Flying Wing UAV Research Articles

Micro UAVs with Fixed Wings: Design, Technological Solutions, and Tests

Considering the advantages of using expanded polystyrene (EPS) reinforced with adhesive tape made of glass fibers, this paper presents a design and technological solution for a functional drone of class C1, meaning a maximum take-off mass of 900 g, and tests validating the use of EPS for small UAVs under flight conditions. The selected profile was MH-49, which had a maximum chord thickness of 10.5%. This profile demonstrated a much lower coefficient of pitching moment than that of the NACA 63215 profile, giving this flying-wing UAV superior governability. This airfoil implies a geometry with greater attenuation of the trailing edge, and the design favors the placement of stress concentrators towards the trailing edge. Due to the use of fiberglass tape reinforcement technology, it is possible to address this profile, implying improved aerodynamic performance. The use of EPS in the disposable UAV industry may bring significant benefits, contributing to the development of high-performance, versatile, and low-cost UAVs, suitable for a wide range of tactical and logistic applications. This study presents the design, the fabrication, and testing of this drone, highlighting the advantages and possible challenges of the proposed solution.

Aerodynamic Assessment of Flying Wing UAV and Impact of Dimples on its Performance

Aerodynamic Assessment of Flying Wing UAV and Impact of Dimples on its Performance

Stealth Unmanned Aerial Vehicle Penetration Efficiency Optimization Based on Radar Detection Probability Model

Aerodynamic/stealth optimization is a key issue during the design of a stealth UAV. Balancing the weight of different incident angles of the RCS and combining stealth characteristics with aerodynamic characteristics are hotspots of aerodynamic/stealth optimization. To address this issue, this paper introduces a radar detection probability model to solve the weight balance problem of incident angles of the RCS and a penetration efficiency model to transfer the multi-object optimization into single-objective optimization. In this paper, a parameterized model of a flying-wing UAV is selected as the research object. A gradient-free optimization algorithm based on the genetic algorithm is used for maximizing efficiency. The optimization model balances the influence of the RCS mean value and RCS peak value on stealth performance. Moreover, the model achieves an optimal entire life cycle penetration efficiency coefficient by balancing aerodynamic and stealth optimization. The results show that the optimized model improves the penetration efficiency coefficient by 13.84% and increases maximum flight sorties by 1.8%. These results prove that the model has a reasonable combination of aerodynamic and stealth optimization for UAVs undertaking penetration missions.

Event-Triggered Adaptive Fuzzy Fault-Tolerant Attitude Control for Tailless Flying-Wing UAV With Fixed-Time Convergence

Event-Triggered Adaptive Fuzzy Fault-Tolerant Attitude Control for Tailless Flying-Wing UAV With Fixed-Time Convergence

Design and Analysis of a Micro Class UAV Integrating Zimmerman Configuration

This document describes the design and analysis processes followed by team to build a micro class UAV. Through this opportunity our team has gotten a chance to test and enhance our skills by developing a design and fabricate flying wing UAV of Zimmerman configuration. Integrating this flying wing into the blended body micro classed UAV, allowed the team members to push limits further and work extra hard to make the ends meet, with the final design of a lightweight UAV following all the constrains for the SAE ADC 2020.

Longitudinal Attitude Control and Stability Analysis for a Low Aspect Ratio Flying Wing UAV at High Angle of Attack

Longitudinal Attitude Control and Stability Analysis for a Low Aspect Ratio Flying Wing UAV at High Angle of Attack

Uçan Kanat Tipi İHA'larda Kanat Profillerinin Aerodinamik Performans Karşılaştırması

