Variable Sweep Wing Research Articles

Aerodynamic Performance Analysis of a Variable Sweep Wing for Commercial Aircraft Applications

This study presents a detailed study on wing and its configurations and the morphing techniques for the wing. The morphing methods of the wing such as variable chord, variable span variable cambers have been studied in detail. In this study in detail about the effects of morphable sweep wing, the commercial aircraft wing has been designed and it‘s been modelled using the solid works software. To study the aerodynamic performance the wing, the wing has been analysed in ANSYS Fluent software and the results are interpreted in detail to analyze the effect of wing and its shapes. From the results it‘s been clear that at low speed (Mach=0.8) straight wing has high L/D ratio and at the sonic speed (Mach=1) sweep wing has higher L/D ratio and in Supersonic Speed (Mach=1.2) delta wing tends to have higher L/D ratio. Based on these results the wing can be morphed to the configurations to obtain a better performance in each flight regime. Based on these morphing, aircraft performance can be improved in all flight regimes.

Simulation analysis the aerodynamic characteristics of variable sweep wing missile

When the sweep of the variable sweep folding wing missile changes, the lift-drag ratio is improved so that the missile is able to adapt to subsonic flight, transonic flight and supersonic flight. This paper designs different 3D models of variable sweep wing missile aerodynamic shape and completes the aerodynamic characteristics analysis of the missile by CFD according to the Basic Aerodynamics, Missile General Design Principle, Flight Dynamics of Missile. The simulation analysis of the missile in different flight conditions with different sweeps is carried out, and the wing sweeps applicable to different flight conditions are summarized.

Switching control of morphing aircraft based on Q-learning

This paper investigates a switching control strategy for the altitude motion of a morphing aircraft with variable sweep wings based on Q-learning. The morphing process is regarded as a function of the system states and a related altitude motion model is established. Then, the designed controller is divided into the outer part and inner part, where the outer part is devised by a combination of the back-stepping method and command filter technique so that the ‘explosion of complexity’ problem is eliminated. Moreover, the integrator structure of the altitude motion model is exploited to simplify the back-stepping design, and disturbance observers inspired from the idea of extended state observer are devised to obtain estimations of the system disturbances. The control input switches from the outer part to the inner part when the altitude tracking error converges to a small value and linear approximation of the altitude motion model is applied. The inner part is generated by the Q-learning algorithm which learns the optimal command in the presence of unknown system matrices and disturbances. It is proved rigorously that all signals of the closed-loop system stay bounded by the developed control method and controller switching occurs only once. Finally, comparative simulations are conducted to validate improved control performance of the proposed scheme.

Morphing aircraft control based on switched nonlinear systems and adaptive dynamic programming

This paper investigates the control problem of a morphing aircraft with variable sweep wings based on switched nonlinear systems and adaptive dynamic programming (ADP). The longitudinal altitude motion of the morphing aircraft is first modeled as switched nonlinear systems in lower triangular form. Then, the designed controller is comprised of the basic part and supplementary part. For the basic part, the backstepping technique is applied and a modified dynamic surface is introduced to overcome the ‘explosion of complexity’ problem. Disturbance observers inspired from the idea of extended state observer are designed to obtain the estimations of the internal uncertainties and external disturbances. The common virtual control laws of the backstepping method are developed by the disturbance observers and radial basis function neural networks. On the other hand, for the supplementary part, an ADP approach with the name of action-dependent heuristic dynamic programming is used to further decrease the altitude tracking error, which generates an additional control input by observing the differences between the actual and desired values in the backstepping design. Finally, comparative simulations are conducted to demonstrate the improved control performance of the proposed approach.

