wing surface. For aircraft wing, it is needless to say that strength is important to sustain lift and drag for the aircraft during the flight. For morphing wings, in addition, actuators must be mounted inside the wing to enable the morphing deformation. Moreover, for the aircraft wing, weight is quite important. Therefore, carbon fiber reinforced plastic (CFRP) is currently most suitable for aircraft wing structural materials. However, it is difficult to mold CFRP so that it has sufficient strength and can be morphed. In this study, by using CFRP, the morphing wing structure was prototyped with targeting a small unmanned aerial vehicle (UAV) weighing 3 kg. The CFRP lattice structure that enables morphing deformation was designed and manufactured by applying the electrodeposition resin molding (ERM) method which was developed by the authors. In the ERM method, firstly, the carbon fiber was fixed with a jig according to the designed morphing wing structure, and immersed in the electrodeposition solution. Secondly, the epoxy polymer particle in the solution were electrophoresed and impregnated between carbon fibers. After thermal curing, the morphing wing structure was fabricated. Further, the loading-unloading torsion and bending tests of the morphing wing structure were carried out. Smooth morphing deformation and sufficient strength were confirmed.

This paper presents autonomous landing controller design of unmanned aerial vehicles (UAVs). Feedback linearization approach with anti windup scheme is designed as autonomous landing controller. Anti windup basically controls the integrator component accumulation in the autopilot design and restricts the output within saturation limits. Autonomous landing of a fixed wing UAV consists of approach, glideslope and flare phases. During approach, aerial vehicle aligns itself with runway and reduces its lateral deviations with respect to runway central line. In glide slope phase, aerial vehicle maintains a fixed flight path angle and descends with a constant sink rate. In flare phase, aerial vehicle follows an exponential trajectory and descends with a lower sink rate which keeps reducing further as it goes to lower altitudes. Flare controller is designed with integrator and antiwindup scheme is used to handle the controller output within saturation limits. Landing is primarily a longitudinal mode operation but due to disturbances and coupling lateral and directional modes also gets activated. In this paper pitch angle, roll angle and yaw angle nonlinear dynamic equations are linearized to obtain the control commands in terms of elevator deflection, aileron deflection and rudder deflections. Similarly total velocity of aerial vehicle being an important parameter is controlled using thrust command. A first order linearized model of velocity is used to obtain thrust control command. Autonomous landing of UAV with feedback linearization controller and anti windup scheme is simulated with Six-DOF model of AE2 UAV. The algorithm is implemented with wind disturbances to show the autonomous landing performance of UAV.

A tri-copter unmanned aerial vehicle (UAV) with individually tilted main wings for various flight maneuvers is presented. In contrast to conventional tilt wing UAVs, thrust vectoring and control of the directions of main wing surfaces are attained by tilting main wings individually. Such individually tilted main wings allow more efficient maneuvers of the UAV. Even though the UAV is a tri-copter, it can compensate the reaction torque created by the tail motor by the individual control of the main wings. Few publications have appeared in the literature, dealing with tri-copter UAV with individually tilted main wings. Dynamic model of the UAV is established via the Newton-Euler formulation and effect of individual wing control is examined. The PID controller and PI controller for control of flight maneuvers are implemented and used for simulations and experiments. Results of simulations and experiments are presented for validation of flight performance of the UAV, including the roll rate 80% larger than that of the conventional UAV with ailerons as control surfaces.

Actuation failure is one of the causes of loss of control in flight accidents. Aircraft usually have multiple redundant actuators to mitigate failures, and Failure Detection and Isolation Systems (FDIS) are used to diagnose failures and reconfigure software/hardware to enhance safety. However, the large number of redundant actuators interferes with the FDIS. To detect and isolate failures in fixed-wing aircraft with redundant actuators, this work proposes the combined use of two different strategies of the Two-Stage Kalman Filter. A Supervisory Loop is included using heuristics and statistics to diagnose the actuators, and a Feed-Forward Differential is implemented to improve the isolation process without interfering with the aircraft flight. The solution is evaluated in the detection of an aileron failure in a Boeing 747 simulator.

In this paper, a multi-input/multi-output servomechanism primary flight controller is presented. Also discussed are flight data related to the system performance and pilot assessments of handling q...

Freeplay in flight control surface linkages leads to nonlinear aeroelastic behavior including limit cycle oscillations (LCO). LCO due to a constant freeplay has been extensively studied, whereas th...

