Transonic Rudder Buzz on Tailless Flying Wing UAV

Transonic rudder buzz responses based on the computational fluid dynamics or computational structural dynamics(CFD/CSD)loosely method are analyzed for a tailless flying wing unmanned aerial vehicle(UAV).The Reynolds-averaged Navier-Stokes(RANS)equations and finite element methods based on the detailed aerodynamic and structural model are established,in which the aerodynamic dynamic meshes adopt the unstructured dynamic meshes based on the combination of spring-based smoothing and local remeshing methods,and the lower-upper symmetric-Gauss-Seidel(LU-SGS)iteration and Harten-Lax-van Leer-Einfeldt-Wada(HLLEW)space discrete methods based on the shear stress transport(SST)turbulence model are used to calculate the aerodynamic force.The constraints of the rudder motions are fixed at the end of structural model of the flying wing UAV,and the structural geometric nonlinearities are also considered in the flying wing UAV with a high aspect ratio.The interfaces between structural and aerodynamic models are built with an exact match surface where load transferring is performed based on 3Dinterpolation.The flying wing UAV transonic buzz responses based on the aerodynamic structural coupling method are studied,and the rudder buzz responses and aileron,elevator and flap vibration responses caused by rudder motion are also investigated.The effects of attack,height,rotating angular frequency and Mach number under transonic conditions on the flying wing UAV rudder buzz responses are discussed.The results can be regarded as a reference for the flying wing UAV engineering vibration analysis.

Analysis of flow separation control using nanosecond-pulse discharge plasma actuators on a flying wing

Dielectric barrier discharge （DBD） plasma is one of most promising flow control method for its several advantages. The present work investigates the control authority of nanosecond pulse DBD plasma actuators on a flying wing model＇s aerodynamic characteristics. The aerodynamic forces and moments are studied by means of experiment and numerical simulation. The numerical simulation results are in good agreement with experiment results. Both results indicate that the NS-DBD plasma actuators have negligible effect on aerodynamic forces and moment at the angles of attack smaller than 16-. However, significant changes can be achieved with actuation when the model＇s angle of attack is larger than 16° where the flow separation occurs. The spatial flow field structure results from numerical simulation suggest that the volumetric heat produced by NS-DBD plasma actuator changes the local temperature and density and induces several vortex structures, which strengthen the mixing of the shear layer with the main flow and delay separation or even reattach the separated flow.

Experimental investigation of lift enhancement for flying wing aircraft using nanosecond DBD plasma actuators

he effects of the arrangement position and control parameters of nanosecond dielectric barrier discharge （NS-DBD） plasma actuators on lift enhancement for flying wing aircraft were investigated through wind tunnel experiments at a flow speed of 25 m s-1.The aerodynamic forces and moments were obtained by a six-component balance at angles of attack ranging from -4° to 28°.The lift,drag and pitching moment coefficients were compared for the cases with and without plasma control.The results revealed that the maximum control effect was achieved by placing the actuator at the leading edge of the inner and middle wing,for which the maximum lift coefficient increased by 37.8% and the stall angle of attack was postponed by 8° compared with the plasma-off case.The effects of modulation frequency and discharge voltage were also investigated.The results revealed that the lift enhancement effect of the NS-DBD plasma actuators was strongly influenced by the modulation frequency.Significant control effects were obtained atf =70 Hz,corresponding to F＋ ≈ 1.The result for the pitching moment coefficient demonstrated that the plasma actuator can induce the reattachment of the separation flows when it is actuated.However,the results indicated that the discharge voltage had a negligible influence on the lift enhancement effect.

Virtual flight test of pitch and roll attitude control based on circulation control of tailless flying wing aircraft without rudders

Circulation Control(CC) realizes rudderless flight control by driving compressed air jet to generate a virtual rudder surface, which significantly improves low detectability. The layout plan of combined control rudder surface is proposed based on the tailless flying wing aircraft. The closed-loop jet actuator system and stepless rudder surface switching control strategy are used to quantitatively study the control characteristics of circulation actuator for pitch and roll attitude through 3-DOF virtual flight test in a wind tunnel with a powered model at wind speed of 40 m/s. The results show that the combined use of circulation actuators can achieve bidirectional continuous and stable control of the aircraft’s pitch and roll attitude, with the maximum pitch rate of 12.3(°)/s and the maximum roll rate of 21.5(°)/s;the response time of attitude angular rate varying with the jet pressure ratio is less than 0.02 s, which can satisfy the control response requirements of aircraft motion stability for the control system;the jet rudder surface has a strong moment control ability, and the pitch moment of the jet elevator with a pressure ratio of 1.28 is the same as that of the mechanical elevator with 28° rudder deflection, which can expand the flight control boundary.