Prescribed performance control of morphing aircraft based on switched nonlinear systems and reinforcement learning

The prescribed performance control of a morphing aircraft with variable sweep wings is investigated based on switched nonlinear systems and reinforcement learning. Switched nonlinear systems in lower triangular form are first adopted to describe the longitudinal altitude motion, and an error transformation is applied to handle the prescribed performance bound. Then, the designed controller is divided into the basic part and supplementary part. For the basic part, the backstepping method with involvement of the modified dynamic surface control technique is utilized to avoid the “explosion of complexity” problem. Improved disturbance observers inspired from the idea of extended state observer are then designed to estimate the disturbances and combined with radial basis function neural networks to develop the common virtual control laws. Moreover, by using the error variables defined in the backstepping design, the reinforcement learning–based supplementary part controller is devised with the critic-action neural network structure, which can adjust its parameters online and further decrease the altitude tracking error. It is proved that all signals of the closed-loop system are uniformly ultimately bounded, and the prescribed performance bound for convergence of the altitude tracking error can be satisfied. Finally, comparative simulations demonstrate the effectiveness of the proposed control approach.

Analysis of Morphing Modes of Hypersonic Morphing Aircraft and Multiobjective Trajectory Optimization

Aiming at improving the range of wing-body combination aircraft at hypersonic flow conditions and exploring the application of morphing technology in hypersonic aircraft, morphing aircrafts with different morphing modes have been proposed. The aerodynamic characteristics and wing efficiency of morphing aircraft with different morphing modes have been studied for a further application. In order to verify the significance of morphing technology on the trajectory, the trajectory of the glide phase has been optimized through multiobjective optimization method. Through a 3 degree-of-freedom dynamic model and a heat flux model, the range of glide trajectory and the total heat of the leading edge of the wing are calculated and selected as optimization objectives in the multiobjective optimization problem. The optimization variables include the Mach numbers when the aircraft is morphing (morphing timing) and the angle of attack of different phases. Based on the multiobjective evolutionary algorithm based on decomposition, the multiobjective trajectory optimization problem is solved and a uniform Pareto Front is obtained, and through the analysis of typical solutions, it can be seen that a compromise is made to balance the two objectives. Through the comparison of morphing and non-morphing aircraft, morphing aircraft can fly further with a smaller total heat of the leading edge of the wing. Also, it seems that the variable sweep wing morphing mode has a better overall performance. The result in this paper is a verification of the application prospects of the morphing technology under hypersonic environment, which will provide a reference for further application of morphing technology in hypersonic aircraft.

Learning to Perform a Perched Landing on the Ground Using Deep Reinforcement Learning

A UAV with a variable sweep wing has the potential to perform a perched landing on the ground by achieving high pitch rates to take advantage of dynamic stall. This study focuses on the generation and evaluation of a trajectory to perform a perched landing on the ground using a non-linear constraint optimiser (Interior Point OPTimizer) and a Deep Q-Network (DQN). The trajectory is generated using a numerical model that characterises the dynamics of a UAV with a variable sweep wing which was developed through wind tunnel testing. The trajectories generated by a DQN have been compared with those produced by non-linear constraint optimisation in simulation and flown on the UAV to evaluate performance. The results show that a DQN generates trajectories with a lower cost function and have the potential to generate trajectories from a range of starting conditions (on average generating a trajectory takes 174 milliseconds). The trajectories generated performed a rapid pitch up before the landing site is reached, to reduce the airspeed (on average less than 0.5m/s just above the landing site) without generating an increase in altitude, and then the nose dropped just before hitting the ground to allow the aircraft to be recovered without damaging the tail. The trajectories generated by a DQN produced a final airspeed (when it hit the ground) of 3.25m/s (with a standard deviation of 0.97m/s) in the downward direction, which would allow the aircraft to be safely recovered and significantly less than a traditional landing (∼ 10m/s).