Extensive research in recent years has focused on developing flight envelope estimation methods to improve current loss of control prevention and recovery systems. Such methods are practically efficient only if they are able to evaluate in real time the new flight envelope of damaged aircraft based on the altered dynamics. Due to nonlinear dynamics of aircraft, common approaches to estimate the entire flight envelope of high-fidelity models are numerically intensive and their real time implementation is computationally impossible. So current methods are based on reduced complexity models or flight envelopes are determined locally. This paper presents a novel method to estimate the global flight envelope of impaired aircraft in real-time for any unknown failure degree. In the proposed method, first, numerous flight envelopes are evaluated using a high fidelity model at various failure degrees and different flight conditions and prepared as training data. Then multiple feedforward neural networks are trained offline by a Bayesian regularization backpropagation algorithm. Finally, the trained networks are used to estimate flight envelopes in real time. The method is applied to rudder and aileron failure cases of the NASA Generic Transport Model. Results show that the estimated flight envelopes are good approximations of the high fidelity global flight envelopes.

In this paper, the problem of aircraft spin recovery is solved as a trajectory optimization problem using direct multiple shooting method with time and altitude-loss as cost functions to be minimized. A stable oscillatory spin state is chosen as the initial condition and the optimal control inputs required to transfer the aircraft to a steady level-flight trim state are determined. Optimal spin recovery simulations are carried out with scaled bounds of control inputs to determine their effectiveness and feasibility for recovery in the case of total control failures. It is shown that spin recovery is feasible in the case of aileron failure, and only arresting of yaw-rate is possible in the case of rudder and throttle failures. Optimal spin recovery simulations are also carried out to determine the effects of wind and aerodynamic forces. It is found that spin recovery solutions are sensitive to initial conditions in the presence of wind, and increase of lift and decrease of drag reduce the altitude-loss of the maneuver.

Experimental studies into aircraft stability and performance can be enhanced by using a rig in which the aircraft model support approximates free flight within a wind tunnel. Such multi-degree-of-freedom wind tunnel rigs often impose kinematic restrictions on the aircraft model's translational motion. This study investigates these kinematic effects, with particular attention to a spherical constraint where the aircraft is held at the end of afixed length pivoting arm. Here the motions of the aircraft and kinematic constraints are derived as differential-algebraic equations and assessed numerically. The impact is found mainly on translational motion with negligible effect on the aircraft's rotation. A concept to reduce these kinematic effects on the aircraft's motion by applying an external force onto the aircraft is proposed. This compensation, which partially accounts for the constraints on the aircraft motion, is shown to reduce the effects of the arm, allowing for improved physical simulation.

This paper presents the design and manufacturing of a new morphing wing system carried out at the Laboratory of Applied Research in Active Controls, Avionics and AeroServoElasticity (LARCASE) at the ETS in Montréal. This first version of a morphing wing allows the deformation of its trailing edge, denote by Morphing Trailing Edge (MTE). In order to characterize the technical impact of this deformation, we compare its performance with that of a rigid aileron by testing in the LARCASE’s price—Païdoussis subsonic wind tunnel. The first set of results shows that it is possible to replace an aileron by a MTE on a wing, as an improvement was observed for the MTE aerodynamic performances with respect to the aileron aerodynamic performances. The improvement consisted in the fact that the drag coefficient was smaller, and the lift-to-drag ratio was higher for the same lift coefficient. Keywords: Aerodynamic performances, Aileron, Morphing wing, Trailing edge, Wing tunnel tests

This paper deals with the actuation system design of a full-scale morphing aileron for regional aircraft. The aileron is allowed to smoothly change its geometrical configuration and perform the in-flight transition from a baseline shape to a set of optimal morphed ones pre-defined on the basis of aerodynamic requirements. The design of such innovative aileron is aimed not only at substituting the conventional aileron installed on a real aircraft but also to provide additional functionality. The aileron is free to rotate around its main hinge axis and it is also allowed to smoothly modify camber with two independent actuation systems. In such manner it can be used also during cruise with a symmetric deflection between the two half wings in order to reduce drag in off design condition. To accomplish variable aileron shape, a rigid-body mechanism was designed. The proposed aileron architecture is characterised by segmented adaptive ribs rigidly linked each other with spanwise reinforcements such as spars and stringers in a multi-box arrangement. Each rib is split into two movable plates connected by means of rotational hinges in a finger-like mechanism. The mechanism is driven by a load-bearing actuator by means of a kinematic chain opportunely tied based on the structural requirements in terms of shape to be matched and load to be withstood. The proposed device is an innovative arrangement of the quick-return mechanism composed of a beam leverage, commercial linear guides and a crank. The actuator shaft is directly inserted in the crank, which transmits the rotation to the linear guide that slide along a rail moving upward or downward the beam thus resulting in a camber variation. The entire aileron is moved by three leverages internally contained and distributed along the first two bays while the most external ribs are considered passive and their movement slaved. Two actuation layouts are analytically and numerically studied, the analytical theory is presented and validated by means of a multi-body simulation. Moreover, a linear static analysis was carried out under the hypothesis of glued contact between linear guides components simulating a jamming condition. This assumption has been formulated because it represents the most severe condition that envelop all the operative loads to which the actuation system is subjected. The analyses conducted are preliminarily aimed to verify that no failure occur under the imposed loads. In this first design loop, the vertical static force acting on the linear carriage exceeded allowable value and then a new configuration with double-sided linear guides was then investigated.

