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.

Optimized Robust Filter for Uncertain Discrete Time System and Its Application to Flight Test

An optimized robust filtering algorithm for uncertain discrete-time systemsis presented. To get a series of computational equations, the uncertain part generated by theuncertain systematic matrix in the expression of the error-covariance matrix of time update stateestimation is optimized and the least upper bound of the uncertain part is given. By means of theseresults, the equivalent systematic matrix is obtained and a robust time update algorithm is builtup. On the other hand, uncertain parts generated by the uncertain observation matrix in theexpression of the error-covariance matrix of measurement update state estimation are optimized, andthe largest lower bound of the uncertain part is given. Thus both the time update and measurementupdate algorithms are developed. By means of the matrix inversion formula, the expression structuresof both time update and measurement update algorithms are all simplified. Moreover, the convergencecondition of a robust filter is developed to make the results easy to application. The results offlight data processing show that the method presented in this paper is efficient.

Power Supply for a Wireless Sensor Network： Airliner Flight Test Case Study

In this paper, we present hands-on experience related to on-going implementation in aircraft of power supply for a wireless sensor network deployed for aerodynamic flight tests. This autonomous battery-free power supply is capturing, managing and storing primary energy from the environment, using solar light and PV （photovoltaic） cells. For practical purposes, it is also equipped with an auxiliary power input. The specifications are detailed, the general architecture is presented and justified, and test results are discussed.

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.

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.

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.

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.

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.

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。

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.

Initial virtual flight test for a dynamically similar aircraft model with control augmentation system

To satisfy the validation requirements of flight control law for advanced aircraft,a wind tunnel based virtual flight testing has been implemented in a low speed wind tunnel.A 3-degree-offreedom gimbal,ventrally installed in the model,was used in conjunction with an actively controlled dynamically similar model of aircraft,which was equipped with the inertial measurement unit,attitude and heading reference system,embedded computer and servo-actuators.The model,which could be rotated around its center of gravity freely by the aerodynamic moments,together with the flow field,operator and real time control system made up the closed-loop testing circuit.The model is statically unstable in longitudinal direction,and it can fly stably in wind tunnel with the function of control augmentation of the flight control laws.The experimental results indicate that the model responds well to the operator＇s instructions.The response of the model in the tests shows reasonable agreement with the simulation results.The difference of response of angle of attack is less than 0.5°.The effect of stability augmentation and attitude control law was validated in the test,meanwhile the feasibility of virtual flight test technique treated as preliminary evaluation tool for advanced flight vehicle configuration research was also verified.

Airborne test flight of HY-2A satellite microwave scatterometer and data analysis

This paper introduces the background, aim, experimental design, configuration and data processing for an airborne test flight of the HY-2 Microwave scatterometer(HSCAT). The aim was to evaluate HSCAT performance and a developed data processing algorithm for the HSCAT before launch. There were three test flights of the scatterometer, on January 15, 18 and 22, 2010, over the South China Sea near Lingshui, Hainan. The test flights successfully generated simultaneous airborne scatterometer normalized radar cross section(NRCS), ASCAT wind, and ship-borne-measured wind datasets, which were used to analyze HSCAT performance. Azimuthal dependence of the NRCS relative to the wind direction was nearly cos(2w), with NRCS minima at crosswind directions, and maxima near upwind and downwind. The NRCS also showed a small difference between upwind and downwind directions, with upwind crosssections generally larger than those downwind. The dependence of airborne scatterometer NRCS on wind direction and speed showed favorable consistency with the NASA scatterometer geophysical model function(NSCAT GMF), indicating satisfactory HSCAT performance.

动压相似对缩比飞行的影响

模型飞行试验作为空气动力学研究的重要手段之一,正逐步成为新型飞机研发中一种有效和低风险的技术途径。目前低速飞行试验技术已经相对成熟,而高速飞行试验尚存在许多困难。本文分析结果显示,来流动压对缩比飞行试验具有显著影响。动压相似保证缩比飞机与原型机表现出相似的动力学特性,但同时也限制了缩比飞机的飞行性能。研究发现,对于本文中的高升阻比飞机,通过飞行方案设计,缩比飞机能够克服动压相似带来的约束,获得与原型机相仿甚至超过原型机的飞行性能,对后续开展其他缩比模型飞行试验具有一定的借鉴意义,同时可为我国低成本高性能靶标系统建设提供技术支持。

