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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.

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.

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 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.

The purpose of this paper is to develop a six degree-of-freedom flight simulation while considering the flexibility of an aircraft with high aspect ratio wings using by new quasi-steady temporal coupling scheme named ‘snapshot method’. The simulation was based on the quasi-steady temporal coupling scheme for the flow field, elastic structure, and rigid body motion problem. A full three-dimensional finite-element model of unmanned aerial vehicles was used. The object aircraft has main wings with an aspect ratio over 20 and a flight for long endurance in high altitude. To consider the flexibility of these aircraft, the present snapshot method was developed that combines ‘aerodynamic—structural dynamics—flight dynamics’ to analyze dynamic response. By applying this method, MATLAB/SIMULINK, which analyzes the rigid body motion, and MSC FlightLoads, which is responsible for aeroelastic trim analysis, was tightly linked into the present simulation framework. Using the present simulation, aircraft response under various maneuver conditions was simulated. First, the trim analysis for the level cruise was performed. Trim parameters, stability derivative coefficients, and the moment of inertia were also determined. Based on these results, the flight dynamics of the aircraft was simulated according to the operation of the rudder and the aileron control surface. The effect of the rigid and flexible body assumptions was also confirmed. In addition, these results were compared with other existing simulation results based on the multibody dynamics under the same conditions. To predict the response of the aircraft under the gust, the simulation was performed by applying the two-dimensional 1-cosine discrete gust profile. The important aspects, such as numerical prediction methods during stages of the highly flexible aircraft design process for a countermeasure to various challenging problems, were shown.

This paper presents a model-based approach to design aileron endurance test rig (ETR). ETR is a dynamic load simulator which simulates aerodynamic load on-ground for verifying and validating the design, performance and stability of aileron actuator. Aileron actuator is a servo-controlled linear hydraulic actuator used to control the movement of ailerons in aircraft. Aileron is one of the primary flight control surfaces which controls roll of the aircraft. In ETR, Aileron actuator acts as unit under test (UUT) while a double-acting linear hydraulic actuator produces a dynamic load with the help of high pressure fluid source and electro-hydraulic servo valve (EHSV). The design of the test rig depends on load and velocity requirements which vary widely over the whole flight envelop and depends on deflection of surface, angle of attack, aircraft speed and altitude. One of the critical factor in designing ETR is to accurately model the interaction between the UUT and load system. This paper presents a simple yet powerful approach of free body diagram to account the power flow between the two systems. Model-based approach allowed to simulate the complete test rig behavior identifying the values of the critical parameters prior to building it. A high fidelity, non-linear mathematical model of aileron ETR is developed, simulated and verified. An appropriate load actuator and its electro-hydraulic servo valve are chosen to meet load and velocity requirements. The minimum rig structure stiffness is determined to ensure the stability of the load control system. A velocity feed-forward-based load controller along with proportional-integral control is implemented and tuned to meet the load control performance satisfactorily. Finally, the developed model is validated against the experimental data from actual test rig.

We propose a mathematical model of a system of wing profiles of tandem type with elastic ailerons in subsonic flow of an ideal gas (liquid). The aerohydroelasticity problem is reduced to the study of a system of integro-differential equations. The solution is constructed by the Bubnov–Galerkin method. Numerical experiments are presented in the case of two profiles with elastic ailerons. We establish the dynamic stability of the aileron and obtain the stability conditions if only one profile has an elastic aileron.

This paper focuses on the computational study of nonlinear effects of unsteady aerodynamics for non-classical aileron buzz. It aims at a comprehensive investigation of aileron buzz phenomenon under varying flow parameters utilizing the classical approach of describing function technique with multiple input variables. Limit cycle oscillatory behavior of an asymmetrical airfoil section has been studied initially using an expensive high-fidelity model which uses coupled CFD/CSD time marching approach. Using this high fidelity model, a multiple input aerodynamic describing function was developed that can effectively represent lumped nonlinearities of aerodynamic subsystem. It has been demonstrated that limit cycles can be closely identified through aerodynamic describing function representation. Traditional LCO analysis of sensitivity with respect to hinge moment of inertia is performed. Further, outburst points of limit cycle amplitude for varying Mach No. and angle of attacks are investigated. By using developed aerodynamic describing function, it has been demonstrated that limit cycle analysis methods may still be used while including the effects of aerodynamic nonlinearities. Results from both methods, direct time marching and multiple input describing function, have been broadly compared for validation.

