Experimental Flight Test Research Articles

Experimental Comparison of Nonlinear Guidance Laws for Unmanned Aircraft

Two aircraft guidance algorithms from the literature, the Non-Linear Guidance Law (NLGL) and the Nonlinear Differential Geometric Path-Following Guidance Law (NDGPFG), are assessed using a 25 kg unmanned aerobatic aircraft. The paper provides experimental results of the first flight test with the NDGPF. Prior to the real world application, a simulation study is performed to mitigate the risk. For both guidance laws, tracking performance is investigated on a purely kinematic model with unlimited control bandwidth first. Both laws provide superb tracking and the NDGPFG achieves exact path-following, i. e. zero track error. In a second simulation, a high-fidelity model of the aircraft is used. This model includes parasitic dynamics and hence has a finite control bandwidth. In contrast to the results with the purely kinematic model, neither of the algorithms achieves exact path-following with the high-fidelity model. Nevertheless, both guidance laws provide good tracking performance and prove feasible in varying environmental conditions. This observation is finally confirmed in the flight test experiment.

Technical Measures Perspectives in Selection of Handling Qualities Criteria for Flight Control Design

AbstractThe requirements of satisfactory pilot‐vehicle interaction in high workload precision flying tasks plays a very important and dominant role in the design of flight control systems. Based on the extensive flight test experience, quantitative handling qualities specifications to aid control law design have been derived and documented in relevant MIL Standards. In this paper an attempt is made to demonstrate methodology used for validation of Handling Qualities (HQ) requirements for a HighPerformance fighter Aircraft (HPA) using technical measurements in systems engineering i.e. Measure of effectiveness (MOE), Measure of performance (MOP) and Technical performance measure (TPM) as defined by INCOSE.A mission requirement such as Air to Air Refueling (AAR) is considered as a user need in qualitative form and is mapped to MOE as an indicator of achieving the operational objective. The MOPs are arrived from the handling quality specifications from MIL standards for the flight control design. TPMs enable an assessment of the system design by estimating the values of key performance parameters of the design through engineering analyses and tests.A piloted evaluation is performed for the given mission requirement. The flight data together with pilot comments on aircraft handling qualities are assessed to identify any mismatch between predicted HQ versus piloted evaluation which leads to iteration of flight control laws. Thereby, the use of technical measures keeps track of requirements to be met while maintain desired system performance as well anticipate any risk and mitigate the same during early stage of design cycle.

Backstepping approach for design of PID controller with guaranteed performance for micro-air UAV

Flight controllers for micro-air UAVs are generally designed using proportional-integral-derivative (PID) methods, where the tuning of gains is difficult and time-consuming, and performance is not guaranteed. In this paper, we develop a rigorous method based on the sliding mode analysis and nonlinear backstepping to design a PID controller with guaranteed performance. This technique provides the structure and gains for the PID controller, such that a robust and fast response of the UAV (unmanned aerial vehicle) for trajectory tracking is achieved. First, the second-order sliding variable errors are used in a rigorous nonlinear backstepping design to obtain guaranteed performance for the nonlinear UAV dynamics. Then, using a small angle approximation and rigorous geometric manipulations, this nonlinear design is converted into a PID controller whose structure is naturally determined through the backstepping procedure. PID gains that guarantee robust UAV performance are finally computed from the sliding mode gains and from stabilizing gains for tracking error dynamics. We prove that the desired Euler angles of the inner attitude controller loop are related to the dynamics of the outer backstepping tracker loop by inverse kinematics, which provides a seamless connection with existing built-in UAV attitude controllers. We implement the proposed method on actual UAV, and experimental flight tests prove the validity of these algorithms. It is seen that our PID design procedure yields tighter UAV performance than an existing popular PID control technique.

Airborne Visual Detection and Tracking of Cooperative UAVs Exploiting Deep Learning.

