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The robot pilot is a new concept of a robot system that pilots a manned aircraft, thereby forming a new type of unmanned aircraft system (UAS) that makes full use of the platform maturity, load capacity, and airworthiness of existing manned aircraft while greatly expanding the operation and application fields of UASs. In this research, the implementation and advantages of the robot pilot concept are discussed in detail, and a helicopter robot pilot is proposed to fly manned helicopters. The robot manipulators are designed according to the handling characteristics of the helicopter-controlling mechanism. Based on a kinematic analysis of the robot manipulators, a direct-driving method is established for the robot flight controller to reduce the time delay and control error of the robot servo process. A supporting ground station is built to realize different flight modes and the functional integration of the robot pilot. Finally, a prototype of the helicopter robot pilot is processed and installed in a helicopter to carry out flight tests. The test results show that the robot pilot can independently fly the helicopter to realize forward flight, backward flight, side flight, and turning flight, which verifies the effectiveness of the helicopter robot pilot.

A three-wing Flapping Wing Rotor Micro Aerial Vehicle (FWR-MAV) which can perform controlled flight is introduced and an experimental study on this vehicle is presented. A mechanically driven flapping rotary mechanism is designed to drive the three flapping wings and generate lift, and control mechanisms are designed to control the pose of the FWR-MAV. A flight control board for attitude control with robust onboard attitude estimation and a control algorithm is also developed to perform stable hovering flight and forward flight. A series of flight tests was conducted, with hovering flight and forward flight tests performed to optimize the control parameters and assess the performance of the FWR-MAV. The hovering flight test shows the ability of the FWR-MAV to counteract the moment generated by rotary motion and maintain the attitude of the FWR-MAV in space; the experiment of forward flight shows that the FWR-MAV can track the desired attitude.

The morphing wing can improve the flight performance during different phases. However, research has been subject to limitations in aerodynamic characteristics of the morphing wing with a flexible leading-edge. The computational fluid dynamic method and dynamic mesh were used to simulate the continuous morphing of the flexible leading-edge. After comparing the steady aerodynamic characteristics of morphing and conventional wings, this study examined the unsteady aerodynamic characteristics of morphing wings with upward and downward deflections of the leading-edge at different frequencies. The numerical results show that for the steady aerodynamic, the leading-edge deflection mainly affects the stall characteristic. The downward deflection of the leading-edge increases the stall angle of attack and nose-down pitching moment. The results are opposite for the upward deflection. For the unsteady aerodynamic, at a small angle of attack, the transient lift coefficient of the upward deflection, growing with the increase of deflection frequency, is larger than that of the static case. The transient lift coefficient of the downward deflection, decreasing with the increase of deflection frequency, is smaller than that of the static case. However, at a large angle of attack, an opposite effect of deflection frequency on the transient lift coefficient was demonstrated. The transient lift coefficient is larger than that of the static case when the leading edge is in the nose-up stage, and lower than that of the static one in the nose-down stage.

Considering the challenges posed by external disturbances on carrier-based aircraft landing control, higher demands are required for the precision and convergence of the carrier landing control system. First, this paper proposes an Adaptive Terminal Sliding Combined Super Twisting Control (ATS-STC) method to address the issues of low precision, slow convergence, and poor disturbance rejection capability resulting from external disturbances, such as carrier air-wake and deck motion. By introducing a nonlinear term into the sliding surface and employing an integral approach, the proposed ATS-STC method can ensure finite-time convergence and mitigate the chattering problem. An adaptive law is also utilized to estimate the external disturbances, thereby enhancing the anti-disturbance performance. Then, the stability and convergence time analysis of the designed controller are conducted. Based on the proposed method, an Automatic Carrier Landing System (ACLS) is developed to perform the carrier landing control task. Furthermore, a multi-dimensional validation is carried out. For the numerical simulation test, the Terminal Sliding Mode Control (TSMC) method and Proportion Integration Differentiation (PID) method are introduced as comparison, the quantitative assessment results show that the tracking error of TSMC and PID can reach 1.5 times and 2 times that of the proposed method. Finally, the Hardware-in-the-Loop (HIL) test and real flight test are conducted. All the experimental results demonstrate that the proposed control method is more effective and precise.

Morphing wings can improve aircraft performance during different flight phases. Recently research has focused on steady aerodynamic characteristics of the morphing wing with a flexible trailing-edge, and the unsteady aerodynamic and stall characteristics in the deflection process of the morphing wing are worthy further investigation. The effects of the angle of attack and deflection rate on aerodynamic characteristics were examined, and based on the aerodynamic characteristics of the morphing wing, a method was developed to delay stall by using the flexible periodic trailing-edge deflection. The numerical results show that the lift coefficients in the deflection process are smaller than those in the static situation at small angles of attack, and that the higher the deflection rate is, the smaller the lift coefficients will be. On the contrary, at large angles of attack, the lift coefficients are higher than those in the static case, and they become larger with the increase of the deflection rate. Further, the periodic deflection of the flexible trailing-edge with a small deflection amplitude and high deflection rate can increase lift coefficients at the critical stall angle.

