Investigations of the mechanical response of dummy HTPB propellant grain under ultrahigh acceleration overload conditions using onboard flight-test measurements

In this paper,to study the mechanical responses of a solid propellant subjected to ultrahigh acceleration overload during the gun-launch process,specifically designed projectile flight tests with an onboard measurement system were performed.Two projectiles containing dummy HTPB propellant grains were successfully recovered after the flight tests with an ultrahigh acceleration overload value of 8100 g.The onboard-measured time-resolved axial displacement,contact stress and overload values were successfully obtained and analysed.Uniaxial compression tests of the dummy HTPB propellant used in the gunlaunched tests were carried out at low and intermediate strain rates to characterize the propellant's dynamic properties.A linear viscoelastic constitutive model was employed and applied in finite-element simulations of the projectile-launching process.During the launch process,the dummy propellant grain exhibited large deformation due to the high acceleration overload,possibly leading to friction between the motor case and propellant grain.The calculated contact stress showed good agreement with the experimental results,though discrepancies in the overall displacement of the dummy propellant grain were observed.The dynamic mechanical response process of the dummy propellant grain was analysed in detail.The results can be used to estimate the structural integrity of the analysed dummy propellant grain during the gun-launch process.

Automated Flight Test and System Identification for Rotary Wing Small Aerial Platform Using Frequency Responses Analysis

This paper proposes an autopilot system that can be used to control the small scale rotorcraft during the flight test for linear-frequency-domain system identification. The input frequency-sweep is generated automatically as part of the autopilot control command. Therefore the bandwidth coverage and consistency of the frequency-sweep are guaranteed to produce high quality data for system identification. Beside that, we can set the safety parameters during the flight test （maximum roll/pitch value, minimum altitude, etc.） so the safety of the whole flight test is guaranteed. This autopilot system is validated using hardware in the loop simulator for hover flight condition.

UAV flight test of plasma slats and ailerons with microsecond dielectric barrier discharge

Plasma flow control（PFC） is a promising active flow control method with its unique advantages including the absence of moving components, fast response, easy implementation, and stable operation. The effectiveness of plasma flow control by microsecond dielectric barrier discharge（μs-DBD）, and by nanosecond dielectric barrier discharge（NS-DBD） are compared through the wind tunnel tests, showing a similar performance between μs-DBD and NS-DBD. Furthermore, theμs-DBD is implemented on an unmanned aerial vehicle（UAV）, which is a scaled model of a newly developed amphibious plane. The wingspan of the model is 2.87 m, and the airspeed is no less than 30 m/s. The flight data, static pressure data,and Tufts images are recorded and analyzed in detail. Results of the flight test show that the μs-DBD works well on board without affecting the normal operation of the UAV model. When the actuators are turned on, the stall angle and maximum lift coefficient can be improved by 1.3° and 10.4%, and the static pressure at the leading edge of the wing can be reduced effectively in a proper range of angle of attack, which shows the ability of μs-DBD to act as plasma slats. The rolling moment produced by left-side μs-DBD actuation is greater than that produced by the maximum deflection of ailerons,which indicates the potential of μs-DBD to act as plasma ailerons. The results verify the feasibility and efficacy of μs-DBD plasma flow control in a real flight and lay the foundation for the full-sized airplane application.

Match Point Solution for Robust Flutter Analysis in Constant-Mach Prediction

This paper presents a method for robust flutter computation which uses flight altitude as the perturbation variable in order to obtain a match point solution. The air density and sound speed of standard atmosphere model are approximated as the polynomial function of altitude, such that the flight altitude becomes the single perturbation variable that describes the aeroelastic system. The uncertainties of generalized stiffness and damping are considered and the uncertain aeroelastic system can be formulated as linear fractional transformation （LFT） representation which is suitable for/.t analysis framework. Finally, the match point solutions of robust flutter margins can be computed with structured singular value （SSV） theory. The robust flutter analysis method provided in this paper is suitable for constant-Mach flight flutter test and provides valuable reference for flight envelope expansion.

