Flight Test Methods Research Articles

Study on Hot Weather Operation Flight Test of Civil Aircraft Fuel System

Abstract As one of the important on-board systems of an airplane, the fuel system of a civil aircraft mainly supplies fuel to the engine and APU of the aircraft, and it is necessary to ensure that the pressure and flow rate of the fuel supply meets the needs of the engine and APU under each kind of probable operating conditions. This paper studies the specific requirements of the airworthiness regulations for the fuel system to work under hot climate conditions. At the same time, it explores how to show the compliance of the fuel system with the airworthiness regulations through the flight test method, which covers the determination of the flight test parameters, the control of the hot fuel temperature, the test modification, the implementation of the flight test mission and the analysis of the results. This paper provides a reference for the implementation of the hot weather operation of similar types of aircraft fuel systems.

A dimensionality reduction approach in helicopter level flight performance testing

AbstractEvaluation of the power required in level flight is essential to any new or modified helicopter performance flight-testing effort. The conventional flight-test method is based on an overly simplification of the induced and profile power components required for a helicopter in level flight. This simplistic approach incorporates several drawbacks that not only make execution of flight sorties inefficient and time consuming, but also compromise the level of accuracy achieved. This paper proposes an alternative flight-test method for evaluating the level-flight performance of a conventional helicopter while addressing and rectifying all identified deficiencies of the conventional method. The proposed method, referred to as the corrected-variables screening using dimensionality reduction (CVSDR), uses an original list of 36 corrected variables derived from basic dimensional analysis principles. This list of 36 corrected variables is reduced using tools of dimensionality reduction to keep only the most effective level-flight predictors. The CVSDR method is demonstrated and tested in this paper using flight-test data from a MBB BO-105 helicopter. It is shown that the CVSDR method predicts the power required for level flight about 21% more accurately than the conventional method while reducing the required flight time by an estimate of at least 60%. Unlike the conventional method, the CVSDR is not bounded by the high-speed approximation associated with the induced power estimation, therefore it is also relevant to the low airspeed regime. This low-airspeed relevancy allows the CVSDR method to bridge between the level-flight regime and the hover. Although demonstrated in this paper for a specific type of helicopter, the CVSDR method is applicable for level-flight performance flight testing of any type of conventional helicopter.

Parameter Adjustment of the Autothrust Control Law by Flight Tests

Aiming at deriving precise automatic control of airspeed, a comprehensive flight test solution for autothrust control law parameter adjustment was proposed, providing a worthy reference for other aircraft types’ work, including the flight test points and methods, aircraft modification requirements for the test, the system performance analysis and evaluation methods, and the flight test risk analysis. A method invoking test scripts by a flight test interface unit and a control panel to configure parameters of autothrust control law and input excitation signals was proposed for the first time. The test scripts and the test devices cross-link can be fully verified in advance in both iron bird and real aircraft ground tests, which can greatly reduce unexpected risks of the test method based on the flight test engineers’ personal experience. According to the evaluation from three aspects of system response, robustness and throttle lever movement based on the test data, the optimal autothrust control law parameters configuration was given.

Research on the method of negative acceleration flight test based on simulation

Negative acceleration flight test is an important part of civil airplane to obtain airworthiness certification. According to civil aviation regulations, engine, APU, fuel, power supply and other systems need to be verified by flight test. In this paper, the factors affecting negative acceleration flight are evaluated by simulation method, such as gross weight (GW), center of gravity (CG), pitch attitude, calibrated airspeed, and stick travel. Relationship between flight path angle (FPA) rate and the calibrated airspeed is formed by approximate calculation. The duration of each negative acceleration flight is suggested to be a minimum of 7 seconds. In order to maximize the time to maintain the negative overload state, a more effective flight test method is proposed.

A Dimensionality Reduction Approach in Helicopter Hover Performance Flight Testing

The power required to hover a helicopter is fundamental to any new or modified performance flight-testing effort. The conventional method of relating two nondimensional variables (coefficients of power and weight) is overly simplified and neglects compressibility effects in the power required to hover under a wide range of gross weights and atmospheric conditions. An alternative flight-test method for assessing hover performance while addressing this deficiency of the conventional method is proposed. The method uses an original list of 15 corrected variables derived from fundamental dimensional analysis, which is further reduced by means of dimensionality reduction to include only the most essential and effective predictors. The method is demonstrated using data of a Bell Jet-Ranger and shows that at the 95% confidence level; the averaged prediction error is only 0.9 hp (0.3% of the maximum continuous power). Using the same data, the conventional method yields a much larger averaged prediction error of 1.7 hp.

