Flight Test Methods Research Articles

Identification of Rolling Resistance Coefficients for Aircraft Tires on Unsurfaced Airfields

The paper describes four test methods for determining the rolling resistance of aircraft tires on unsurfaced airfields. Evaluations of a flight-test method, an instrumented-vehicle method, a soil stress-deformation method, and a tow-test method were performed on a grassy airfield. The test results showed that rolling resistance is more dependent upon surface roughness than on grass length. The benefits and usefulness of each method are also presented.

Comparison of flight test data with a computational fluid dynamics model of a Scottish Aviation Bulldog aircraft

AbstractThe following paper presents detailed aerodynamic data of a Scottish Aviation Bulldog light aircraft. The data is taken from the pre-stall region of the aircraft flight envelope through two flight test methods and from a geometrically accurate computational fluid dynamics (CFD) model of the full scale aircraft, which was meshed in Ansys ICEM CFD and solved in Ansys Fluent. The fidelity of the CFD model was achieved by development of a CATIA solid model with surfaces matching a spatial point cloud of the aircraft taken using a 3D laser scanner. Following a CFD verification process, a 3·4m hybrid mesh with a Spalart-Allmaras (SA) turbulence model was found to give the best overall lift and drag characteristics. Further detailed comparisons with the glide flight test data showed the CFD drag polar to have 63% lower zero lift drag, although this discrepancy was related to the simplification of the original CATIA surface model, which excluded the undercarriage, aerials and other protuberance drags. Inclusion of estimates of these sources of drag resulted in a match in zero lift drag to within 15% and a maximum lift to drag of 10:1 which was within 11% of the glide flight test result. The remaining drag discrepancy is attributed to other effects including trim drag and the surface finish of the actual aircraft.

MIMO System Stability Margin and its Flight Test Method

Aiming at the multiple input multiple output (MIMO) flight control system evaluation, put forward MIMO system stability margin definition, also put forward to get the MIMO system stability margin of the flight test method. The simulation results show that, the method is feasible.

Designing similarity measurement with distance measure and application on laterally directional mode flight test

Similarity measure construction has been proposed as fault detection of flight test method in order to obtain the primary control surface stuck and the combination stuck of primary control. Similarity measures were obtained through analyzing the certainty and uncertainty of fuzzy membership functions, which were designed based on well-known Hamming distance. It was applied to the fault detection of primary control surface stuck of uninhabited aerial vehicle (UAV). At post-failure control surface, if the UAV is controllable and trimmable using other control surfaces, the UAV is able to fly or return to the safety region through reconfiguration of flight control system. To detect the fault, similarity measure computations were carried out. This result could be applicable with the real-time parameter estimation method. By monitoring the value of coefficients due to the control surface deviation, it becomes aware that the control surface fault occurs or not. The control surface stuck position and value were separated by comparing the trim value with the reference value. This is the advantage of increasing in reliability without adding sensors or with additional low cost.

Performance Flight Testing of Small, Electric Powered Unmanned Aerial Vehicles

Flight testing methods are developed to find the drag polar for small unmanned aerial vehicles powered by electric motors with fixed-pitch propellers. Wind tunnel testing was used to characterize the propeller-motor efficiency. The drag polar was constructed using data from flight tests. This drag polar was then used to find the following performance parameters: maximum velocity, minimum velocity, velocity for maximum range, velocity for maximum endurance, maximum rate of climb, maximum climb angle, minimum turn radius, maximum turn rate, and maximum bank angle. The developed flight testing methods are used to characterize the performance of a small UAV.