Perched landing manoeuvres with a variable sweep wing UAV

A variable sweep wing UAV is developed utilising off the shelf components with a custom mechanism for the wing box. The movement of the wing sweep in flight enables large pitching moments suitable for performing perching manoeuvres. Wind tunnel data is presented that confirms the favourable characteristics expected from sweeping the wing and achieving high pitch rates. Whilst only small sweep changes are required during flight, the design allows up to 30∘ forward sweep for significant pitching moments during the flare.A new collection of controllers is developed based on observations from similar landing techniques performed by birds and hang-gliders onto flat ground. The three-stage landing process takes the aircraft along an approach path, through a roundout procedure during which airspeed decays and concludes with rapid pitch up. Flight test results are presented during which it is found that the airspeed can be reduced to, on average, under 3 m/s in the final moments before landing – well below the stall speed of 9 m/s.

Switching LPV control design with MDADT and its application to a morphing aircraft

In flight control of a morphing aircraft, the design objective and the dynamics may be different in its various configurations. To accommodate different performance goals in different sweep wing configurations, a novel switching strategy, mode dependent average dwell time (MDADT), is adopted to investigate the flight control of a morphing aircraft in its morphing phase. The switching signal used in this note is more general than the average dwell time (ADT), in which each mode has its own ADT. Under some simplified assumptions the control synthesis condition is formulated as a linear matrix optimization problem and a set of mode-dependent dynamic state feedback controllers are designed. Afterwards the proposed approach is applied to a morphing aircraft with a variable sweep wing to demonstrate its validity.

The Definition of the Load Spectrum for SU-22 Fighter-Bomber Full Scale Fatigue Test

Abstract The Su-22 fighter-bomber is a military aircraft used in the Polish Air Force since the mid 1980’s. By the decision of the Polish Ministry of Defense the predicted service life for this type of aircraft will be extended to 3200 flight hours. Due to the fact that some aircraft were nearing the end of the service life guaranteed by the manufacturer, the actual service life, determined based on the flight profile in the Polish Air Force, had to be validated. Consequently, the Full Scale Fatigue Test (FSFT) had to be carried out in order to verify that the required service life was attainable. This article describes the process of preparation of the load spectra used in the Su-22 FSFT. Due to the fact that the Su-22 has a variable sweep wing the whole test was divided into three Stages (landing, flight and flap loads) carried out at different wing sweep angles (30°/45°/30°). The spectra were developed using the historical data gathered from Flight Data Recorders (FDR), strain signals acquired during the Operational Load Monitoring program (OLM) and aerodynamic calculations.

Numerical Simulation on the Radar Cross Section of Variable-Sweep Wing Aircraft

The influence of variable-sweep wing on the aircraft’s radar cross section (RCS) characteristics has been studied to reduce the aircraft’s RCS as well as its detection probability by the hostile radar. With the help of CATIA, a 3-D digital model of the variable-sweep wing aircraft is built to generate a series of digital grids. Using MATLAB, a numerical simulation on the RCS of variable-sweep wing aircraft is conducted based on physical optics (PO) method and equivalent currents method (ECM). The results of mathematical statistics and comparative analysis show that: (1) the RCS peak value in the head direction of the aircraft decreases non-linearly with the sweep angle of the wing’s leading edge; (2) the azimuth angle corresponding to one of the peak values of the aircraft’s RCS is equal to the leading edge’s sweep angle; (3) when the leading edge’s sweep angle is 33°, the arithmetic average value of the RCS values in the head direction of the aircraft is 0.644% of the average value when the sweep angle is 0°; (4) the larger the sweep angle is, the lower the probability that the aircraft is detected.

Multi-body dynamic modelling and flight control for an asymmetric variable sweep morphing UAV

AbstractIn this paper, the multi-body dynamic model of an asymmetric variable sweep wing morphing UAV is built based on Kane’s method. This model describes the UAV’s transient behaviour during morphing process and the dynamic characteristic of the variable sweep wings. An integrated design of trajectory tracking control via constrained backstepping method is presented then. The idea of aircraft roll control through asymmetric wing sweep angle changes rather than traditional aileron is explored and used in the fight control design. The control of variable sweep wings is designed as well based on the presented dynamic model. Command filters are used in the backstepping design procedure to accommodate magnitude, rate and bandwidth constraints on virtual states and actuator signals. Stability of the closed-loop system can be proved in the sense of Lyapunov. Simulation of tracking a desired trajectory which contains two manoeuvres demonstrates the feasibility of the proposed protocol and the morphing wing roll controller.