The performance achievable by using Unmanned Aerial Vehicles (UAVs) for a large variety of civil and military applications, as well as the extent of applicable mission scenarios, can significantly benefit from the exploitation of formations of vehicles able to fly in a coordinated manner (swarms). In this respect, visual cameras represent a key instrument to enable coordination by giving each UAV the capability to visually monitor the other members of the formation. Hence, a related technological challenge is the development of robust solutions to detect and track cooperative targets through a sequence of frames. In this framework, this paper proposes an innovative approach to carry out this task based on deep learning. Specifically, the You Only Look Once (YOLO) object detection system is integrated within an original processing architecture in which the machine-vision algorithms are aided by navigation hints available thanks to the cooperative nature of the formation. An experimental flight test campaign, involving formations of two multirotor UAVs, is conducted to collect a database of images suitable to assess the performance of the proposed approach. Results demonstrate high-level accuracy, and robustness against challenging conditions in terms of illumination, background and target-range variability.

Unsteady Detached-Eddy Simulation (DES) of the Jetstream 31 aircraft in One Engine Inoperative (OEI) condition with propeller modelling

Numerical results from a three-dimensional (3D) computational fluid dynamics (CFD) model of the Jetstream 31 aircraft in conditions of one engine inoperative are presented. The objective of this work is to analyse the performance of the Jetstream 31 aircraft and provide transient data using an unsteady Detached-Eddy Simulation (DES) CFD approach and a numerical propeller model to compare computational results with a single engine flight test experiment. The propeller modelling approach has been implemented through User-Defined Functions using C programming language to replicate the propeller effect. Different angles of attack and sideslip are studied, based on records from flight test data, both with unsteady (DES) and steady-state Reynolds-Averaged Navier-Stokes (RANS) models. An error analysis on the flight test data provides an error band from 2% to 16% among all cases, high values due to the lack of many data samples. Across the RANS approach, an average deviation of 6.9% and 3.8% for respectively lift and drag coefficients is achieved. By applying the DES turbulence modelling approach, a better lift prediction is achieved (5.4%) despite a slightly worse drag (4.5%). It has also been found that 80% of the numerical results are within the error band defined. A close agreement has been found within moment coefficients, with average percentage deviations from 3.3% to 7.0%. Overall, an analysis has been carried out in the present work, both on the flight test and computational sides to provide reliable numerical results of these aerodynamic properties.

Assessment of a small UAV speed polar graph by conducting flight tests

Purpose The purpose of this paper is to present an approach to a polar graph measurement by a flight testing technique and to propose a baseline research method for future tests of UAV polar graphs. The method presented can be used to demonstrate a conceptual and preliminary design process using a scaled, unmanned configuration. This shows how results of experimental flight tests using a scaled flying airframe may be described and analysed before manufacturing the full scale aircraft. Design/methodology/approach During the research, the flight tests were conducted for two aerodynamic configurations of a small UAV. This allowed the investigation of the influence of winglets and classic vertical stabilizers on the platform stability, performance and therefore polar graphs of a small unmanned aircraft. Findings A methodology of flight tests for the assessment of a small UAV’s polar graph has been proposed, performed and assessed. Two aerodynamic configurations were tested, and it was found that directional stability had a large influence on the UAV’s performance. A correlation between the speed and inclination of the altitude graph was found – i.e. the higher the flight speed, the steeper the altitude graph (higher descent speed, steeper flight path angle). This could be considered as a basic verification that the recorded data have a physical sense. Practical implications The polar graph and therefore glide ratio of the aircraft is a major factor for determining its performance and power required for flight. Using the right flight test procedure can speed-up the process of measuring glide ratio, making it easier, faster, robust, more effective and accurate in future research of novel, especially unorthodox configurations. This paper also can be useful for the proper selection of requirements and preliminary design parameters for making the design process more economically effective. Originality/value This paper presents a very efficient method of assessing the design parameters of UAVs, especially the polar graph, in an early stage of the design process. Aircraft designers and producers have been widely performing flight testing for years. However, these procedures and practical customs are usually not wide spread and very often are treated as the company’s “know how”. Results presented in this paper are original, relatively easily be repeated and checked. They may be used either by professionals, highly motivated individuals and representatives of small companies or also by ambitious amateurs.