Climate warming and the increased demand in air travels motivate the aviation industry to urgently produce technological innovations. One of the most promising innovations is based on the smoothly continuous morphing leading-edge concept. This study proposes a two-step process for the design of a morphing leading-edge, including the optimization of the outer variable-thickness composite compliant skin and the optimization of the inner kinematic mechanism. For the compliant skin design, an optimization of the variable thickness composite skin is proposed based on a laminate continuity model, with laminate continuity constraint and other manufacturing constraints. The laminate continuity model utilizes a guiding sequence and a ply-drop sequence to describe the overall stacking sequence of plies in different thickness regions of the complaint skin. For the inner kinematic mechanism design, a coupled four-bar linkage system is proposed and optimized to produce specific trajectories at the actuation points on the stringer hats of the compliant skin, which ensures that the compliant skin can be deflected into the aerodynamically optimal profile. Finally, a morphing leading-edge is manufactured and tested. Experimental results are compared with numerical predictions, confirming the feasibility of the morphing leading-edge concept and the overall proposed design approach.

The arresting process of carrier-based aircraft is widely recognized as a challenging task, characterized by the highest accident rate among all carrier-based aircraft operations. Dynamic simulation plays a crucial role in assessing the intricate responses of the arresting process, favoring the design of carrier-based aircraft. An efficient and accurate rigid-flexible coupling model for analyzing the dynamics response of the arresting process is proposed. By combining the dynamic characteristics of airframe, landing gear, arresting hook and arresting gear system, the rigid-flexible coupling model is established to reflect the relative motion of the coupling parts and arresting load. The dynamics model is verified through simulations of landing gear landing drops and by comparing the arresting simulation results with corresponding data in the US military standard. Additionally, simulations of the arresting process under different off-center distance and aircraft yaw angle are conducted to obtain the dynamic response of the aircraft during the arresting process. The result indicates that the rigid-flexible coupling model proposed is effective for analyzing the arresting dynamics response of carrier-based aircraft. The axial force of the arresting cable on both sides of the hook engagement point, pitch and yaw angle of aircraft are inconsistent under yaw and off-center arresting. The analysis method and obtained results provide valuable references for assessing the dynamic responses of carrier-based aircraft during arresting process and provide valuable in-sights in the design of carrier-based aircraft.

“Ionic wind” generators are used as the main propulsion system in ion propulsion unmanned aerial vehicles (UAVs). Owing to the large size and poor stiffness of the electrode array in the propulsion system, the electrode array is prone to deformation under the flight load. In this work, the thrust characteristics and static aeroelastic properties of "ionic wind" propulsion systems were analysed in detail. The simulation model for an "ionic wind" propulsion system was established by coupling a two-dimensional gas discharge model with a gas dynamics model. The influences of electrode voltage, spacing, size, and shape on the performance of the propulsion system were investigated. The fluid-solid interaction method was used to solve static aeroelastic characteristics under deformation. The aerodynamic and thrust performances of the elastic state and the rigid state were compared. It was found that the operating voltage, the distance between two electrodes, and the emitter radius had greater impacts on the thrust of the propulsion system. The propulsion system had a small contribution to the lift but a large contribution to the drag. In the elastic state, the lift coefficient accounted for 12.2%, and the drag coefficient accounted for 25.8%. Under the action of the downwash airflow from the wing, the propulsion system formed a upward moment around the center of mass, which contributed greatly to the pitching moment derivative of the whole aircraft. In the elastic state, the pitching moment derivative accounted for 29.7%. After elastic deformation, the thrust action point moved upward by 28.7 mm. Hence, the no lift pitching moment is reduced by 0.104 N·m, and the pitching moment coefficient is reduced by 0.014, causing a great impact on the longitudinal trimming of the whole aircraft.

Describing spatial safety status is crucial for high-density air traffic involving multiple unmanned aerial vehicles (UAVs) in a complex environment. A probabilistic approach is proposed to measure safety situation in congested airspace. The occupancy distribution of the airspace is represented with conflict probability between spatial positions and UAV. The concept of a safety envelope related to flight performance and response time is presented first instead of the conventional fixed-size protected zones around aircraft. Consequently, the conflict probability is performance-dependent, and effects of various UAVs on safety can be distinguished. The uncertainty of a UAV future position is explicitly accounted for as Brownian motion. An analytic approximate algorithm for the conflict probability is developed to decrease the computational consumption. The relationship between safety and flight performance are discussed for different response times and prediction intervals. To illustrate the applications of the approach, an experiment of three UAVs in formation flight is performed. In addition, an example of trajectory planning is simulated for one UAV flying over airspace where five UAVs exist. The validation of the approach shows its potential in guaranteeing flight safety in highly dynamic environment.