Pitching Dynamic Response of Variable Sweep Wing Aircraft

The dynamic response of Variable Sweep Wing Aircraft (VSWA) with the wing sweeping is presented. The center of gravity (cg) of the aircraft, location of each wing partition , and moment of inertia alter significantly due to the wing morphing, resulting in considerably change of the dynamics of the aircraft. The extended equations of motion (EOMs) suitable for morphing wing aircraft are derived. Compared with the traditional EOMs, there are 4 additional forces and moments exhibiting in the extended EOMs due to the wing morphing. The results show that the additional forces and moments can affect the flight control considerably.

Impact blasting with glass beads

A variable sweep wing UAV is developed utilising off the shelf components with a custom mechanism for the wing box. The movement of the wing sweep in flight enables large pitching moments suitable for performing perching manoeuvres. Wind tunnel data is presented that confirms the favourable characteristics expected from sweeping the wing and achieving high pitch rates. Whilst only small sweep changes are required during flight, the design allows up to 30∘ forward sweep for significant pitching moments during the flare.A new collection of controllers is developed based on observations from similar landing techniques performed by birds and hang-gliders onto flat ground. The three-stage landing process takes the aircraft along an approach path, through a roundout procedure during which airspeed decays and concludes with rapid pitch up. Flight test results are presented during which it is found that the airspeed can be reduced to, on average, under 3 m/s in the final moments before landing – well below the stall speed of 9 m/s.

Static stability and control characteristics of scissor wing configurations

A scissor wing geometry is introduced as an alternative to variable sweep and oblique wing designs. It is shown that this configuration offers certain enhancements to the stability and control of the aircraft in addition to aerodynamic advantages associated with sweep. It is shown that a scissor wing configuration can maintain a constant static margin throughout its flight Mach numbers. The dependence of the motion of the aircraft neutral point on the sweep angle is shown to be a function of the chord and span ratios. The control authority is studied and shown not to suffer with changes in the sweep angle. It is also suggested that with the use of wing mounted elevons, additional pitch and attitude control can be obtained over a range of sweep angles.

Transonic unsteady aerodynamic and aeroelastic calculations about airfoils and wings

Some of the research in the area of computational unsteady transonic flow about airfoils and wings, including aeroelastic effects, is surveyed. Results are presented for the unsteady, transonic, small-disturbance potential equation. Several calculations are shown of unsteady transonic flow over airfoils and wings, including cases that employ viscous modeling and cases that compare with experimental data. First an early result of transonic flow with moving shock waves over an airfoil is described. Next a calculation is compared with an experiment of flow over an NACA 0012 airfoil, for which there is a large shock movement. For flow over wings, first calculations are shown of flow over a transport wing, including comparisons with experiment of aeroelastic results. Next calculations of flow over a low aspect ratio wing are compared with experimental results. Finally, aerodynamic calculations of flow over a variable-sweep wing are compared with experiment, and the effect of varying the sweep on the aeroelastic damping is studied.

Vortex-induced bending oscillation of a swept wing

An analysis of existing experimental results shows that the observed bending oscillations of a highly swept wing cannot have been caused by the induced effects of a single leading-edge vortex, but rather by the interaction between two leading-edge vortices, one generated by a thick inner wing or wing/body glove and the other by the outboard, thin, variable-sweep wing. This inner/outer wing vortex interaction is similar to that occurring on a double-delta planform wing. The cause of the negative aerodynamic damping from the vortex interaction is the sensitivity to the angle of attack and its time derivative of the spanwise location of a leadingedge vortex.