System identification of cropped delta UAVs from flight test methods using particle Swarm-Optimisation-based estimation

AbstractIn the era of Unmanned Aerial Systems (UAS), an onboard autopilot occupies a prominent place and is inevitable for many of their modern applications. The efficacy of autopilot heavily relies upon the accuracy of the sensors employed and the capability of the onboard flight controller. In general, aerodynamic behaviour and flight dynamic capabilities of Unmanned Aerial Vehicles (UAVs) govern the selection and the design of flight controllers. Precise modeling of linear aerodynamic characteristics from flight data can be achieved using many of the existing classical parameter estimation techniques such as Output Error Method (OEM), Equation Error Method (EEM), and Filter Error Method (FEM). However, all the classical methods may not be readily applicable for aerodynamic modeling in nonlinear flight envelopes. The current manuscript is an attempt to exploit the capabilities of the Artificial Intelligence (AI) technique, named Particle Swarm Optimisation (PSO), in combination with Least Squares (LS) cost function to perform linear as well as nonlinear aerodynamic parameter estimation. The aforementioned task is accomplished by considering flight data from manoeuvers pertaining to linear angles of attack, moderate and near stall flight envelopes of two different UAVs with cropped delta planform geometry. Parameters estimated using the proposed LS-PSO method are consistent with minimum standard deviation and are on a par with OEM estimates. The proposed LS-PSO method enhances the capabilities of LS-based EEM while estimating stall characteristic parameters, which was not possible with LS alone. The longitudinal and lateral-directional static parameters estimated from the full-scale wind tunnel testing of the two UAVs were also used to corroborate the results obtained from the flight data using the LS-PSO method.

Research on UAV Flight Test Technology Based on Embedded Commands

This paper presents the current situation that the pilot’s manual operation method can’t fully meet the requirements of UAV flight test. To solve the problem, a UAV flight test technology based on embedded command is proposed. By analysing the verification contents and assessment methods of UAV flight test, the typical actions in UAV flight test are extracted, the embedded command modules are designed, and the embedded command flight test method is formulated. The flight test results show that compared with the manual operation method, the embedded command flight test technology can alleviate pilot’s operation burden, reduce the flight test risks and improve the flight test efficiency.

Design and Development of a High-Speed UAS for beyond Visual Line-of-Sight Operations

This paper provides details on best practices for vehicle design for BVLOS missions, with a focus on robust control links and real-time health monitoring.The development work presented herein focuses on a specific mission to set the speed and distance world record for a fully autonomous vehicle, requiring development of analytic tools for estimating vehicle range and endurance, as well as models for predicting command and control link integrity while operating BVLOS. Range and endurance estimates are developed using a novel real-time system identification algorithm onboard the vehicle for aerodynamic parameter estimation, and propulsion system characterization. Flight-testing methods and results are presented demonstrating the capabilities of these algorithms, with results from a unique test-case involving an unintended fault during the record flight. In addition to the vehicle performance assessment, Friis-based models are introduced to provide a pseudo real-time communications link range assessment. The predictive communications range models were refined using specific antenna and system characterization data measured in a compact range facility resulting in a detailed link budget useful for addressing BVLOS command and control requirements. The combination of aircraft and radio-frequency performance prediction models with supporting flight-testing techniques applied to small unmanned aerial systems provides critical insight into the design of future vehicles, and validation of existing systems tasked with performing BVLOS missions.

A Flight Load Test Method for Helicopter Rotor Blade

This paper studies a flight load test method for helicopter blade, including the structural force of helicopter blade, strain gauge bridge modification, load calibration method and the establishment of load calibration equation in the typical maneuvers. First, blade structure of the helicopter was analyzed and the strain gauges were pasted in the load profile which were considered to be important positions to analysis the flight safe of helicopter. Second, Load calibration equations were obtained by ground calibration test, which would be used to convert the strain time curve to the load time curve. The helicopter flighted in different maneuvers, such as hovering, horizontal flight and climbing. At the same time, flight parameters were recorded, such as helicopter altitude, pitch angle, roll angle and yaw Angle. The flight parameters can be used to analysis the flight conditions. Then the blade flight load in different flight maneuvers was analyzed. The result was that the time-domain waveform was neat, good periodicity, strong regularity, no hybrid interference and jump point. It was illustrated that the blade load test flight data were accurate and reliable, which met the accuracy requirements of Engineering. The blade load data obtained through different flight maneuvers can be used to structural modification, fatigue analysis and blade design.