Design and Flight Evaluation of Primary Control System for Learjet-25B Aircraft

In this paper, a multi-input/multi-output servomechanism primary flight controller is presented. Also discussed are flight data related to the system performance and pilot assessments of handling qualities. The data were collected during flight-test experiments (March 2018) on the Calspan Variable Stability In-Flight Simulator Learjet-25B aircraft. Aircraft body angular rates and linear accelerations are used to enable decoupled command tracking in pitch, roll, and yaw axes. Angle of attack and angle of sideslip measurements are not required for the system to operate. The developed primary flight controller translates pilot controls into the corresponding pitch, roll, and yaw commands for the aircraft to track. The system consists of a gain-scheduled robust linear baseline augmented with a direct adaptive model reference output feedback. The baseline is designed via the Observer-Based Loop Transfer Recovery method. The adaptive component incorporates the closed-loop reference model, whereby a state observer from the baseline Observer-Based Loop Transfer Recovery system also acts as the reference model for the adaptive augmentation. In March 2018, the Observer-Based Loop Transfer Recovery controller was successfully flight verified on the Calspan Learjet-25B aircraft, at the U.S. Air Force Test Pilot School, Edwards Air Force Base, California. Flight-test results are discussed, and a brief summary is given at the end.

Dynamic Modeling, Control System Design and MIL–HIL Tests of an Unmanned Rotorcraft Using Novel Low-Cost Flight Control System

Unmanned helicopters have gained great importance during recent years due to their special abilities such as hover flight, vertical take-off and landing, maneuverability and superior agility. The advances in electronic devices technologies lead to more powerful and lighter processors to be used in avionic systems which have attracted more attention to these UAVs. The first steps of utilizing an unmanned helicopter are dynamic modeling, control system design and performing model-in-the-loop (MIL) and hardware-in-the-loop (HIL) tests which are presented in this paper. In this research, MIL and HIL tests of an unmanned helicopter are done using novel Linux-based flight control system built on Raspberry Pi board (different from normally used PC-104 and STM- or Arduino-based systems). Dynamic modeling, robust hierarchical control design, flight control system hardware and software architecture and MIL and HIL test results are reported here. By succeeding in these tests, it is shown that the proposed platform can be used in experimental flight tests in next steps.

Robust Attitude Control for Quadrotors Based on Parameter Optimization of a Nonlinear Disturbance Observer

This paper presents a nonlinear disturbance observer (NDOB) for active disturbance rejection in the attitude control loop for quadrotors. An optimization framework is developed for tuning the parameter in the NDOB structure, which includes the infinity-norm minimization of the weighted sum of noise-to-output transfer function and load disturbance sensitivity function. Subsequently, the minimization generates an optimal value of the parameter based on the tradeoff between disturbance rejection and noise propagation in the system. The proposed structure is implemented on PIXHAWK, a real-time embedded flight control unit. Simulation tests are carried out on a custom built, high-fidelity simulator providing physically accurate simulations. Furthermore, experimental flight tests are conducted to demonstrate the performance of the proposed approach. The system is injected with step, sinusoidal, and square wave disturbances, and the corresponding system tracking performance is recorded. Experimental results show that the proposed algorithm attenuates the disturbances better compared to just a baseline controller implementation. The proposed algorithm is computationally cheap, an active disturbance rejection technique and robust to exogenous disturbances.

Structural design and verification of an innovative whole adaptive variable camber wing

A whole adaptive variable camber wing (AVCW) equipped with an innovative double rib sheet (DRS) structure is experimentally and numerically studied in this work. The DRS structure adopts the surface contact mode for the force transmission of changeable camber wing instead of the conventional rigid hinge joint contact. The whole AVCW design allows to adjust the shape of wing in a real-time at various flight conditions, which is of great interest for Unmanned Aerial Vehicle (UAV) applications. The flight-test experiments demonstrate that the total AVCW carrying the developed adaptive control system (ACS) can enhance UAV flight efficiency by 14.1% comparing to a traditional fixed-wing of Talon UAV. In addition, it indicates that employing the whole AVCW structure can sustain a larger flight load, without increasing the weight of entire wing structure except for the actuator device and adhesives, which is promising for future engineering applications.