In this letter, a piezoaeroelastic energy harvester based on an airfoil with double plunge degrees of freedom is proposed to additionally take advantage of the vibrational energy of the airfoil pitch motion. An analytical model of the proposed energy harvesting system is built and compared with an equivalent model using the well-explored pitch-plunge configuration. The dynamic response and average power output of the harvester are numerically studied as the flow velocity exceeds the cut-in speed (flutter speed). It is found that the harvester with double-plunge configuration generates 4%–10% more power with varying flow velocities while reducing 6% of the cut-in speed than its counterpart.

Gliding is an important flight mode for insects because it saves energy during long distance flight without ing flapping. In this study, we investigated the influence of locust wing corrugation on the aerodynamic performance in gliding mode at low Reynolds number. Numerical simulations using two-dimensional Navier-Stokes equations are applied to study the gliding flight, which reveals the interaction between forewing and hindwing. The lift of the corrugated airfoil in a locust wing decreases from the wing root to the tip. Simulation results show that the pressure drags on the forewing and hindwing increase with an increase in wing thickness; while the lift-drag ratio of the airfoil is marginally affected by the corrugation on the airfoil. Geometric parameters analysis of the locust wing is also carried out, which includes the corrugation height, the corrugation placement and the shapes of leading and trailing edges.

An integrated design concept for crashworthy fuselage using sine-wave beam and strut is proposed and investigated. The finite element model of aircraft fuselage is built first. The structures above cabin floor, occupant and seat are simplified as two rigid blocks. The fuselage frame is redesigned, and the sine-wave beam is arranged under the frame. The impact dynamic performance of the aircraft with bottom sine-wave beam structure is studied and compared with that of conventional type. To obtain better crashworthiness performance, different rigidity of strut is combined with the sine-wave beam bottom structure. Numerical simulation result shows that the proposed sine-wave beam bottom structure could not only dissipate more proportion of impact kinetic energy but also reduce the initial peak acceleration. The structure and rigidity of strut have great influence on the crashworthiness performance. To give a better fuselage structure, both of the strut and bottom structure should be properly integrated and designed.

Gliding is an important flight mode for insects because it saves energy during long distance flight without wing flapping. In this study, we investigated the influence of locust wing corrugation on the aerodynamic performance in gliding mode at low Reynolds number. Numerical simulations using two-dimensional Navier-Stokes equations are applied to study the gliding flight, which reveals the interaction between forewing and hindwing. The lift of the corrugated airfoil in a locust wing decreases from the wing root to the tip. Simulation results show that the pressure drags on the forewing and hindwing increase with an increase in wing thickness; while the lift-drag ratio of the airfoil is marginally affected by the corrugation on the airfoil. Geometric parameters analysis of the locust wing is also carried out, which includes the corrugation height, the corrugation placement and the shapes of leading and trailing edges.

Design and implementation of an effective control mechanism is the key to enable the controlled flight of flapping wing rotor micro aerial vehicles (FWR-MAVs). In order to address this open challenge, in this paper, a triple-wing FWR-MAV with its particular control system is proposed. The corresponding experimental study of the vehicle is conducted. As a result, the triple-wing propulsion system demonstrates advantages in both lift generation and flight stability based on the force measurement. The control torques of the proposed FWR-MAV are generated by a particular slip flow rudder mechanism. The effectiveness of such a design combination has been validated experimentally. A series of flight tests are conducted, including sustained hovering and forward-backward maneuvering. To the best of our knowledge, such a design yields the world’s first FWR-MAV capable of sustained controlled flight.

Insect-like flapping flight offers a power-efficient and highly manoeuvrable basis for a micro air vehicle. In this paper, investigation was made into the aerodynamic analysis of a flapping wing applicable to micro aerial vehicles. First, the three dimensional aerodynamic calculation and flow field simulation of a flapping wing were studied. Then, the generation of reduced order models for computing the unsteady and nonlinear aerodynamic loads on the flapping wing from flapping motions is described. The models considered are based on Volterra theory using nonlinear kernels. One of the biggest challenges in creating these reduced order modeling techniques is accurate identification of the Volterra kernels. The validity of models studied was assessed by comparison of the model outputs with unsteady CFD result. It was proved that the percent error between CFD and reduced order model is 8.55.