Boundary Layer Transition Measurement Flight Test Based on Infrared Image Technique

According to the requirements of boundary layer transition detection measurement flight test, the study on the detection technique of infrared thermal image in real flight conditions was carried out. In this paper, the basic principle of infrared measurement technique was introduced firstly. Through the analysis of the main factors, such as emissivity, radiation incidence angle and temperature difference between the measured object and environment, the test means were determined, which includes surface treatment, infrared light path adjustment, solar radiation and internal heating of tested parts. Then the whole flight test method of boundary layer transition measurement was designed, including selecting test object and test instrument, determining test procedure and image processing procedure, and giving transition location criterion. Finally, the validity of this flight test method was demonstrated through testing results, the research of which would be helpful and meaningful in boundary layer transition measurement.

Внедрение летно-моделирующего комплекса и выбор летающей лаборатории с целью снижения издержек проведения испытаний авиационного оборудования

In the face of growing competition in high-tech products, the feasibility of introducing product and technological innovations into Russian enterprises is revealed. The article conducted a study of the Russian and foreign markets. The analysis of the state program of development of the military-industrial complex and the aviation industry. The main stages of development and production of further aviation aircraft are given. The scheme of testing new equipment and aircraft weapons was considered. The article determines the feasibility of using flight-model complexes for conducting summer tests of aircraft units and assemblies, its composition and main tasks. Based on the analysis of current flight test methods, the determining factors for the choice of a flying laboratory were identified in order to reduce the cost of conducting tests of aircraft equipment.

CИНТЕЗ РОБАСТНОГО ПІД-РЕГУЛЯТОРА ДЛЯ НЕВИЗНАЧЕНОЇ НЕЛІНІЙНОЇ МОДЕЛІ БПЛА З ВИКОРИСТАННЯМ ШТУЧНОЇ МОДЕЛІ ПОСИЛЕННЯ І ТИМЧАСОВОЇ ЗАТРИМКИ

Estimations of the flight vehicles aerodynamic coefficients through the theoretical, numerical, wind tunnel and flight-test methods have always errors and uncertainties. Nonlinear dynamics of the unmanned aerial vehicle due to speed variations, as well as variability and uncertainty of the aerodynamic coefficients can be considered as an uncertain model. A robust proportional-integral-derivative controller is designed based on the nonlinear optimization in the time domain for the uncertain nonlinear dynamical model of the unmanned aerial vehicle. Artificial gain and time-delay models are added to achieve required stability margins as system robustness during the robust proportional-integral-derivative control design. The aerodynamic coefficients are divided into two groups on the basis of the output vector's sensitivity to the aerodynamic coefficients. Nonlinear optimization of the criterion is performed in two steps for the robust proportional-integral-derivative controller design. In the first step of the design, nominal values are used for coefficients with low-sensitivity, and upper and lower limits are used only for high sensitive aerodynamic coefficients. In the second step, the upper and lower limits are applied all of the aerodynamic coefficients to evaluate and re-adjust the robust proportional-integral-derivative controller parameters. The robust controller is designed for the uncertain nonlinear roll and lateral dynamic model of the Skywalker X8 flying wing to show the effectiveness of the proposed method to guarantee the stability margins. In the given example during the design of the robust controller, with adding of the artificial gain and time-delay model in the control loop the controller parameters are changed. This regulator increases the phase stability margin by about 10 degrees and the gain stability margin by about two for the family of the equivalent uncertain linear models.

ON THE QUESTION OF SAFETY OPERATION OF A HELICOPTER AFTER CARGO DISCHARGE FROM EXTERNAL SLING

When helicopters fly with suspended load, there may be a need for a load drop. This can be associated with both the abnormal situations in flight and with a regular load drop to a predetermined area. However, under certain conditions, after the external load drop, there may be significant changes in the parameters of the helicopter movement. In this regard, when planning the helicopter flight with suspended load, it is necessary to make sure of safety in case of a load drop. In carrying out this work, theoretical and experimental research methods have been used. The theoretical methods of research include analytical calculations and computational experiments conducted with the help of special software. Experimental methods include flight test methods. Flight experiments allowed to confirm the adequacy of theoretical methods for researching the helicopter dynamics after external load drop. The proposed methodology for ensuring the safety of helicopter operation involves a staged research of the helicopter dynamics after external load drop. At the first stage, an analytical determination of the normal vertical load factor after the load drop according to the formula derived earlier was carried out. At the second stage of the research, computational experiments were performed using the HeliCargo software package, which allowed not only to clarify the values of normal vertical load factor obtained by the analytical calculations, but also to consider the movements of the helicopter mass center and its angular movements after load drop. The conducted research shows that the helicopter dynamics after load drop can be satisfactorily investigated by the previously developed theoretical methods, which is confirmed by comparing the obtained data with the results of the flight experiments. In addition, this paper presents concrete results, stating the safety of external load drop, which can be used in planning the production of work on the delivery of suspended load to a predetermined area.