Store Separation: Photogrammetric Solution for the Static Ejection Test

The process of developing and certifying aircraft and aeronautical systems requires the execution of experimental flight test campaigns to determine the actual characteristics of the system being developed and/or validated. In this process, there are many campaigns that are inherently dangerous, such as the store separation. In this particular case, the greatest risk is the collision of the store with the fuselage of the aircraft. To mitigate the risks of this campaign, it is necessary to compare the actual trajectory of a separation with its simulated estimates. With such information, it is possible to decide whether the next store release can be done with the required safety and/or whether the model used to estimate the separation trajectory is valid or not. Consequently, exact determination of the trajectory of the separation is necessary. Store separation is a strategic, relevant, and complex process for all nations. The two main techniques for determining the quantitative store trajectory data with 6DoF (six degrees of freedom) are photogrammetry and instrumented telemetry packages (data obtained from inertial sensors that are installed in the store). Each presents advantages and disadvantages. In regard to photogrammetry, several market solutions can be used to perform these tests. However, the result of the separation trajectory is only obtained after the test flight, and therefore, it is not possible to safely carry out more than one on the same flight. In this context, the development and validation of a solution that will allow the realization of near real-time separation analysis are in fact an innovative and original work. This paper discusses the development and validation, through actual static ejection tests, of the components that will compose a new onboard optical trajectory system for use in store separation campaigns. This solution includes the implementation of a three-dimensional (3D) calibration field that allows calibration of the optical assembly with just one photo per optical assembly, development of a complete analytical model for camera calibration, and development of specific software for identification and tracking of targets in two-dimensional (2D) coordinate images and three-dimensional (3D) coordinate trajectory calculation. In relation to the calibration, the analytical model is based on a pinhole type camera and considers its intrinsic parameters. This allowed for a mean square error smaller than ±3.9 pixels @1σ. The 3D analysis software for 6DoF trajectory expression was developed using photogrammetry techniques and absolute orientation. The uncertainty associated with the position measurement of each of the markers varies from ±0.02 mm to ±8.00 mm @1σ, depending on the geometry of the viewing angles. The experiments were carried out at IPEV (Flight Test Research Institute)/Brazil, and the results were considered satisfactory. We advocate that the knowledge gained through this research contributes to the development of new methods that permit almost real-time analysis in store separation tests.

Effects of Attitude on HIFiRE-5b Boundary-Layer Transition

HIFiRE-5 is a hypersonic flight-test experiment designed to investigate the aerothermodynamics of a three-dimensional geometry. The vehicle is an elliptic cone with a 2:1 aspect ratio and 2.5 mm nose radius. The HIFiRE-5b achieved a Mach number of 7.7 to 7.9 and freestream unit Reynolds number of to during descent. Heat flux was calculated from thermocouple pairs, and boundary-layer transition locations were identified from these heating rates. A multilobed transition front was observed; it is suspected that a different mechanism is responsible for each of these lobes. The first round of flight-test data analysis did not control for attitude variation, resulting in oscillating heat-flux rates. The current work subsamples the data based on angle of attack and yaw, enabling a more precise investigation of other relevant variables. Regions of the flowfield especially sensitive to attitude variation can be better analyzed. The heating rate at the leading edge exhibits the expected dependence on yaw angle and insensitivity to angle of attack. Data from the crossflow-dominated portion of the flowfield show the region’s sensitivity to angle of attack and dependence on yaw. Transition near the centerline exhibited a small dependence on angle of attack, but a large effect from yaw.

