Flying Wing Research Articles

Researches on the aerodynamic and stealth characteristics for flying wing fighter

Flying wing configuration fighter is very promising for the advanced fighter in the future. This paper studies the aerodynamic and stealth characteristics of flying wing fighter. Firstly, the effect of sharp leading edge on the subsonic, transonic lift and drag characteristics and stealth characteristics of the fighter are studied. The results show that the sharpen of leading edge of the inboard wing could greatly improve the lateral stealth characteristics while maintaining the subsonic and transonic lift characteristics. In order to furtherly improve its aerodynamic and forward stealth characteristics, the aerodynamic/stealth collaborative design optimization method based on discrete adjoint is carried out. And the results presented that there are obvious contradictions in the subsonic, transonic and supersonic lift characteristics of the flying wing fighter, and the improvement of the forward stealth characteristics will also damage the subsonic lift characteristics. Therefore, it is necessary to make a compromise between the subsonic, transonic and supersonic lift characteristics, drag characteristics and stealth characteristics in the shape design of flying wing fighter.

Aerodynamic Assessment of a Control Strategy Based on Twist Morphing Wing in a Flying Wing Aircraft

Aerodynamic Assessment of a Control Strategy Based on Twist Morphing Wing in a Flying Wing Aircraft

Flight control of a flying wing aircraft based on circulation control using synthetic jet actuators

To achieve the nice stealth performance and aerodynamic maneuverability of a Flying Wing Aircraft (FWA), a longitudinal aerodynamic control technology based on circulation control using trailing-edge synthetic jet actuators was proposed without the movement of rudders. Effects on the longitudinal aerodynamic characteristics of a small-sweep FWA were investigated. Then, flight tests were carried out to verify the control abilities, providing a novel technology for the design of a future rudderless FWA. Results show that synthetic jets could narrow the dead zone area, improve the flow velocity near the trailing edge, and then move the trailing-edge separation point and the leading-edge stagnation point downwards, which make the effective Attack of Angle (AOA) increase, thereby enhancing the pressure envelope area. Circulation control based on synthetic jets could improve the lift, drag and nose-down moment. The variations of lift and nose-down moment decrease with the growth of AOA caused by the improved reverse pressure gradient and the weakened circulation control efficiency. Finally, synthetic jet actuators were integrated into the trailing edge of a small-sweep FWA, which could realize the roll and pitch control without deflections of rudders during the cruise stage, and the maximum roll and pitch angular velocity are 12.64 (°)/s and 8.51 (°)/s, respectively.

Novel yaw effector of a flying wing aircraft based on reverse dual synthetic jets

For achieving the nice stealth performance and aerodynamic maneuverability of a Flying Wing Aircraft (FWA), a novel yaw effector based on Reverse Dual Synthetic Jets (RDSJ) was proposed without the movement of rudders. Effects on aerodynamic characteristics of a small-sweep FWA and control mechanism were investigated by numerical simulations. Finally, reverse dual synthetic jet actuators were integrated into a real FWA and flight tests were firstly carried out. Numerical results show that RDSJ could make drag coefficient increase and weaken lift coefficient, which generate a yawing moment and a rolling moment in the same direction, realizing control of heading attitudes, but strong coupling with the pitching moment occurs at large angles of attack. For control mechanism, RDSJ could produce two reverse synthetic jets out of phases, improve the reverse pressure gradient and hence form alternate recirculation zones or even early large-area separation, which cause the rise of pressures before exits and the dip of pressures behind exits, achieving improvement of drag and the yawing moment. The results of flight tests support that RDSJ could realize control of heading attitudes without deflections of rudders during the cruise stage and achieve the maximal yaw angular velocity of 10.12(°)/s, verifying the feasibility of this novel yaw effector.

ArticleGust Alleviation by Active–Passive Combined Control of the Flight Platform and Antenna Array for a Flying Wing SensorCraft

SensorCraft is an intelligence, surveillance, and reconnaissance (ISR) system that integrates unmanned flight platforms and airborne antenna arrays. Under gust loads, the high–aspect–ratio, light–wing structure of SensorCraft has considerable bending and torsion deformation, affecting the flight performance of unmanned flight platforms and leading to the loss of antenna arrays’ electromagnetic performance. Taking SensorCraft as the background, a wing conformal antenna array was designed, an aircraft model with a passive wingtip device was established, a control law was developed by the LQG/LTR method, and a gust alleviation active–passive combined control method of a “LQG/LTR active controller + passive wingtip device” was proposed. By constructing an unsteady aerodynamic reduced–order model (ROM) based on the Volterra series and a conformal array pattern fast method based on the modal form, the effectiveness of the gust alleviation active–passive combined control method on the aircraft platform and antenna array was analyzed. The results show that structural deformation of the wing conformal antenna leads to changes in the main lobe gain, beam direction, and sidelobe level. The active–passive gust alleviation method has obvious advantages. Compared with the LQG/LTR active gust alleviation method, the peak value of wingtip displacement is reduced by 15.6%, and the peak value of the gain loss is reduced by 0.72 dB, which is conducive to better performance of the airborne conformal antenna array.