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 nosedown 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.

Aerodynamic performance enhancement of a flying wing using nanosecond pulsed DBD plasma actuator

Experimental investigation of aerodynamic control on a 35° swept flying wing by means of nanosecond dielectric barrier discharge （NS-DBD） plasma was carried out at subsonic flow speed of 20-40 m/s, corresponding to Reynolds number of 3.1 × 10^5-6.2× 10^5. In control condition, the plasma actuator was installed symmetrically on the leading edge of the wing. Lift coefficient, drag coefficient, lift-to-drag ratio and pitching moment coefficient were tested with and without control for a range of angles of attack. The tested results indicate that an increase of 14.5% in maximum lift coefficient, a decrease of 34.2% in drag coefficient, an increase of 22.4% in maximum lift-to-drag ratio and an increase of 2° at stall angle of attack could be achieved compared with the baseline case. The effects of pulsed frequency, amplitude and chord Reynolds number were also investigated. And the results revealed that control efficiency demonstrated strong dependence on pulsed fre- quency. Moreover, the results of pitching moment coefficient indicated that the breakdown of lead- ing edge vortices could be delayed by plasma actuator at low pulsed frequencies.

Aerodynamic investigation of twist angle variation based on wing smarting for a flying wing

In this paper, the effects of twist angle variation on aerodynamic coefficients and flow field on the wing with wing smarting approach are studied using numerical simulation. The simulation was performed using incompressible Reynolds-Averaged Navier-Stokes(RANS) equations based on the two-equation k-x Shear Stress Transport(SST) turbulent model for flow speed30 m/s and a Reynolds number of 69000. Investigations have been carried out for several twist angles and at a specific range of angles of attack. The twist applied is the type of geometric twist(wash-out), which is linearly distributed along the span. The test case is a lambda-shaped tailless aircraft with a wing fracture on the trailing edge, and a sweep angle 56°. The results show that with increasing twist angle, the aerodynamic efficiency improves over a wide range of angles of attack,but at 0° angle of attack it will decrease significantly. By increasing the angle of attack, the effect of twist on the flow field and aerodynamic coefficients will gradually decrease;hence, at a certain amount of angle of attack, the effect of twist will stop, that angle is called the neutral brink angle.Longitudinal stability analysis shows that by growing the twist angle, the conditions required for longitudinal stability are satisfied, and the pitch-up phenomenon will be delayed.

Transonic wing stall of a blended flying wing common research model based on DDES method

Numerical simulation of wing stall of a blended flying wing configuration at transonic speed was conducted using both delayed detached eddy simulation（DDES） and unsteady Reynolds-averaged Navier-Stokes（URANS） equations methods based on the shear stress transport（SST） turbulence model for a free-stream Mach number 0.9 and a Reynolds number 9.6 × 10. A joint time step/grid density study is performed based on power spectrum density（PSD） analysis of the frequency content of forces or moments, and medium mesh and the normalized time scale0.010 were suggested for this simulation. The simulation results show that the DDES methods perform more precisely than the URANS method and the aerodynamic coefficient results from DDES method compare very well with the experiment data. The angle of attack of nonlinear vortex lift and abrupt wing stall of DDES results compare well with the experimental data. The flow structure of the DDES computation shows that the wing stall is caused mainly by the leeward vortex breakdown which occurred at x/x= 0.6 at angle of attack of 14°. The DDES methods show advantage in the simulation problem with separation flow. The computed result shows that a shock/vortex interaction is responsible for the wing stall caused by the vortex breakdown. The balance of the vortex strength and axial flow, and the shock strength, is examined to provide an explanation of the sensitivity of the breakdown location. Wing body thickness has a great influence on shock and shock/vortex interactions, which can make a significant difference to the vortex breakdown behavior and stall characteristic of the blended flying wing configuration.

Reconfigurable Flight Control Design for Combat Flying Wing with Multiple Control Surfaces

With control using redundant multiple control surface arrangement and large-deflection drag rudders,a combat flying wing has a higher probability for control surface failures.Therefore,its flight control system must be able to reconfigure after such failures.Considering three types of typical control surface failures（lock-in-place（LIP）,loss-of-effectiveness（LOE） and float）,flight control reconfiguration characteristic and capability of such aircraft types are analyzed.Because of the control surface redundancy,the aircraft using the dynamic inversion flight control law already has a control allocation block.In this paper,its flight control configuration during the above failures is achieved by modifying this block.It is shown that such a reconfigurable flight control design is valid,through numerical simulations of flight attitude control task.Results indicate that,in the circumstances of control surface failures with limited degree and the degradation of the flying quality level,a combat flying wing adopting this flight control reconfiguration approach based on control allocation could guarantee its flight safety and perform some flight combat missions.