Model Identification of a Micro Air Vehicle

This paper is focused on the model identification of a Micro Air Vehicle （MAV） in straight steady flight condition. The identification is based on input-output data collected from flight tests using both frequency and time dorrtain techniques. The vehicle is an in-house 40 cm wingspan airplane. Because of the complex coupled, multivariable and nonlinear dynamics of the aircraft, linear SISO structures for both the lateral and longitudinal models around a reference state were derived. The aim of the identification is to provide models that can be used in future development of control techniques for the MAV.

A Miniature Robotic Plane Meteorological Sounding System

This article presents a miniature robotic plane meteorological sounding system (RPMSS), which consists of three major subsystems: a miniature robotic plane, an air-borne meteorological sounding and ?ight control system, and a ground-based system. Take-o? and landing of the miniature aircraft are guided by radio control, and the ?ight of the robotic plane along a pre-designed trajectory is automatically piloted by an onboard navigation system. The observed meteorological data as well as all ?ight information are sent back in real time to the ground, then displayed and recorded by the ground-based computer. The ground-based subsystem can also transmit instructions to the air-borne control subsystem. Good system performance has been demonstrated by more than 300 hours of ?ight for atmospheric sounding.

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.

Novel roll effector based on zero-mass-flux dual synthetic jets and its flight test

The autonomous and controllable Dual Synthetic Jet Actuator(DSJA)is firstly integrated into the Unmanned Aerial Vehicle(UAV),and flight tests without the deflection of rudders are carried out to verify the viability of DSJA to control the attitudes of UAV during cruising.DSJA is improved into an actuator with two diaphragms and three cavities,which has higher energy levels.Actuators,differentially distributed on both sides of the wings,are installed on the trailing edge close to the wing tips.Flight tests,containing Differential Circulation Control(DCC)using double-side actuators,Positive Circulation Control(PCC)using left-side actuators and Negative Circulation Control(NCC)using right-side actuators,are implemented at cruising speed of 25 m/s.Results show that roll attitude control without rudders could be realized by DSJAs.DCC and NCC can generate the rightward roll and yaw angular velocity,prompting UAV to turn right.The stronger controlling ability can be achieved by DCC,with the maximum roll angular velocity of 15.62(°)/s.PCC can generate a rightward roll moment,but a leftward yaw moment will be produced at the same time.Leftward yaw induces the leftward rolling moment,which weakens the roll control effect,making UAV keep to yaw to the left with a small slope.

Multiple hierarchy risk assessment with hybrid model for safety enhancing of unmanned subscale BWB demonstrator flight test

To explore the low-speed characteristics of the Blended-Wing-Body(BWB)configuration for future civil aircraft,a series of unmanned subscale demonstrators have been developed and tested by our research team.During this process,specific safety risks deriving from uncertain design features,system unreliability,and insufficient personnel experience caused continuous flight test mishaps and the risk mechanism was not clear.Local and trial-and-error learning driven safety improvements took few effects on mishap prevention,so our focus was turned to look for systematic safety strategies.This paper establishes a systems theory based hybrid model to integrate the physical system reliability analysis techniques with the system dynamics method for illustrating the multiple risk interactions of the demonstrator flight test involving organizational,human resource and technical system factors.Using the prior BB-5 demonstrator as a case,the hybrid model simulation represents its historical risk evolution process,which verifies the model rationality.Derived risk control strategies reduced the mishap rate of a new demonstrator called BB-6 Sprit.The paper also shows the extended hybrid model can be applied on safety management of unmanned aerial vehicles from the initial period of vehicle development.

Conceptual design and flight test of two wingtipdocked multi-body aircraft

To overcome the drawbacks such as large wing deformations,poor performance encountering gusts,limits in taking off and landing,inconvenience of transportation of High-Altitude Long-Endurance(HALE)Unmanned Aerial Vehicles(UAVs),a new conceptual aircraft called wingtip-docked Multi-Body Aircraft(MBA)has attracted lots of attentions.Aiming to investigate the feasibility of this concept,two UAV models were designed,manufactured and connected by a wingtip-docking mechanism,which only allows the relative roll motion between the two aircraft.The trim solution of the two connected aircraft is firstly obtained by solving the developed nonlinear flight dynamic equations,followed by the stability analysis based on the linearized model.The results show that the connected aircraft is inherently unstable and cannot fly without a reasonable flight control system.A set of Proportional-Integral-Derivative(PID)control laws was then developed and implemented in the two experimental aircraft.The success of the flight tests show that the flight control can effectively eliminate the unstable motion and the wingtip-docked MBA is controllable and feasible。