A parallelized flight simulator for the dynamic analysis of airborne wind energy (AWE) systems for ground- and fly-generation configurations is presented. The mechanical system comprises a kite or fixed-wing drone equipped with rotors and linked to the ground by a flexible tether. The time-dependent control vector of the simulator mimics real AWE systems and it includes the length of the main tether, the geometry of the bridle, the torque of the motor controllers of the rotors, and the deflections of ailerons, rudder and elevator. The use of a lagrangian formulation with a minimal coordinate approach and discretizing the main tether as a chain of inelastic straight rods linked by ideal (dissipative-less) rotational joints, yielded a non-stiff set of ordinary differential equations free of algebraic constraints. Several verification tests, including a reel-in maneuver that admits an analytical solution, are presented. The efficiency of the parallelization with the number of tether segments, and trade-off analysis of the lagrangian and hamiltonian formulations are also considered. The versatility of the simulator is highlighted by analyzing two maneuvers that are relevant for AWE scenarios. First, the simulator is used to compute periodic figure-of-eight trajectories with an open-loop control law that varies the geometry of the kite's bridle, as frequently done in ground-generation AWE systems. Second, an unstable equilibrium state of a tethered drone equipped with two rotors for energy harvesting is stabilized by implementing a closed-loop control strategy for the deflection of the control aerodynamic surfaces. This work was supported by the Ministerio de Economía, Industria y Competitividad of Spain and the European Regional Development Fund under the project ENE2015-69937-R (MINECO/FEDER, UE). GSA work is supported by the Ministerio de Economía, Industria y Competitividad of Spain under the Grant RYC-2014-15357. RS was partially supported by the EU projects AWESCO (H2020-ITN-642682) and REACH (H2020-FTIPilot-691173). Publicado

A steady growth of aviation transportation (4-5% per year) causes excessive saturation at numerous major airports. As a result, many flights are delayed. One of the ways to deal with this growing problem is to transfer regional propeller aircraft maintenance to suburban airports. It will require both a modernization of local airports and the design of a new generation of regional aircraft with short takeoff and landing (STOL). The aircraft ability to operate from short runways depends not only on wing unit loading and on high-lift capacities but also it is determined by the control surfaces efficiency. The latter often becomes one of the major reasons for limitation of the amount of lift used in STOL configuration. Thus, the successful application of high-lift devices stipulates the necessity for both the efficiency increasing of existing aircraft control surfaces and the development of some alternative form of lateral control not requiring a significant wingspan proportion. The forms of a lateral control, this article considers, include the interceptor, drooped ailerons, ailerons fitted with mini-flap and one of the alternative forms which uses differential flap section deflection. Several mini-flaps with a various chord are also considered to increase the available rudder yawing moment. The efficiency of the above-mentioned control surfaces has been studied in TsAGI low speed wind tunnel on a model of a twin-engine light aircraft with an enhanced level of lifting capacity on take-off and landing configurations. The tests were conducted at a Reynolds number of 1.0×10 6 and Mach number of 0.15.