Design and Feasibility Study of Novel Flying Wing Carrier for Launching Small Satellites in Low Earth Orbit

In recent years, there has been a considerable growth in the demand for low-orbit satellites, leading to a need for more flexible and cost-effective launch systems. This study presents a low-cost “carrier-launcher” configuration designed for space missions in low earth orbit. The carrier is a remote-controlled unmanned flying wing that can fulfil the role of the first stage of a multi-stage earth-to-orbit launcher rocket. Making the carrier a flying wing increases its effectiveness and efficiency compared to other state-of-the-art options. The flying wing architecture allows for a significantly lighter carrier compared to the traditional aircraft. The launcher is carried on the wing’s upper surface and is released during a high-altitude almost “zero g” parabolic manoeuvre. A state-of-the-art analysis has been conducted to initialize and develop the carrier’s conceptual configuration. The aerodynamics and flight mechanics of the flying wing carrier were studied using the potential aerodynamic code Athena Vortex Lattice. The high-altitude launcher’s release manoeuvre has been investigated to properly assess the required installed thrust. Finite element analyses were also performed using NASTRAN to preliminarily evaluate the aeroelastic behaviour of the proposed “carrier-launcher” configuration. The overall results show the conceptual feasibility of the flying wing carrier for launching small satellites in low earth orbit. This study provides valuable insights into the development of cost-effective launch systems for the growing demand in the low-orbit satellite sector. Our proposed design has a maximum take-off mass of 122,000 kg, uses 4 Rolls-Royce UltraFan model engines, has a wingspan of 54 m, and can carry a 10,000 kg launcher to put a 460 kg payload in LEO. As it is an initial conceptual study, this investigation establishes an initial benchmark for forthcoming inquiries, hence providing a starting point of a breakthrough concept to foster its future development.

Parametric Investigation of Canards on a Flying Wing UAV Using the Taguchi Method

The current work investigates the effect of canard geometric characteristics on the performance of a lightweight flying wing Unmanned Aerial Vehicle (UAV), capable of both conventional and Vertical Take-Off and Landing (VTOL) flight. The canards are sized as horizontal stabilizers to enhance the UAV’s longitudinal stability and minimize trimming requirements during cruise. Using a Design of Experiments (DOE) approach and, specifically, the Taguchi method, six canards’ design parameters are investigated on three different levels. These parameters are the sweep angle (Λ), aspect ratio (AR), taper ratio (λ), vertical position in relation to the main wing (vpos), incidence angle (ic), and dihedral angle (Γ). An L27 orthogonal array (OA) is used to investigate the influence of these key design parameters using two performance criteria, namely the Lift-to-Drag ratio (L/D) and the pitching moment coefficient (Cm), at cruise conditions (designated as L/Dcruise and |Cm|cruise). The investigation is conducted by using high-fidelity Computational Fluid Dynamics (CFD) methods for each of the 27 configurations defined by the L27 OA, over a range of angles of attack. Based on the CFD results, two distinct combinations are derived for maximum L/Dcruise and minimum |Cm|cruise using the Signal-to-Noise ratio (SNR) analysis. The optimal design parameter combinations for the two performance criteria are A2B1C1D1E1F1 and A2B2C1D2E2F3, respectively. Finally, the Pareto Analysis of Variance (ANOVA) technique is conducted to define the contribution of each of the six design parameters on the L/Dcruise and the |Cm|cruise. More specifically, ic seems to have the most significant effect on L/Dcruise, whereas Λ is the most important parameter for |Cm|cruise.