测耙数量对进气道稳态畸变特性测量的影响

在现代战斗机进气道相关试验中,受低速辅助进气门打开和隐身大S弯管道、大迎角侧滑角飞行姿态等复杂因素影响,进气道流场畸变较大,不同测耙数量的测试结果相差较大。国内外对进气道出口流场稳态畸变的测量及计算有较多研究,但鲜有测耙数量对进气道稳态畸变测量影响的全面对比研究。本文采用测耙测量进气道大压力梯度区和低压区并计算其稳态畸变,针对进气道风洞试验、全尺寸进发联合台架试验和进发匹配试飞,对常用的6支耙和8支耙形式测量结果进行对比研究。结果表明,对于形状复杂的战斗机进气道而言,进气道出口流场的测试需要尽可能采用8支测耙的测量方式,才能取得满意的结果。特别是舰载战斗机在舰面环境下,受舰首来流及偏流板反射等因素影响,进气道流场更加复杂,且发动机工作在特殊加力状态,对进发匹配稳定性要求也更加苛刻,更有必要采用8支测耙的测量方式。

基于GA-BP神经网络的大型客机气流角估计方法

为了解决硬件冗余难以克服的气流角传感器共因故障问题,进一步提高飞机气流角信号的可靠性,研究了基于GABP神经网络的气流角估计方法。通过BP神经网络融合姿态角、加速度、风速等参数来实现不依赖气流角传感器的气流角估计;引入遗传算法对神经网络权值和阈值进行全局优化,提高估计精度;对某大型客机的试飞数据预处理后用于模型的训练和测试。仿真结果表明,训练完成的GA-BP神经网络模型对气流角的估计值贴近实际值,稳定性和精度明显高于BP神经网络。上述方法给飞机增加一个余度的气流角信号,可用于传感器故障时为飞机提供可靠的气流角信号。

机载激光武器试验评估技术研究

由于毁伤机理和使用方法的不同,现有机载武器和传感器的试验方法不适用于机载激光武器。为研究其特殊的试验与评估技术,从国外典型机载激光武器的试验情况着手,分析了主要的试验阶段、试验内容、试验类别、管理特点等,对气动光学特性、出光特性、目标探测识别跟踪、光束控制和杀伤性试验等主要科目的内容、方法进行研究。研究表明,美国为了避免试验技术落后研制技术,采取了多计划安排一体推进、大力构建试验机、建设试验设施和循序渐进开展试验等管理措施促进相关的研究。建议加强试验技术的投入,适时建设体系化试验机,加强测试和靶标等领域的研究,以期为新质武器的成熟和应用提供帮助。

基于模型的空投任务飞行品质试飞设计及评估

重装空投飞行试验过程中,货物运动会导致飞机姿态变化过大,影响飞行安全;为了保证重装空投飞行试验安全,建立了空投过程数学模型,预测空投过程飞机响应、使用“预测→验证→比较→预测”的方法指导飞行试验循序渐进实施;在空投试飞前,分析空投货物时飞机响应特点,确定载机平台飞行品质试飞内容,评价空投构型飞行品质,作为空投飞行试验的前提;对于空投任务飞行试验设计从试飞状态选择、试飞员操纵方法设计两方面开展研究,以飞机姿态控制为目标,确定重力空投、牵引空投过程飞机飞行状态、给出飞行员操纵建议,设计了空投过程评价要素表,形成面向空投任务的飞行试验设计及飞行品质评价技术;提出的基于模型的试验设计方法及飞行品质评价方法可以有效保障空投任务执行,并全面充分地评价空投任务下的飞行品质。

民用飞机飞行中推力确定研究

确定飞行中发动机安装推力对飞机性能计算、发动机推力管理、经济性评价等具有重要意义。飞行中推力无法直接测量,需要在动力装置内部适当位置安装传感器测量性能参数,计算得出飞行中推力。研究了民用飞机动力装置飞行中推力确定方法,给出发动机安装净推力公式。介绍了推力计算的流程,利用某大涵道比涡扇发动机性能试飞数据,计算了该动力装置的安装净推力,并与参考值进行了对比分析,可为民用飞机型号设计提供参考。