Investigation on correlation between wind tunnel and flight for boundary layer stability and transition of MF-1 blunt cone

Hypersonic boundary-layer transition (BLT) is affected by multifarious factors, some of which cannot be simulated in ground wind tunnels, which makes the BLT data of wind tunnel experiments are usually different to that of flight experiments. In order to convert BLT data from wind tunnels to real flight cases, and to improve BLT prediction accuracy, a blunt cone model with zero angle-of-attack and 7° semi-cone angle is selected to contrastive analysis the boundary-layer stability and transition between wind tunnels and flights by using linear stability theory and eN method, based on the data set of our MF-1 flight test and hypersonic wind tunnel experiments, as well as an abroad flight test and wind tunnel experiments. The results show that the difference of stability and transition between wind tunnels and flights is essential. The wall temperature ratio in flights is smaller than that in wind tunnels, which leads to the second mode frequency in flight tests are greater than that of wind tunnel experiments. Two empirical formulas are developed to compute the most unstable frequency of the second mode. These formulas are shown suitable for both flight conditions and wind tunnel conditions. It is also shown that the N value ( N T) at the start of BLT is greatly affected by the nose Reynolds number. The larger the nose Reynolds number, the smaller the N T value. When the nose-tip bluntness is large enough, the BLT occurs even without boundary-layer mode instability. The front bow shock waves of blunt bodies will induce entropy layers, and the bluntness will affect the location (swallow point) where the entropy layers merge into the boundary layers. The relative location of transition point and swallow point is shown in good correlation with the N T, and the trend is similar for both wind tunnels and flights. A correlation between the transition-swallow relative location and the N T is developed by the least square method. This correlation is expected to conveniently predict hypersonic BLT of similar conditions.

Design to Validation of a New Heat Flux Source with Uncertainty Analysis

High-speed, transient aerothermal studies produce excessive temperatures and heat fluxes in laboratory, ground, and flight-test experiments. Experiments are performed for 1) understanding the response of test vehicles under various induced thermal environments, 2) verifying computational codes and input parameters, and 3) evaluating thermophysical and mechanical characteristics of new materials. In addition to these rationales, the reconstruction of the surface heat flux from in-depth or backside instrumentation is an important inverse application. Recent research has led to an alternative view of inverse heat conduction based on calibration principles performed in the frequency domain leading to a novel “parameter-free” inverse heat conduction measurement equation. The final mathematical framework reveals that the resolution of a first-kind Volterra integral equation for the surface heat flux prediction contains only experimental data sets. This type of mathematical formulation is highly ill posed. This paper focuses on the design, fabrication, and preliminary test campaign of a new electrical heating cell for producing a verifiable heat flux source. The feasibility investigation demonstrates the merits of system through modeling, experimental design, and transient uncertainty propagation. Validation studies are presented, leading to favorable outcomes. Slug calorimetry, composed of a pure copper, is used in the validation campaign.

HIFIRE-5b Heat Flux and Boundary-Layer Transition

Hypersonic International Flight Research Experimentation (HIFiRE)-5 is a hypersonic flight-test experiment designed to investigate the aerothermodynamics of a three-dimensional geometry. The vehicle is a 7 deg minor-axis half-angle elliptic cone with a aspect ratio and 2.5 mm nose radius. The HIFiRE-5b achieved a Mach number of 7.7–7.9 and a freestream unit Reynolds number of during descent. Heat flux was calculated from thermocouple pairs, and boundary-layer transition locations were identified from these heating rates. A multilobed transition front was observed, with earliest transition at the centerline, on the leading edges, and part way in between. It is suspected that a different mechanism is responsible for each of these lobes. Transition along HIFiRE-5b’s leading edges occurred at roughly constant length-based Reynolds numbers, suggesting a different mechanism than the roughness-induced transition suspected for HIFIRE-5a. Transition near the planes of symmetry was found to be less sensitive to attitude variation than elsewhere on the vehicle surface. In particular, hot streaks suspected to be due to stationary crossflow vortices were very sensitive to small changes in attitude (roughly angle of attack and yaw). Data from 128 pairs of thermocouples show the overall temperature, heat flux, and boundary-layer state during flight.

Flight-testing of a cooperative UGV-to-UAV strategy for improved positioning in challenging GNSS environments

We present an experimental flight test evaluation of a cooperative navigation strategy in which an Unmanned Aerial Vehicle (UAV) that is subjected to very poor GNSS satellite geometry is provided ranging updates from Unmanned Ground Vehicle (UGV). Central to the design of this approach, the UGV's motion planning is designed to provide the most favorable positioning geometry for the UAV. During a set of field tests, the positioning error of a UAV that is confronted with unfavorable GNSS satellite geometry is shown to be reduced by more than five-fold through the use of ranging updates from a UGV.