A Flight Test Method of Flammable Liquid Leakage and Discharge for Certain Aircraft

To guarantee the successful completion of airworthiness certification test for a certain type of aircraft, the leakage and discharge compliance criteria of flammable liquid were designed by studying airworthiness requirements and the design characteristic of the civil aircraft, where flammable liquid leakage zones and hazardous zones in which liquids are not allowed to enter were determined. Further the test equipments and the dyeing liquor for replacing flammable fluid were developed, and ground and flight test methods were proposed. Moreover, the above test equipment and methods were applied to flight tests of flammable liquid leakage and discharge for certain civil aircraft. The test results show that the test equipment is stable and reliable, and the flight test method can accurately demonstrate the flammable liquid leakage and discharge.

Developments of Aerodynamic Flight Test and Scale Effect Compensation Methods toward Aerodynamic Prediction

Developments of Aerodynamic Flight Test and Scale Effect Compensation Methods toward Aerodynamic Prediction

Non-linear aerodynamic modelling of unmanned cropped delta configuration from experimental data

ABSTRACTThe paper presents the aerodynamic characterization of a low-speed unmanned aerial vehicle, with cropped delta planform and rectangular cross section, at and around high angles-of-attack using flight test methods. Since the linear models used for identification from flight data at low and moderate angles of attack become unsuitable for accurate parameter estimation at high angles of attack, a non-linear aerodynamic model has to be considered. Therefore, the Kirchhoff's flow separation model was used to incorporate the non-linearity in the aerodynamic model in terms of flow separation point and stall characteristic parameters. The Maximum Likelihood (ML) and Neural Gauss-Newton (NGN) methods were used to perform the parameter estimation on one set of low angle-of-attack and one set of near-stall flight data. It is evident from the estimates that the NGN method, which does not involve solving equations of motion, performs on a par with the classical ML method. This may be attributed to the reason that NGN method uses a neural network which has been trained by performing point to point mapping of the measured flight data. This feature of NGN method enhances its application over a wider envelope of high angles of attack flight data.

Power-Required Flight Testing of Unmanned Aerial Systems Using Non-Steady-State Flight

Power-required flight testing has traditionally been done using the classic power independent of weight, velocity independent of weight method. This method is well suited for manned aircraft but has some shortcomings when applied to unmanned systems that are remotely piloted and/or operating on small test ranges. Maximum-likelihood estimation was used to identify the lift and drag models of two test aircraft using data acquired through acceleration and deceleration runs. The calculated power-required data were reduced using the classic technique to remove weight and air density effects. Flight-test data were collected using a Raspberry Pi single-board computer because of its low cost, weight, and good performance. Data points using the classic flight-test method were also collected for comparison against the dynamic method. Because of the small test range used, a third of all points collected using the classic method had to be eliminated because of assumption violations. The flight-test results show good agreement between the deceleration runs and the classic method. The acceleration runs generally had a poorer overall fit and a larger spread due to excessive sensor noise caused by imbalances in the motor/prop combination. Overall, the dynamic power analysis was much easier to perform and required significantly less time and airspace.

Stability characteristics of the single-engine tractor propeller airplane in climb

This paper presents the analysis of effects on the airplane longitudinal stability reserve by the single tractor propeller at high engine power settings. The analysis was done within the terminal flight phases. The airplane has the original and modified horizontal tail shape. The horizontal tail modifications were applied to the second prototype of the “LASTA” airplane during the development of flight testing. The airplane stability problem was initiated by the rise of pilot average mass diverging in practice from the specification of 86 kg (190 lbs), as defined in aviation regulations and installation of measurement equipment. The “LASTA” airplane model is here analyzed for two different horizontal tail shapes by the combination of the calculating aerodynamics and flight testing methods. The paper makes the presentation of the introduction phase of the airplane design and construction process.

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.