Application of energy saving concepts of fighter/attack design

A study of 20 energy saving concepts as applied to an advanced fighter/attack aircraft intended for an initial operational capability (IOC) data of 1995 has been carried out. The results show that the use of surface launched air targeted missiles, advanced engines, conformal external fuel tanks, variable sweep wings, advanced airfoils, relaxed static margin, and the use of advanced structural materials can reduce significantly the fuel consumption of such an aircraft at a modest increase in life cycle costs.

Arrow Wings for Supersonic Cruise Aircraft

The arrow wing planform has a far greater range potential than the delta wing planform for utilization on a commercial supersonic cruise vehicle. A supersonic cruise vehicle concept must be configured to favor cruise efficiency. The arrow wing planform cruise lift-drag ratio is approximately one unit higher than the delta wing. The small weight advantage of the delta wing cannot offset the large cruise efficiency advantage of the arrow wing. The low-speed aerodynamic characteristics of the arrow wing are acceptable and can be further improved by continued research and development. For this, additional analytical tools need to be developed and wind tunnel tests conducted to provide verifications and empirical adjustments to the analytical tools. The major emphasis for arrow wing development should be in the low-speed regime. The design challenge is to seek out design features and refinements that remove any deficiencies of the arrow wing, while not sacrificing the superior cruise efficiency. ING planform selection is a classic problem for commercial supersonic cruise vehicles (SCV's). This fact is highlighted by the various proposed and actual supersonic aircraft that were studied during the USA FAA/SST program, the British/French Concorde program, and the Russian TU 144 program of the 1960's. The NASA SCAT series included arrow, delta, and variable-sweep concepts. Lockheed, Boeing, Douglas, and North American addressed many of the same concepts and modifications thereof, as well as their own concepts. The final Lockheed and Boeing configurations were fixed-wing delta concepts. Both the Concorde and the TU 144 configurations are fixed-wing delta (ogive) concepts. The basic conclusion to be drawn from these SST planform selections is that, based on the available technology, design philosophy, and design requirements of the 1960's, a fixed delta-type wing planform was the unanimous choice. Numerous considerations enter into the selection of planform parameters that lead to the final choice of wing planform. These include such items as cruise lift-drag ratio (LID}, weight, low-speed aerodynamic characteristics, and simplicity. Based on the mission requirements of a given supersonic aircraft, trade studies of the various considerations just mentioned establish the optimum wing planform. These trade studies must be reiterated every time technology is advanced, design philosophy reevaluated, and/or design requirements updated. The first-generation SST developed by Lockheed during the FAA/SST program of the 1960's featured a double-delta planform, tailless concept. The delta wing planform was Lockheed's first choice due to its lighter weight, superior low- speed aerodynamic characteristics, and simplicity. The arrow wing planform was Lockheed's second choice. This planform displayed superior cruise efficiency (cruise lift-to-drag ratio), but was heavier than the delta planform. Also, there was concern with regard to the low-speed aerodynamic charac- teristics of this planform. The variable-sweep wing concept

Concept and conduct of proof test of F-111 production aircraft

F-111 aircraft are in service with both the Tactical Air Command and the Strategic Air Command in the USAF. The aircraft performs primarily in the fighter-bomber role for TAC and as a strategic bomber capable of high and low altitude penetration with SAC. Over 100 000 flight hours have been accumulated by the F-111 fleet including a limited amount of combat flying. The broad spectrum of service requirements for the F-111 include short field take-off and landing capability, all weather and deep interdiction, close air support with large payloads, high altitude high speed air defence, and world-wide deployment with minimum tanker support. Major technical advancements were necessary in meeting these requirements, including most notably a variable-sweep wing. The unswept high aspect ratio positions provide for maximum range and endurance and short take-off and landing distances. The wing in moderate sweep positions provides an efficient configuration for sustained high subsonic and high altitude supersonic flight. Supersonic flight at sea level at speeds in excess of the capability of any other aircraft plus supersonic speeds at altitude in excess of Mach 2.0 are achieved in a most highly swept low aspect ratio wing position.