Ground-Based Sense-and-Avoid System for Small Unmanned Aircraft

In this Paper, we present a complete, proof-of-concept sense-and-avoid solution for small unmanned aircraft systems, including a small low-cost ground-based radar system, multitarget tracking and estimation, collision detection, and an avoidance planner. We describe the development of a small frequency-modulated continuous-wave phased-array radar system that provides a three-dimensional surveillance volume. The radar measurements are processed using the recursive random sample consensus algorithm, producing tracks for the intruders and the ownship. We propose a collision-detection algorithm based on the geometric relationship between encountering aircraft. If a collision threat is detected, a collision-free new path is generated for the ownship using a two-step path-planning algorithm. In the first step, an initial suboptimal path is generated using an search. In the second step, the path is refined using a variant of the potential fields technique, adapted to the sense-and-avoid scenario. The performance of the complete system is demonstrated with flight-test experiments.

Adaptive Control of Precision Guided Airdrop Systems with Highly Uncertain Dynamics

The bulk of research in the field of precision guided airdrop systems has focused on improving landing accuracy in the presence of atmospheric winds that can exceed vehicle airspeed. One important challenge of parafoil systems is their highly uncertain flight dynamic behavior and control response, which can result from canopy degradation or an offnominal inflation event. This significantly impacts the ability to reach the target and can often lead to very large miss distances. This work addresses guided airdrop system model uncertainty with a novel combined direct and indirect adaptive control strategy to quickly characterize vehicle dynamics and lateral control sensitivity in flight. Extensive simulation and experimental flight testing indicate that the proposed adaptive algorithm is capable of high-accuracy landing in a large variety of degraded conditions, including unknown nonlinear changes in control sensitivity as well as control reversals. In comparison, current industry standard algorithms experience over an order of magnitude decrease in accuracy when tested under identical scenarios.

Global Positioning System Denied Navigation of Autonomous Parafoil Systems Using Beacon Measurements From a Single Location

Precision-guided airdrop systems have shown considerable accuracy improvements over more widely used unguided systems through high-quality position, velocity, and time feedback provided by global positioning system (GPS). These systems, like many autonomous vehicles, have become solely dependent on GPS to conduct mission operations. This necessity makes airdrop systems susceptible to GPS blackout in mountainous or urban terrain due to multipathing issues or from signal jamming in active military zones. This work overcomes loss of GPS through an analysis of guidance, navigation and control (GNC) capabilities using a single radio frequency (RF) beacon located at the target. Such a device can be deployed at the target by ground crew on site to retrieve package delivery. Two novel GNC algorithms are presented, which use either range from or direction to a RF beacon. Simulation and experimental flight testing results indicated that beacon-based methods can achieve similar results as GPS-based methods. This technology provides a simple and elegant solution to GPS blackout with best method studied showing only a 21% decrease in landing accuracy in comparison to GPS-based methods.

Full control of a quadrotor using parameter-scheduled backstepping method: implementation and experimental tests

This paper proposes a nonlinear controller for a quadrotor helicopter, under the control of which the system is globally asymptotically stabilized with good control quality. The proposed controller is synthesized by Command Filtered Backstepping method with a novel parameter scheduling scheme. By scheduling controller parameters within the system-stabilizing region, the convergence speeds of errors in each step are adaptively adjusted based on different flight conditions. Amplitudes of control signals are reduced during fast tracking progress to avoid actuator saturation, which is hardly modeled and may cause instability. The controller also allows more aggressive tuning that achieves better regulation accuracy. The technology to implement the proposed controller is illustrated in detail, and key parameters of quadrotor model and the actuator’s dynamics are identified by experiments. To validate the proposed method, experimental flight tests are conducted under three typical flight conditions. Results comparing to other methods such as PID, sliding mode control and dynamic surface control are demonstrated, showing that the proposed controller is practical to an actual quadrotor system and can achieve good control performance.