Flush Airdata System on a Flying Wing Based on Machine Learning Algorithms

By using an array of pressure sensors distributed on the surface of an aircraft to measure the pressure of each port, the flush airdata sensing (FADS) system is widely applied in many modern aircraft and unmanned aerial vehicles (UAVs). Normally, the pressure transducers of the FADS system should be mounted on the leading edge of the aircraft, where they are sensitive to changes in pressure. For UAVs, however, the leading edge of the nose and wing may not be available for pressure transducers. In addition, the number of transducers is limited to 8–10, making it difficult to maintain accuracy in the normal method for FADS systems. An FADS system model for an unmanned flying wing was developed, and the pressure transducers were all located outside the regions of the leading edge areas. The locations of the transducers were selected by using the mean impact value (MIV), and ensemble neural networks were developed to predict the airdata with a very limited number of transducers. Furthermore, an error detection method was also developed based on artificial neural networks and random forests. The FADS system model can accurately detect the malfunctioning port and use the correct pressure combination to predict the Mach number, angle of attack, and angle of sideslip with high accuracy.

Coordinated Roll Control of Conformal Finless Flying Wing Aircraft

Attitude control of tailless aircraft is typically achieved through a system of trailing-edge flaps that produce coupled control effects in pitch, roll and yaw due to changes in lift and drag distributions. This paper presents a novel control allocation method which allows conformal trailing-edge controls to generate yawing moments with no coupling, thereby simplifying the required control suite. The method is demonstrated through modelling of a coordinated roll manoeuvre of a tailless aircraft in which yaw control is provided by trailing-edge flaps only. For a typical tailless aircraft configuration, it is shown that a coordinated rolling manoeuvre can be achieved with a reduction of at least 30% of the required control deflection compared to solutions using non-conformal controls such as split drag rudders.

Investigation on the Reynolds Number Effect of a Flying Wing Model with Large Sweep Angle and Small Aspect Ratio

The effect of the Reynolds number on the flying wing common research model with a large sweep angle and small aspect ratio is studied by numerical simulation. The helicity cross-flow transition correction of the Langtry–Menter model is established in view of the deficient prediction accuracies of existing transition prediction models in simulating cross-flow transition, and the helicity parameters are calibrated to improve the prediction accuracy of the numerical method. The transition prediction method is verified by the test results of the standard model transition of the DLR-F5 wing and small-aspect-ratio flying wing. At the same time, the numerical method is effectively verified by comparing it to the aerodynamic/torque results of the standard model of the small-aspect-ratio flying wing. On this basis, the variation laws with the Reynolds number are analyzed for the aerodynamic and flow field characteristics of the flying wing common research model with a large sweep angle and small aspect ratio. The numerical simulation is mainly carried out in the high subsonic speed range (M = 0.9), and the simulated Reynolds number range is from 2 × 106 to 80 × 106, which includes the test Reynolds number and flight Reynolds number. The influence of turbulence on the surface flow pattern and drag characteristics of the small-aspect-ratio standard model is analyzed. Calculation results show that when the turbulence is greater than 0.4%, it has a considerable influence on the surface laminar flow range and total drag coefficient. The results reveal that the Reynolds number mainly affects the magnitude of friction in the cruise state (small angle of attack). The problem of the Reynolds number in the self-aligning region of the flying wing standard model with a small aspect ratio is further studied, and the Reynolds number is found to be greater than 10 × 106. The drag and lift coefficients after the approach are close to the Reynolds number in the self-aligning region of the flying wing standard model with a small aspect ratio.

Flutter Modelling and Computation of a Flying Wing Aircraft

Flying wing aircraft has many encouraging advantages, but there still were many unsteady aerodynamic uncertainty of the structures. Dynamic stability is a most significant analysis object. This paper wants to propose the flutter characteristic of a real flying wing aircraft. Firstly, an efficient finite element modelling method was taken to find basic dynamic features of the presented aircraft with effective mode analysis. Typical dynamic characteristics including bending and twisting modes of the structure were analyzed to propose possible dynamic stability of the aircraft. Secondly, a frequency domain fluid-structure interaction method was proposed to solve flutter analysis problem of the aircraft. With different heights and Ma number, flutter features of the aircraft were presented. It was seen from the computational results that dynamic stability analysis is necessary for the flying wing aircraft with proper flight velocity and height. The present dynamic modelling method and flutter computation process can give an effect reference for the design of flying wing aircraft. Also it can provide useful suggestions for existing flying wing aircraft structures.