Progress in flight tests of hypersonic boundary layer transition

Boundary layer transition(BLT)can cause a sharp rise in heat flux and skin friction,which can seriously affect the flight performance and safety of hypersonic flight vehicles.Therefore,the mechanism,prediction and control of transition have become important issues that must be dealt with for the development of advanced flight vehicles,and it is also a research hotspot of particular interest to major aerospace countries.Compared to other transition research approaches,model flight tests can better present the transition problems under real flight conditions,thus have been carried out extensively over the past 30 years.The United States,Germany,France,Australia,and other countries have carried out transition research based on flight tests,such as the Pegasus wing-glove crossflow transition and the Hypersonic Boundary Layer Transition(HyBOLT)transition control flight test of the United States,the joint research project of the Hypersonic International Flight Research and Experimentation-1(HIFiRE-1)circular cone and the HIFiRE-5 elliptic cone transition flight tests between the United States and Australia,the flight test of compression surface transition of the scramjet forebody(LEA)in France and so on.Although these flight tests suffered various setbacks,they still obtained valuable transition data.Recently,the United States is carrying out the concave-surface transition flight tests of Hypersonic Boundary Layer Transition(BOLT)and BOLT-II.Since its first model flight test mission for verification purpose launched successfully in 2015,several hypersonic BLT flight tests have been conducted by China Aerodynamics Research and Development Center(CARDC).The flight tests have measured valid transition data under flight conditions,obtained the transition front and its dynamical variation on blunt cones at various angles of attack and a lifting body Hypersonic Transition Research Vehicle(HyTRV).The crossflow traveling waves in high-altitude flight were measured for the first time,and our understanding of hypersonic BLT has been greatly improved.

Virtual flight test techniques to predict a blended-wing-body aircraft in-flight departure characteristics

As one of the promising configurations of the next generation of commercial aircraft,research on departure characteristics of the Blended-Wing-Body(BWB)is of great signification to safe flight limits.A three-degree-of-freedom(3-DOF)virtual flight test in a wind tunnel has been implemented for a candidate configuration to predict the departure characteristics.The support mechanism,the test model and the control law of the virtual flight test are introduced.In order to show the relationship between virtual flight test and actual flight test,the similarity criterion is also given.In open loop,the model has mild oscillations in the longitudinal and lateral directions,which are stable in closed-loop.The effect of flight control has been verified in virtual flight and actual subscale flight test.The analysis of system identification results indicate that the model has a good response to the excitation signal,and the response is in reasonable agreement with the flight test.Finally,the virtual flight departure test results are compared with the flight test.It shows that there is a good correspondence between the angle of attack and the elevator deflection at departure.This gives promising evidence of the practicability of virtual flight testing to predict departure of a BWB.

Preliminary study on heat flux measurement data of TT-0 flight test

With the explosive development of aerospace science,the design of the new generation airliner at higher speeds is attracting more attentions.To achieve this goal,it is necessary to achieve accurate prediction of the aerodynamic heating/force loads and successful reduction of drag and heat flux.As a remedy for the existing studies which are based upon the CFD and wind tunnel tests,this study presents a flight test for the drag and heat reduction spike technology.The principal goals of this flight test were to provide reference for verifying the accuracy of the prediction technology on ground and promote the development of the drag and heat reduction technology.By adopting the OS-X rocket,the TT-0 test vehicle designed by Shenyang Aircraft Design&Research Institute reached a maximum Mach number of 5.8 and a maximum altitude of 38 km.Hypersonic and supersonic pressure data by pressure scanning valves and heat fluxes by gauges at different locations were obtained successfully.Also,heat fluxes obtained by in-house CFD code are illustrated in comparison with the flight data.The results indicate that the numerical errors are large in most cases.More technologies,such as more CFD codes and more numerical procedures,should be adopted to conduct studies on this issue in the future.