This paper concerns design, development, and testing of a morphing wing for an unmanned aerial vehicle with the aim of improving the vehicle’s performance over its flight envelope. For this purpose, a variable-span morphing wing (VSMW) model that allows changes in wing span, was developed and exposed to wind tunnel testing. The morphing wing model was tested in a low-speed wind tunnel facility at a variety of Reynolds numbers from 250 × 103 to 620 × 103 under steady wind speeds (i.e., 35, 60, and 80 m/s). Experimental wind tunnel results were then compared against base wing test specimen to evaluate the performance of the morphing wing. The experiment demonstrated an improvement in the aerodynamic efficiency of the VSMW in the form of a reduction in the drag and a subsequent increase of 5 and 17 percent in flight range and endurance, respectively. It was also found that the maximum aerodynamic efficiency of the wing model occurs at 4 degree angle of attack, which was 17.4 for the morphing wing with 100 percent increase in the wing span. This research also shows how the VSMW can be used as an alternative and effective method of roll control instead of the conventional method using aileron control surfaces.

Aircraft are inevitably affected by gust during flight, which disturbs the regular operations of pilots and worsens the ride quality. In more grievous cases, flight mission cannot be completed and the flight safety may be disserved. In order to improve the ride quality and the fatigue life of the plane structure under the affect of gust, it is necessary to explore and validate the gust alleviation schemes. Through the low-speed wind tunnel test, the gust alleviation active control technology applied to elastic aircraft is studied. For a large-type passenger plane configuration with high aspect ratio wing, a test system was designed and three gust alleviation control schemes with PID controllers were proposed. Finally the gust alleviation control low-speed wind tunnel test was carried out in the FD-09 wind tunnel. Test results showed that at certain speed and gust frequency, all of the 3 control schemes can alleviate the acceleration at fuselage and wing-tip to a certain extent, as well as the bending moment of wing-root. The gust alleviation control scheme, which uses aileron, elevator and canard as control surfaces synthetically, gives the most satisfying gust alleviation effect.

High temperature superconducting (HTS) magnetic levitation (maglev) vehicle is a fascinating transportation tool due to its unique inherent stability. In order to improve performance of present HTS maglev vehicle system, a novel HTS maglev vehicle (aero-vehicle) utilizing aerodynamic designs from modern aircraft is proposed, which is driven by an engine positioned on the top of the car body. The aerodynamic designs of the aero-vehicle include the wings and the empennage, and these new designs are used to adjust and control the pitch, roll, and yaw of the car body. To analyze the feasibility of the aero-vehicle, its characteristics are investigated in the paper. Depending on ailerons of wings, elevator and rudder of the empennage, the aero-vehicle is capable to advance safety and ride comfort of present HTS maglev vehicle.

The aeroelastic responses of a high-aspect-ratio wing are investigated based on nonlinear experimental aerodynamic forces. The influences of nonlinear experimental aerodynamic forces and dynamic pressures on the wing loads are studied in the longitudinal and lateral maneuver states. The flight loads of the wing fixed at the root are calculated at different angles of attack. The aileron efficiency with respect to the dynamic pressures and aileron deflections are also studied. The results indicate that the flight loads of the wings vary nonlinearly with the angle of attack and dynamic pressure. Due to the high-lift aerofoil, elastic components are a large portion of the wing loads, especially for small angles of attack and high dynamic pressure conditions. The aileron efficiency is significantly affected by aileron deflections, dynamic pressures and angles of attack when the nonlinear experimental aerodynamic forces are used for calculation. In states with high dynamic pressures and large aileron deflections, aileron reversal can occur. The aileron deflection and angle of attack have a nonlinear effect on the aileron efficiency. An efficient method for analyzing the flight loads and structural design of high-aspect-ratio wings is derived in this study, and the analysis can provide insight into the distribution of flight loads for high-aspect-ratio wings.