Computer Simulation Research of Self-adaptive Compensation Control Based on Neural Network for Control Surface Failure of Uncertain Tailless Flying Wing Aircraft

The conventional aircraft layout will contain the elevator, aileron, rudder three control surfaces, and thus generate three control moments to adjust the aircraft trajectory and flight attitude. However, the aircraft with multi-control plane flying wing layout needs to choose redundant new control plane configuration, which is quite different from the conventional aircraft layout in both control method and operation mode. Since there are many problems in the aircraft motion mathematical model during the ownership period, which will directly affect the control effect of the flight control system, in order to reduce dynamic error and improve the robustness and accuracy of the flight control system, this paper proposes an adaptive neural network algorithm. Combined with the current operating status of tailless flying wing aircraft and the research status of aircraft control system analysis, this paper mainly discusses the design differences between the adaptive control system and the original control system, based on the empirical case, to clarify the compensation effect of the adaptive control system.

Research on Command Allocation Method for Flying Wing Aircraft

Flying wing has nice aerodynamic efficiency, good loading and stealth characteristics, it is an ideal aerodynamic configuration for reconnaissance, transportation and bombing type aircraft. However, flying wing aircraft suffers from poor stability and control characteristics, which makes it difficult to design and control. Due to elimination of horizontal and vertical tails, flying wing aircraft has poor longitudinal and directional dynamic characteristics. Unlike most of the conventional aircraft, flying wing aircraft usually have a large number of control effectors, the redundant effectors are designed and required to provide adequate lift, pitch and yawing force or moments. There-fore, flying wing aircraft usually need to adopt reasonable control allocation algorithms to exploit the control capability of redundant effectors. Especially for large-aspect ratio flying wing with very high level of effector redundancy, the pursuit of performances, e.g., gust load alleviation control and maximum lift mode, makes control allocation extremely important and indispensable. This paper summarizes the typical characteristics of command allocation methods, and its applicability in different flight missions, providing a reference for the design of command allocation method for flying wing aircraft.

Research on multi-task command allocation method for flying wing aircraft

Flying wing aircraft adopts redundant multiple control surfaces configuration, especially the new type of drag based yaw control device is used in directional axis. Therefore, its command allocation problem has strong nonlinear and multi-axis coupling characteristics. According to the command allocation characteristics of flying wing aircraft, a multi-objective optimization based nonlinear programming method is used for command allocation, and the differences of the control surface command allocation results under different optimization targets is analyzed. The results are also carried out in the attitude tracking simulation based on flight control system designed with nonlinear dynamic inversion. It demonstrates that the different optimization targets have great and quite different impacts on flight states. It proves that the multi-objective optimization based control allocation method is a highly adaptable and flexible command allocation method for flying wing aircraft.

Airfoil Selection Procedure, Wind Tunnel Experimentation and Implementation of 6DOF Modeling on a Flying Wing Micro Aerial Vehicle.

Airfoil selection procedure, wind tunnel testing and an implementation of 6-DOF model on flying wing micro aerial vehicle (FWMAV) has been proposed in this research. The selection procedure of airfoil has been developed by considering parameters related to aerodynamic efficiency and flight stability. Airfoil aerodynamic parameters have been calculated using a potential flow solver for ten candidate airfoils. Eppler-387 proved to be the most efficient reflexed airfoil and therefore was selected for fabrication and further flight testing of vehicle. Elevon control surfaces have been designed and evaluated for longitudinal and lateral control. The vehicle was fabricated using hot wire machine with EPP styrofoam of density 50 Kg/. Static aerodynamic coefficients were evaluated using wind tunnel tests conducted at cruise velocity of 20 m/s for varying angles of attack. Rate derivatives and elevon control derivatives have also been calculated. Equations of motion for FWMAV have been written in a body axis system yielding a 6-DOF model. It was found during flight tests that vehicle conducted coordinated turns with no appreciable adverse yaw. Since FWMAV was not designed with a vertical stabilizer and rudder control surface, directional stability was therefore augmented through winglets and high wing leading edge sweep. Major problems encountered during flight tests were related to left rolling tendency. The left roll tendency was found inherent to clockwise rotating propeller as ‘P’ factor, gyroscopic precession, torque effect and spiraling slipstream. To achieve successful flights, many actions were required including removal of excessive play from elevon control rods, active actuation of control surfaces, enhanced launch speed during take off, and increased throttle control during initial phase of flight. FWMAV flew many successful stable flights in which intended mission profile was accomplished, thereby validating the proposed airfoil selection procedure, modeling technique and proposed design.