지상관측법 및 DGPS 기법을 활용한 이/착륙 성능 비행시험 비교

비행시험은 개발 항공기의 다양한 성능검증을 포함하여 감항기술기준에 대한 안전성을 입증하는 최종시험으로서 중요성이 크다. 특히, 민간 항공기의 비행시험은 개발업체의 경험 및 노하우에 의해 기술 보호성격이 강하며, 다양한 시험항목과 비행조건에 의해 비용이나 시간측면에서 효율적인 비행시험 수행 및 비행시험 기법의 확보는 필수적이다. 본 논문에서는 이륙 및 착륙성능 입증을 위한 이착륙 비행시험에 대해 지상관측법과 DGPS를 활용한 비행시험 기법을 제안하고 실제 비행시험을 수행하여 그 결과를 비교하였다. The flight test is last means of compliance to satisfy airworthiness standards and important to evaluate the performance and safety of the developed aircraft. The flight test technologies are obtained from great numbers of experiences and know-hows and protected. In addition, flight test should be conducted efficiently since its various test conditions and items. Therefore, it is requisite to secure efficient flight test methods. This paper discusses the flight test methods for take-off and landing performance and two kinds of techniques are proposed. By performing real flight tests, they are compared with each other and analyzed through the flight analysis.

Fault detection and isolation of DURUMI-II using similarity measure

This paper describes the flight test method for studying the primary control surface stuck condition and the combination stuck of the primary control. An aircraft must show controllability and trimmability under post-failure conditions. An aircraft is successfully tested under various fault conditions. It is recognized that a control surface fault is detected by monitoring the value of the coefficients related to the control surface deviation. The control surface stuck position is determined by comparing the trim value with the reference value. To detect and isolate the fault, an analysis that employs the real-time parameter estimation method is used. If the flight control system is reconfigured using online estimates of aircraft parameters from a real-time parameter estimation scheme, the reliability increases without the addition of sensors or additional cost.

자동화 비행시험기법에 의한 소형 무인헬리콥터의 파라메터 추정

In this paper dynamic modeling parameters were estimated using a frequency domain estimation method. A systematic flight test method was employed using preprogrammed multistep excitation of the swashplate control input. In addition when one axis is excited, the autopilot is engaged in the other axis, thereby obtaining high-quality flight data. A dynamic model was derived for a small scale unmanned helicopter (CNUHELI-020, developed by Chungnam National University) equipped with a Bell-Hiller stabilizer bar. Six degree of freedom equations of motion were derived using the total forces and moments acting on the small scale helicopter. The dynamics of the main rotor is simplified by the first order tip-path plane, and the aerodynamic effects of fuselage, tail rotor, engine, and horizontal/vertical stabilizer were considered. Trim analysis and linearized model were used as a basic model for the parameter estimation. Doublet and multistep inputs are used to excite dynamic motions of the helicopter. The system and input matrices were estimated in the frequency domain using the equation error method in order to match the data of flight test with those of the dynamic modeling. The dynamic modeling and the flight test show similar time responses, which validates the consequence of analytic modeling and the procedures of parameter estimation.

IN-FLIGHT IDENTIFICATION OF THE AUGMENTED FLIGHT DYNAMICS OF THE RMAX UNMANNED HELICOPTER

The flight dynamics of the Yamaha RMAX unmanned helicopter has been investigated, and mapped into a six degrees of freedom mathematical model. The model has been obtained by a combined black-box system identification technique and a classic model-based parameter identification approach. In particular, the closed-loop behaviour of the built-in attitude control system has been studied, to support the decision whether to keep it as inner stabilization loop or to develop an own stability augmentation system. The flight test method and the test instrumentation are described in detail; some samples of the flight test data are compared to the model outputs as validation, and an overall assessment of the built-in stabilization system is supplied.

Parameter identification with a controlled free flying model of a spaceplane

A flight test method for the investigation of flying qualities of a spaceplane with a radio controlled model is presented in this paper. Therefore a relatively simple model is equipped with several sensors and a telemetry transmitter. The model is dropped from a radio controlled carrier vehicle at an altitude of about 700 m. Because of the instability and the high wing loading a dynamically scaled model of the investigated spaceplane configuration PHOENIX cannot be controlled by vision from ground. A computer on ground is used to operate the model automatically by computing feedforward and feedback control algorithms. To avoid the high risk of a landing the model is recovered by a parachute. The measured data are used to estimate the parameters of a quasi-steady aerodynamic model with an equation error method. A comparison of these parameters with available wind tunnel data indicates, that the quality of the estimated parameters is comparatively high.