Time-domain approaches are presented for analysis of the dynamic response of aeroservoelastic systems to atmospheric gust excitations. The continuous and discrete gust inputs are defined in the time domain. The time-domain approach to continuous gust response uses a state-space formulation that requires the frequency-dependent aerodynamic coefficients to be approximated with the rational function of a Laplace variable. A hybrid method which combines the Fourier transform and time-domain approaches is used to calculate discrete gust response. The purpose of this approach is to obtain a time-domain state-space model without using rational function approximation of the gust columns. Three control schemes are designed for gust alleviation on an elastic aircraft, and three control surfaces are used: aileron, elevator and spoiler. The signals from the rate of pitch angle gyroscope or angle of attack sensor are sent to the elevator while the signals from accelerometers at the wing tip and center of gravity of the aircraft are sent to the aileron and spoiler, respectively. All the control laws are based on classical control theory. The results show that acceleration at the center of gravity of the aircraft and bending-moment at the wing-root section are mainly excited by rigid modes of the aircraft and the accelerations at the wing-tip are mainly excited by elastic modes of the aircraft. All the three control schemes can be used to alleviate the wing-root moments and the accelerations. The gust response can be alleviated using control scheme 3, in which the spoiler is used as a control surface, but the effects are not as good as those of control schemes 1 and 2.

A modeling and control approach for an advanced configured large civil aircraft with aeroservoelasticity via the LQG method and control allocation is presented. Mathematical models and implementation issues for the multi-input/multi-output (MIMO) aeroservoelastic system simulation developed for a flexible wing with multi control surfaces are described. A fuzzy logic based optimization approach is employed to solve the constrained control allocation problem via intelligently adjusting the components of output vector and find a proper vector in the attainable moment set (AMS) autonomously. The basic idea is to minimize the L2 norm of error between the desired moment and achievable moment using the designing freedom provided by redundantly allocated actuators and control surfaces. Considering the constraints of control surfaces, in order to obtain acceptable performance of aircraft such as stability and maneuverability, the fuzzy weights are updated by the learning algorithm, which makes the closed-loop system self-adaptation. Finally, an application example of flight control designing for the advanced civil aircraft is discussed as a demonstration. The studies we have performed showed that the advanced configured large civil aircraft has good performance with the proper designed control law designed via the proposed approach. The gust alleviation and flutter suppression are applied with the synergetic effects of elevator, ailerons, equivalent rudders and flaps. The results show good closed loop performance and meet the requirement of constraint of control surfaces.

This paper details a quantitative 3D investigation using LDA into the interaction aerodynamics on a sub-scale open wheel race car inverted front wing and wheel. Of primary importance to this study was the influence of changing wing angle of attack and span on the resulting near-field and far-field flow characteristics. Results obtained showed that both variables do have a significant influence on the resultant flow-field, particularly on wing vortex and wheel wake development and propagation.

The most fundamental concept in designing multi-lane smart electromechanical actuation systems, besides meeting performance requirements, is the realization of high integrity. The main aim of this paper is to discuss fundamental consolidation designs and monitoring schemes in different architectures and to address threshold settings methodologies, inherent randomness, lane equalization, and control strategy. The analysis is based on a 4-lane actuation system capable of driving aerodynamic and inertial loads (with 2 lanes failed) of an aileron control surface similar to that of the Sea Harrier.

The paper is concerned with mechanical phenomena accompanying ground run of aircraft with near-critical slip angles. Such phenomena (reduced effectiveness of directional control, self-excited frictional vibrations of a landing gear, internal resonance of an aileron, transverse oscillations of the legs of a multileg landing gear) are due to the nonlinear dependence of the side force on the slip angles

The results of cyclic tests of structural elements (SEs) with a stringer, which are made of the composite material (CM) Organit-10T and are parts of an aileron from an e-80Ee aircraft, are considered. As SEs are broken, information is obtained simultaneously from acoustic-emission (AE) and tensometric equipment and measurements of the fatigue crack’s opening are carried out. An insignificant increase (by 0.25 µm) of the crack opening generates a considerable flow of AE signals. These signals are processed using cluster analysis, which makes it possible to separate the signals emitted by a propagating fatigue crack in the stringer region from the signals arriving from the region of the fracturing gripping jaw.