Lateral Stability and Control of a Flying Wing Configuration Aircraft

A numerical investigation was conducted on an Unmanned Combat Air Vehicle (UCAV) model which is a moderately swept, tailless and flying wing configuration. A wind tunnel test was performed to validate the numerical results. Model forces and moments were measured using a six-component strain-gauged force balance. The computations were performed using a commercial Reynolds-Averaged Navier-Stokes (RANS) flow solver. The numerical results show that the lateral unstable of the model is caused by asymmetrical breakdown of the body vortex, and the flaps installed can induce earlier breakdown and intensity loss of the leeward vortex resulting in the enhancement of lateral stability.

Full-Span Flying Wing Wind Tunnel Test: A Body Freedom Flutter Study

Aiming at the experimental test of the body freedom flutter for modern high aspect ratio flexible flying wing, this paper conducts a body freedom flutter wind tunnel test on a full-span flying wing flutter model. The research content is summarized as follows: (1) The full-span finite element model and aeroelastic model of an unmanned aerial vehicle for body freedom flutter wind tunnel test are established, and the structural dynamics and flutter characteristics of this vehicle are obtained through theoretical analysis. (2) Based on the preliminary theoretical analysis results, the design and manufacturing of this vehicle are completed, and the structural dynamic characteristics of the vehicle are identified through ground vibration test. Finally, the theoretical analysis model is updated and the corresponding flutter characteristics are obtained. (3) A novel quasi-free flying suspension system capable of releasing pitch, plunge and yaw degrees of freedom is designed and implemented in the wind tunnel flutter test. The influence of the nose mass balance on the flutter results is explored. The study shows that: (1) The test vehicle can exhibit body freedom flutter at low airspeeds, and the obtained flutter speed and damping characteristics are favorable for conducting the body freedom flutter wind tunnel test. (2) The designed suspension system can effectively release the degrees of freedom of pitch, plunge, and yaw. The flutter speed measured in the wind tunnel test is 9.72 m/s, and the flutter frequency is 2.18 Hz, which agree well with the theoretical results (with flutter speed of 9.49 m/s and flutter frequency of 2.03 Hz). (3) With the increasing of the mass balance at the nose, critical speed of body freedom flutter rises up and the flutter frequency gradually decreases, which also agree well with corresponding theoretical results.

Control Authority of a Camber Morphing Flying Wing

Flying wings achieve their roll, pitch, and yaw controllability exclusively through spanwise variations of lift and drag due to the lack of an empennage. Instead of using conventional control surfaces, which increase the aerodynamic drag due to gaps and discontinuities, a morphing structure can be employed. This work considers a previously developed camber morphing flying wing, which uses 10 electromechanical actuators to achieve spanwise-varying trailing edge deflections, resulting in variations of the lift and drag distributions and generating moments around the roll, pitch, and yaw axes. To investigate the control authority of the aircraft across all flight maneuvers, a series of multi-objective aerostructural optimizations is carried out. The performance for the individual maneuvers, as well as the interaction of simultaneous roll, pitch, and yaw moments and the resulting control authority, is assessed. Because the same actuators are used for all maneuvers, a high degree of interaction has been found between the achieved rolling and yawing moments. By minimizing the additional drag with the morphing actuation, the undesired effect of adverse yaw is minimized. The results describe a possible control strategy for flying wing aircraft with optimal drag performance.

R-functions in the Analytical Description of the Surface of a Flying Wing Unmanned Aerial Vehicle

R-functions in the Analytical Description of the Surface of a Flying Wing Unmanned Aerial Vehicle

Experimental Study on Plasma Flow Control of Symmetric Flying Wing Based on Two Kinds of Scaling Models

The symmetric flying wing has a simple structure and a high lift-to-drag ratio. Due to its complicated surface design, the flow field flowing through its surface is also complex and variable, and the three-dimensional effect is obvious. In order to verify the effect of microsecond pulse plasma flow control on the symmetric flying wing, two different sizes of scaling models were selected. The discharge energy was analyzed, and the force and moment characteristics of the two flying wings and the particle image velocimetry (PIV) results on their surface flow field were compared to obtain the following conclusions. The microsecond pulse surface dielectric barrier discharge energy density is independent of the actuator length but increases with the actuation voltage. After actuation, the stall angle of attack of the small flying wing is delayed by 4°, the maximum lift coefficient is increased by 30.9%, and the drag coefficient can be reduced by 17.3%. After the large flying wing is actuated, the stall angle of attack is delayed by 4°, the maximum lift coefficient is increased by 15.1%, but the drag coefficient is increased. The test results of PIV in the flow field of different sections indicate that the stall separation on the surface of the symmetric flying wing starts first from the outer side, and then the separation area begins to appear on the inner side as the angle of attack increases.