System identification and stability evaluation of an unmanned aerial vehicle from automated flight tests

This paper presents a consequence of the systematic approach to identify the aerodynamic parameters of an unmanned aerial vehicle (UAV) equipped with the automatic flight control system. A 3-2-1-1 excitation is applied for the longitudinal mode while a multi-step input is applied for lateral/directional excitation. Optimal time step for excitation is sought to provide the broad input bandwidth. A fully automated programmed flight test method provides highquality flight data for system identification using the flight control computer with longitudinal and lateral/directional autopilots, which enable the separation of each motion during the flight test. The accuracy of the longitudinal system identification is improved by an additional use of the closed-loop flight test data. A constrained optimization scheme is applied to estimate the aerodynamic coefficients that best describe the time response of the vehicle. An appropriate weighting function is introduced to balance the flight modes. As a result, concurrent system models are obtained for a wide envelope of both longitudinal and lateral/directional flight maneuvers while maintaining the physical meanings of each parameter.

무인항공기의 시스템 식별을 위한 비행시험기법

In this paper, a flight test method is described for the system identification of the unmanned aerial vehicle equipped with an automatic flight control system. Multistep inputs are applied for both longitudinal mode and lateral/directional excitation. Optimal time step for excitation is sought to provide the broad input bandwidth. A programmed mode flight test method provides high-quality flight data for system identification using the flight control computer with the longitudinal and lateral/directional autopilot which enables the separation of each motion during the flight test. In addition, exact actuating input that is almost equivalent to the designed one guarantees the highest input frequency attainable. Several repetitive flight tests were implemented in the calm air in order to extract the consistent system model for the air vehicle. The enhanced airborne data acquisition system endowed the high-quality flight data for the system identification. The flight data were effectively used to the system identification of the unmanned aerial vehicle.

Flight Test Identification and Simulation of a UH-60A Helicopter and Slung Load

Helicopter slung-load operations are common in both military and civil contexts. Helicopters and loads are often qualified for these operations by means of flight tests, which can be expensive and time consuming. There is significant potential to reduce such costs both through revisions in flight-test methods and by using validated simulation models. To these ends, flight tests were conducted at Moffett Field to demonstrate the identification of key dynamic parameters during flight tests (aircraft stability margins and handling-qualities parameters, and load pendulum stability), and to accumulate a data base for simulation development and validation. The test aircraft was a UH-60A Black Hawk, and the primary test load was an instrumented 8- by 6- by 6-ft cargo container. Tests were focused on the lateral and longitudinal axes, which are the axes most affected by the load pendulum modes in the frequency range of interest for handling qualities; tests were conducted at airspeeds from hover to 80 knots. Using telemetered data, the dynamic parameters were evaluated in near real time after each test airspeed and before clearing the aircraft to the next test point. These computations were completed in under 1 min. A simulation model was implemented by integrating an advanced model of the UH-60A aerodynamics, dynamic equations for the two-body slung-load system, and load static aerodynamics obtained from wind-tunnel measurements. Comparisons with flight data for the helicopter alone and with a slung load showed good overall agreement for all parameters and test points; however, unmodeled secondary dynamic losses around 2 Hz were found in the helicopter model and they resulted in conservative stability margin estimates.