A low speed wind tunnel test was conducted for full-scale model of an unmanned aerial vehicle (UAV) in Korea Aerospace Research Institute (KARI) Low Speed Wind Tunnel (LSWT). The purpose of the presented paper is to illustrate the general aerodynamic and performance characteristics of the UAV that was designed and fabricated in KARI. Since the testing conditions were represented minor portions of the load-range of the external balance system, the repeatability tests were performed at various model configurations to confirm the reliability of measurements. Variations of drag-polar by adding model components such as tails, landing gear and test boom are shown, and longitudinal and lateral aerodynamic characteristics after changing control surfaces such as aileron, flap, elevator and rudder are also presented. To explore aerodynamic characteristics of an UAV with model components build-up and control surface deflections, lift curve slope, pitching moment variation with lift coefficients and drag-polar are examined. The discussed results might be useful to understand the general aerodynamic characteristics and drag pattern for the given UAV configuration.

This paper is concerned with designing an optimum composite wing structure with enhanced roll maneuver capability at high dynamic pressures using a control system to retwist and recamber the wing. Antisymmetric twist and camber distribution on a composite flexible wing, weight optimized with constraints on strength and aileron efficiencies also satisfying flutter requirements, were determined to counter act the detrimental twisting moment of the aileron at high dynamic pressure. The retwisting and recambering of the wing is achieved by providing control forces obtained from a technique referred to as “Fictitious Control Surfaces” (FCS). Two different optimally designed wings with unacceptable roll performance were selected for roll performance improvement investigation. The control energy necessary to retwist the wing to improve the unacceptable roll performance for subsonic design condition was calculated to provide the structural designer an alternative, to the traditional increase of stiffness to satisfy roll performance.

The Euler and Navier-Stokes equation for describing complex flow phenomena around a wing with aileron is effectively solved using the combination of domain decomposition technique and patched grids. An internal boundary condition satisfied the conservation of flux is given. In computation, Van Leer's scheme and central difference scheme are used for inviscid and viscous flux terms respectively and a new limiter function is constructed to preserve the scheme having the TVD character. Numerical results show that present method is an effective one for describing the flow around a 3-D wing with control surfaces.

Modern flying vehicles (FV) have broad ranges of variation of the exploitational characteristics, possess elevated structure flexibility, use numerous load modifications that result in a substantial variation in the elastic mass and aerodynamic parameters in flight and an increase in the possibility of the appearance of dangerous aeroelastic vibrations. One of the new means for eliminating aeroelastic vibrations is the application of the socalled active control systems. Such a system (Fig. i) consists of sensors i that yield information about the FV vibrations, an on-board digital computer 2 in which the aileron 3 control algorithm is set up using the drivers 4. An indubitable advantage of active control systems is the possibility of adapting the control law to flight condition changes and assuring its optimality during changes in the FV characteristics.

For the limiting case when the ratio of the characteristic wing dimension to the characteristic aileron dimension tends to infinity, an exact general solution to the problem of linear wing theory expressed in quadratures is obtained. Examples are given of calculations of the distribution of the pressure and the hinge moments for rectangular and trapeziform ailerons.

When an aircraft has developed a spin, the pilot is usually concerned not only with recovering from the spin but doing so with minimum loss of height. This paper is mainly concerned with the mathematical formulation of this problem, including the simplifications that can reasonably be made. The Pontryagin minimum principle is used to obtain necessary conditions for a solution to the problem. The amount of computation required may be prohibitive; on the other hand, if certain simplifying assumptions may be regarded as valid, it is indicated how a solution may be obtained. The theory is restricted to the case in which three controls are actively used: aileron, elevator and rudder. However, further controls may easily be included, with only a slight increase in complexity.