To improve the survivability during an emergency situation, an algorithm for aircraft forced landing trajectory planning is proposed. The method integrates damaged aircraft modelling and trajectory planning into an optimal control framework. In order to deal with the complex aircraft flight dynamics, a solving strategy based on Gauss pseudospetral method (GPM) is presented. A 3-DOF nonlinear mass-point model taking into account the wind is developed to approximate the aircraft flight dynamics after loss of thrust. The solution minimizes the forced landing duration, with respect to the constraints that translate the changed dynamics, flight envelope limitation and operational safety requirements. The GPM is used to convert the trajectory planning problem to a nonlinear programming problem (NLP), which is solved by sequential quadratic programming algorithm. Simulation results show that the proposed algorithm can generate the minimum-time forced landing trajectory in event of engine-out with high efficiency and precision.

The crashworthiness is an important design factor of civil aircraft related with the safety of occupant during impact accident. It is a highly nonlinear transient dynamic problem and may be greatly influenced by the uncertainty factors. Crashworthiness uncertainty analysis is conducted to investigate the effects of initial conditions, structural dimensions and material properties. Simplified finite element model is built based on the geometrical model and basic physics phenomenon. Box–Behnken sampling and response surface methods are adopted to obtain gradient information. Results show that the proposed methods are effective for crashworthiness uncertainty analysis. Yield stress, frame thickness, impact velocity and angle have great influence on the failure behavior, and yield stress and frame thickness dominate the uncertainty of internal energy. Failure strain and tangent modulus have the smallest influence on the initial peak acceleration, and gradients of mean acceleration increase because the appearance of material plastic deformation and element failure.

An investigation was made into the nonlinear aeroelastic behavior of a composite wing with morphing trailing edge actuated by curved beams. Based on the design concept, an experimental wing model was built and a finite element model was created using MSC Patran/Nastran software. Impact test of the experimental model was carried out and the test data were used to validate the finite element models. In order to investigate the effect of freeplay nonlinearity between discs attached to the curved beams and wing skins, a nonlinear aeroelastic equation in modal coordinate was created. Generalized structural mass, damping, and stiffness matrixes were printed with DMAP language in Nastran. Doublet lattice method and Roger’s approximation were used to calculate unsteady aerodynamic influence coefficients matrix. In the analysis, the effect of morphing stiffness on critical flutter speed was studied. In addition, Hopf bifurcation and limit cycle oscillation were detected for the aeroelastic system with freeplay. The results show that the freeplay nonlinearity may reduce convergence speed and accelerate divergence of aeroelastic responses.

Abstract: In order to improve airfoil performance under different flight conditions and to make the performance insensitive to off-design condition at the same time, a multi-objective optimization approach considering robust design has been developed and applied to airfoil design. Non-uniform rational B-spline (NURBS) representation is adopted in airfoil design process, control points and related weights around airfoil are used as design variables. Two airfoil representation cases show that the NURBS method can get airfoil geometry with max geometry error less than 0.0019. By using six-sigma robust approach in multi-objective airfoil design, each sub-objective function of the problem has robustness property. By adopting multi-objective genetic algorithm that is based on non-dominated sorting, a set of non-dominated airfoil solutions with robustness can be obtained in the design. The optimum robust airfoil can be traded off and selected in these non-dominated solutions by design tendency. By using the above methods, a multi-objective robust optimization was conducted for NASA SC0712 airfoil. After performing robust airfoil optimization, the mean value of drag coefficient at Ma0.7–0.8 and the mean value of lift coefficient at post stall regime (Ma0.3) have been improved by 12.2% and 25.4%. By comparing the aerodynamic force coefficients of optimization result, it shows that: different from single robust airfoil design which just improves the property of drag divergence at Ma0.7–0.8, multi-objective robust design can improve both the drag divergence property at Ma0.7–0.8 and stall property at low speed. The design cases show that the multi-objective robust design method makes the airfoil performance robust under different off-design conditions.

Gust response of aircraft can become a potential energy source thanks to energy harvesting (EH) methods, one of which can transform mechanical energy to electrical energy, applying piezoelectric ceramic transducers (PZTs). Harvested electrical energy needs to be evaluated for reuse, but current beam-model-based evaluation methods of EH performance for a plate model are insufficient because the plate is two-dimensional. This paper proposes two types of EH evaluation methods to analyze the gust exciting low-aspect-ratio plate wing model. One method focuses on the ratio of electrical energy to kinetic energy, and the other reflects energy output density per unit weight. These two methods can reveal the higher energy harvesting efficiency positions when utilizing PZTs on the flight EH system. Plate model, three-dimensional doublet lattice unsteady aeroelastic method and piezoelectric equation are used to build piezoaeroelastic wing model, and 1-cosine discrete gust load is the base excitation. A time-domain aero-electro-elastic state-space equation of the low-aspect-ratio piezoaeroelastic plate wing model is established, and the time-history analysis is used to solve the EH output. Results show that EH outputs are influenced by various parameters including load resistance, PZT thickness, gust intensity and wavelength, and PZT location variation. The optimal values of the proposed EH efficiency evaluation methods are found and explained.