Robust Flutter Margins of an F/A-18 Aircraft from Aeroelastic Flight Data

It is essential to determine eight conditions at which aeroelastic instabilities occur. Flight-test methods can be dangerous and costly, whereas analytical methods may not accurately predict the eutter boundary. An approach to computing eutter boundaries based on the structured singular value is presented. The aeroelastic system is formulated inarobust stabilityframework byparameterizingaround dynamicpressure and introducinguncertainty operators to account for modelingerrors. Flight data can be used tovalidate the robust systemmodel and increase accuracy of the eutter margin estimate. Flutter margins are generated for the F/A-18 Systems Research Aircraft using this method and traditional methods. Extensive e ight data from conditions throughout the eight envelope areanalyzed todeterminethelevelsof uncertaintyinatheoreticalmodel. Therobust marginsdemonstratethat the e ight conditions for e utter may lie closer to the eight envelope than previously estimated by traditional analysis.

Effects of thrust line offset on neutral point determination in flight testing

On aircraft with high or low thrust lines, conventional stability flight test methods result in shifted neutral points, which do not correspond to the actual pitch stability neutral points of the aircraft. Specifically, e.g., an elevator-position neutral point extrapolated from flight test data of an aircraft with a high thrust line, may be significantly behind the actual stick-fixed neutral point, causing a potential hazard. This implies that slopes of elevator position and stick force vs velocity diagrams do not necessarily mean that the aircraft is stable in pitch.

A method for flight-test determination of propulsive efficiency and drag

A flight-test method is described from which propulsive efficiency as well as parasite and induced drag coefficients can be directly determined using relatively simple instrumentation and analysis techniques. The method uses information contained in the transient response in airspeed for a small power change in level flight in addition to the usual measurement of power required for level flight. Measurements of pitch angle and longitudinal and normal acceleration are eliminated. The theoretical basis for the method, the analytical techniques used, and the results of application of the method to flight-test data are presented. Flight-test data showed performance parameters measured with a standard deviation of about 0.8% for propulsive efficiency, 0.3% for parasite drag coefficient, and 8% for the airplane efficiency factor, e.

Practical flight test method for determining reciprocating engine cooling requirements

It is pointed out that efficient and effective cooling of air-cooled reciprocating aircraft engines is a continuing problem for the general aviation industry. Miley et al. (1981) have reported results of a study regarding the controlling variables for cooling and installation aerodynamics. The present investigation is concerned with experimental methods which were developed to determine cooling requirements of an instrumented prototype or production aircraft, taking into account a flight test procedure which has been refined and further verified with additional testing. It is shown that this test procedure represents a straightforward means of determining cooling requirements with minimal instrumentation. Attention is given to some background information, the development history of the NACA cooling correlation method, and the proposed modification of the NACA cooling correlation.

A new concept of static stability and its flight testing in supersonic flight

The current concept of static stability is not adequate for supersonic flight since it does not account for the influence of forces and moments due to altitude perturbations, which have a significant effect on the longitudinal motion. A new concept is proposed that accounts for altitude influence in an adequate manner. This concept, which is closely related to constant energy consideration, is based on a dynamic stability analysis with particular reference to the exponential characteristic modes of the airplane. Furthermore, it is shown that the well-known relation between static stability and the variation of elevator angle with speed is not valid in supersonic flight. As a consequence, the flight test methods currently used for determining static stability also are not adequate for supersonic flight. A new flight test method is proposed that provides an indication of static stability using the variation of elevator angle with altitude (or dynamic pressure, respectively). The points addressed may be of particular significance in regard to existing flying qualities requirements and criteria.

A method for in-flight measurement of ground effect on fixed-wing aircraft.

Flight test method for evaluating ground effect on fixed-wing aircraft in which pilot flies at constant angle of attack and power setting during landing approach

Automatic Built-in Test of Advanced Avionics Systems Part II: Design Technioues for Automatic Built-in-Test

Need and justification for automatic, in-flight, thorough built-in test capabilities in advanced avionics systems is discussed and the need for a systems design approach is stressed. The role of an airborne digital computer for sequencing, conducting and evaluating the built-in test and the great flexibility of this computer, its characteristics, program and interface are developed and applicable test methods and design techniques are illustrated by typical examples. Short comings of methods used to continuously monitor all functions in flight and alternate builtin test methods are covered; evaluation logic and time-sharing of tests are discussed.