Three-axis coupled flight control law design forflying wing aircraft using eigenstructure assignment method

Due to elimination of horizontal and vertical tails,flying wing aircraft has poor longitudinal and directional dynamic characteristics.In addition,flying wing aircraft uses drag rudders for yaw control,which tends to generate strong three-axis control coupling.To overcome these problems,a flight control law design method that couples the longitudinal axis with the lateraldirectional axes is proposed.First,the three-axis coupled control augmentation structure is specified.In the structure,a‘‘soft/hard"cross-connection method is developed for three-axis dynamic decoupling and longitudinal control response decoupling from the drag rudders;maneuvering turn angular rate estimation and subtraction are used in the yaw axis to improve the directional damping.Besides,feedforward control is adopted to improve the maneuverability and control decoupling performance.Then,detailed design methods for feedback and feedforward control parameters are established using eigenstructure assignment and model following technique.Finally,the proposed design method is evaluated and compared with conventional method by numeric simulations.The influences of control derivatives variation of drag rudders on the method are also analyzed.It is demonstrated that the method can effectively improve the dynamic characteristics of flying wing aircraft,especially the directional damping characteristics,and decouple the longitudinal responses from the drag rudders.

Dihedral influence on lateral–directional dynamic stability on large aspect ratio tailless flying wing aircraft

The influence of dihedral layout on lateral–directional dynamic stability of the tailless flying wing aircraft is discussed in this paper. A tailless flying wing aircraft with a large aspect ratio is selected as the object of study, and the dihedral angle along the spanwise sections is divided into three segments. The influence of dihedral layouts is studied. Based on the stability derivatives calculated by the vortex lattice method code, the linearized small-disturbance equations of the lateral modes are used to determine the mode dynamic characteristics. By comparing 7056 configurations with different dihedral angle layouts, two groups of stability optimized dihedral layout concepts are created. Flight quality close to Level 2 requirements is achieved in these optimized concepts without any electric stability augmentation system.

Experimental Investigation on Aerodynamic Control of a Wing with Distributed Plasma Actuators

Experimental investigation of active flow control on the aerodynamic performance of a flying wing is conducted. Subsonic wind tunnel tests are performed using a model of a 35° swept flying wing with an nanosecond dielectric barrier discharge （NS-DBD） plasma actuator, which is installed symmetrically on the wing leading edge. The lift and drag coefficient, lift-to- drag ratio and pitching moment coefficient are tested by a six-component force balance for a range of angles of attack. The results indicate that a 44.5% increase in the lift coefficient, a 34.2% decrease in the drag coefficient and a 22.4% increase in the maximum lift-to-drag ratio can be achieved as compared with the baseline case. The effects of several actuation parameters are also investigated, and the results show that control efficiency demonstrates a strong dependence on actuation location and frequency. Furthermore, we highlight the use of distributed plasma actuators at the leading edge to enhance the aerodynamic performance, giving insight into the different mechanism of separation control and vortex control, which shows tremendous potential in practical flow control for a broad range of angles of attack.

Simulation on the dynamic stability derivatives of battle-structure-damaged aircrafts

Accurately evaluating the aerodynamic performance of a battle-structure-damaged aircraft is essential to enable the pilot to optimize the flight control strategy. Based on CFD and rigid dynamic mesh techniques,a numerical method is developed to calculate the longitudinal and longitudinal-lateral coupling forces and moments with small amplitude sinusoidal pitch oscillation, and the corresponding dynamic derivatives of two fragment-structure-damaged and two continuous-rod-damaged models modified from the SACCON UAV. The results indicate that, at the reference point set in this paper, additional positive damping is generated in fragment-damaged configurations;thus, the absolute values of the negative pitch dynamic derivative increase. The missing wingtip induces negative pitch damping on the aircraft and decreases the value of the pitch dynamic derivative. The missing middle wing causes a noticeable increase in the absolute value of the pitch dynamic derivative;the missing parts on the right wing cause the aircraft to roll to the right side in the dynamic process, and the pitch-roll coupling cross dynamic derivatives are positive. Moreover, the values of these derivatives increase as the damaged area on the right wing increases, and an optimal case with the smallest cross dynamic derivative can be found to help improve the survivability of damaged aircraft.