A practical filter error method for aerodynamic parameter estimation of aircraft in turbulence

It is common for aircraft to encounter atmospheric turbulence in flight tests.Turbulence is usually modeled as stochastic process noise in the flight dynamics equations.In this paper,parameter estimation of nonlinear dynamic system with both process and measurement noise was studied,and a practical filter error method was proposed.The linearized Kalman filter of first-order approximation was used for state estimation,in which the filter gain,along with the system parameters and the initial states,constituted the parameter vector to be estimated.The unknown parameters and measurement noise covariance were estimated alternately by a relaxation iteration method,and the sensitivities of observations to unknown parameters were calculated by finite difference approximation.Some practical aspects of the method application were discussed.The proposed filter error method was validated by the flight simulation data of a research aircraft.Then,the method was applied to the flight tests of a subscale aircraft,and the aerodynamic stability and control derivatives were estimated.All the estimation results were compared with the results of the output error method to demonstrate the effectiveness of the approach.It is shown that the filter error method is superior to the output error method for flight tests in atmospheric turbulence.

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.

Intelligent modeling and identification of aircraft nonlinear flight

In this paper, a new approach has been proposed to identify and model the dynamics of a highly maneuverable fighter aircraft through artificial neural networks（ANNs）. In general, aircraft flight dynamics is considered as a nonlinear and coupled system whose modeling through ANNs, unlike classical approaches, does not require any aerodynamic or propulsion information and a few flight test data seem sufficient. In this study, for identification and modeling of the aircraft dynamics, two known structures of internal and external recurrent neural networks（RNNs） and a proposed structure called hybrid combined recurrent neural network have been used and compared.In order to improve the training process, an appropriate evolutionary method has been applied to simultaneously train and optimize the parameters of ANNs. In this research, it has been shown that six ANNs each with three inputs and one output, trained by flight test data, can model the dynamic behavior of the highly maneuverable aircraft with acceptable accuracy and without any priori knowledge about the system.

Design and experimental study of a new flapping wing rotor micro aerial vehicle

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.

Longitudinal Control for Balloon-Borne Launched Solar Powered UAVs in Near-Space

Aiming to improve the pull-up control performance in the process of releasing balloonborne solar powered UAVs(Unmanned Aerial Vehicles),this paper establishes the full flight mechanics equations with flexible modes,and proposes the control method suitable for engineering application.To be specific,the authors first calculate the real aerodynamic force on horizontal stabilizer by comparing the fuselage deformation in ballooning test with that in static loading test.Furthermore,considering fuselage elastic deformation,the pitching moment coefficient is obtained and the influence of airspeed and elevator angle on pitching moment coefficient and control surface efficiency are analysed.Second,the authors establish a complete flight mechanics model,including elastic structural dynamic model and rigid flight dynamic model,by comprehensively considering the aerodynamic data,the relationship between fuselage deformation and load,as well as the ballooning test.Third,the authors perform the numerical simulation and comparison study on control performance between rigid model and flexible model.Moreover,the authors implement model modification based on the low altitude flight test and steady-state point analysing.Finally,a scaled UAV is used to complete the balloon-borne launching test.The results show that the longitudinal control method can analyse the longitudinal aerodynamics and control characteristics accurately,and could be effectively utilized in the pull-up control of the balloon-borne solar powered UAV.

Investigation of lateral-directional aerodynamic parameters identification method for fly-by-wire passenger airliners

A new identification method is proposed to solve the problem of the influence on the loaded excitation signals brought by high feedback gain augmentation in lateral-directional aerodynamic parameters identification of fly-by-wire（FBW） passenger airliners. Taking for example an FBW passenger airliner model with directional relaxed-static-stability, through analysis of its signal energy distribution and airframe frequency response, a new method is proposed for signal type selection, signal parameters design, and the appropriate frequency relationship between the aileron and rudder excitation signals. A simulation validation is presented of the FBW passenger airliner＇s lateral-directional aerodynamic parameters identification. The validation result demonstrates that the designed signal can excite the lateral-directional motion mode of the FBW passenger airliner adequately and persistently. Meanwhile, the relative errors of aerodynamic parameters are less than 5%.