PurposeThe purpose of the paper is to analyze the active suppression of the aeroelastic vibrations of ailerons with strongly nonlinear characteristics by neural network/reinforcement learning (NN/RL) control method and comparing it with the classic robust methods of suppression.Design/methodology/approachThe flexible wing and aileron with hysteresis nonlinearity is treated as a plant-controller system and NN/RL and robust controller are used to suppress the nonlinear aeroelastic vibrations of aileron. The simulation approach is used for analyzing the efficiency of both types of methods in suppressing of such vibrations.FindingsThe analysis shows that the NN/RL controller is able to suppress the nonlinear vibrations of aileron much better than linear robust method, although its efficiency depends essentially on the NN topology as well as on the RL strategy.Research limitations/implicationsOnly numerical analysis was carried out; thus, the proposed solution is of theoretical value, and its application to the real suppression of aeroelastic vibrations requires further research.Practical implicationsThe work shows the NN/RL method has a great potential in improving suppression of highly nonlinear aeroelastic vibrations, opposed to the classical robust methods that probably reach their limits in this area.Originality/valueThe work raises the questions of controllability of the highly nonlinear aeroelastic systems by means of classical robust and NN/RL methods of control.

Purpose The separation of energy conversion and propulsor is a promising aspect of hybrid-electric propulsion systems, allowing for increased installation efficiencies and setting the basis for distributed propulsion concepts. University of Stuttgart’s Institute of Aircraft Design has a long experience with electrically powered aircraft, starting with Icaré 2, a solar-powered glider flying, since 1996. Icaré 2 recently has been converted to a three-engine motor glider with two battery-powered wing-tip propellers, in addition to the solar-powered main electric motor. This adds propulsion redundancy and will allow analyzing yaw control concepts with differential thrust and the propeller-vortex interaction at the wing-tip. To ensure airworthiness for this design modification, new ground vibration tests (GVTs) and flutter calculations are required. The purpose of this paper is to lay out the atypical approach to test execution due to peculiarities of the Icaré 2 design such as an asymmetrical aileron control system, the long wing span with low frequencies of the first mode and elevated wing tips bending under gravity and thus affecting the accuracy of the wing torsion frequency measurements. Design/methodology/approach A flutter analysis based on GVT results is performed for the aircraft in basic configuration and with wing tip propulsors in pusher or tractor configuration. Apart from the measured resonant modes, the aircraft rigid body modes and the control surface mechanism modes are taken into consideration. The flutter calculations are made by a high-speed, low-cost software named JG2 based on the strip theory in aerodynamics and the V-g method of flutter problem solution. Findings With the chosen atypical approach to GVT the impact of the suspension on the test results was shown to be minimal. Flutter analysis has proven that the critical flutter speed of Icaré 2 is sufficiently high in all configurations. Practical implications The atypical approach to GVT and subsequent flutter analysis have shown that the effects of wing-tip propulsors on aeroelasticity of the high aspect ratio configuration do not negatively affect flutter characteristics. This analysis can serve as a basis for an application for a permit to fly. Originality/value The presented methodology is valuable for the flutter assessment of aircraft configurations with atypical aeroelastic characteristics.

The purpose of this project is to optimize the design, dimensions along with changing in material of the bracket structure to ensure the structural fatigue life increase along with reduction in stress and displacement of the component. Optimization study of components gives a great opportunity for material saving which leads to save cost and man power. The major bracket design parameters were explained in detail and the bracket configuration has been described. Different types of loads acting on the aircrafts bracket are determined and the moments, displacements, etc., are also determined. The bracket structure was also explained and functions of each component and their arrangement are also studied. The methodology of finite element method and the detailed description about various FEM tools have been studied and implemented in this work. The procedure of finite element method was followed to analyse the model. The analysing part of this project is done using CAE tool package and the results were discussed. In this Project, we designed Aircraft engine bracket using CATIA V5 and carried out linear static analysis using MSC Software. For Pre-Processing used MSC PATRAN and MSC NASTRAN for solving followed by Fatigue calculations.

A closed-loop coupling model at the system-level is developed to analyze the effects of joint clearance on the dynamic responses of electromechanical aileron actuation system. The proposed model considers the coupling effects between the electromechanical actuator (EMA) control performance and dynamic characteristics of linkage mechanism with joint clearance. Besides, the experiments are conducted in a test rig, which verifies the effectiveness of the proposed closed-loop coupling model. The nonlinear contact force model and modified Coulomb friction model are adopted in the joint clearance of the linkage mechanism, and the influences of clearance size on the dynamic behaviors of electromechanical aileron actuation system are studied. The numerical and experimental results indicate that the novel closed-loop coupling model, considering the EMA control performance and dynamics of linkage mechanism with joint clearance at the same time, is an effective model to predict the dynamic characteristics of electromechanical aileron actuation with joint clearance, which provides a practical method to analyze the dynamic performance of electromechanical coupling multibody systems with joint clearance.