The aim of the study is to investigate how the choice of airfoil affects the aerodynamic characteristics of a flying wing UAV. For this purpose, comparative analyzes were performed for four different airfoils: MH60, TL54, Eppler 339, and TsAGI 12%. Given the maximum range performance (maximum lift /drag ratio), the best aerodynamic efficiency is given by the flying wing UAV with MH60 and TL54 airfoil. Based on their maximum lift-to-drag ratio, the flying wing UAVs made with MH60 and TL54 airfoils exhibited the best aerodynamic efficiency. Specifically, the maximum lift-to-drag ratio for the flying wing with the MH60 airfoil was 33.1, while that for the flying wing with the TL54 airfoil was 32.7. Considering the pitching moment coefficient, the flying wing made with the MH60 airfoil and TsAGI 12% exhibited a more stable characteristic than the TL54 and Eppler 339 airfoils. Based on the results of the study, it was found that the flying wing UAVs made with the TL54 and MH60 airfoils outperformed those made with the Eppler 339 and TsAGI 12% airfoils in terms of maximum range, minimum descent rate, and maximum endurance performance.

Adaptive fractional order non-singular terminal sliding mode anti-disturbance control for advanced layout carrier-based UAV

Advanced layout carrier-based UAV has grabbed increasing attention in aerospace field for its unique characteristics such as fewer casualties, lower labor costs and more capable of carrying out various tasks in shipping environment. However, the complex body structure of advanced layout UAV can bring about insufficiency in terms of stability, making it especially susceptible to changeable working circumstances. To overcome the object complexity and external disturbances mainly consisting of air turbulence that the UAV may come across, a composite adaptive fractional-order non-singular terminal sliding mode control strategy is proposed to achieve finite-time attitude tracking control and disturbance rejection. The distinctive features of the proposed strategy are embodied in active disturbance compensation, finite-time convergence, higher control accuracy and chattering suppression. The mathematical model of flying wing UAV is established first, then a fixed-time disturbance observer applying higher-order sliding mode is designed to obtain the estimation of unknown disturbance so that the carrier-based environment factors can be considered in the later control process. Next, an adaptive fractional-order non-singular terminal sliding mode controller is developed based on fractional order calculus and adaptive double power exponential reaching law to realize finite-time tracking control of advanced layout carrier UAV with great robustness against external disturbances due to the estimation from the proposed observer, and the stability analysis of closed-loop system is implemented by utilizing Lyapunov theory. Finally, a comparison study of the proposed strategy with other sliding mode control methods is presented by numerical simulation, to illustrate its feasibility and superiority.

Study on tolerance control of flying-wing layout vectored thrust UAV

The flying-wing layout vectored thrust UAV often uses multiple sets of aerodynamic rudder surfaces and vector nozzles to jointly generate the required control force and torque, which belongs to a typical overdrive system. Multiple sets of rudder surfaces give UAV a higher redundancy configuration, but also result in multiplying the rudder surface failure rate. Especially for the statically unstable flying-wing UAV in lateral direction, the moment imbalance caused by the failure of the executive rudder will seriously affect the flight safety. Therefore, this paper takes the failure of elevator and rudder damage as an example, and proposes a kind of control distribution method of aerodynamic-based control and vectored thrust as auxiliary. The binary vector deflection mechanism is used to generate vector thrust, compensate damaged actuator mechanism rudder effect, release the control margin of the rudder actuator, and restore the maneuverability of the aircraft. Under the failure of the aerodynamic rudder actuator damage, the stable control UAV is realized.

Simulation Analysis of Detection Area for Flying Wing Layout UAV Using X-Band Bistatic Radar

The shape design of flying wing UAV can effectively reduce the detection and tracking probability of single-station radar. Bistatic radar has the advantage of anti-stealth for stealth targets due to the characteristics of multi-station distribution. However, there is no literature to study the detection area of flying wing UAV by bistatic radar. To solve this problem, FEKO electromagnetic simulation software is used to establish the X-band electromagnetic model of X-47B UAV, and the bistatic RCS of 0° pitch angle is simulated. According to the bistatic radar equation, the effective detection area of X-47B UAV lateral flight is simulated. The results show that the detection range of bistatic radar is mainly in the vicinity of 5 banded regions and bistatic radar. The five strip regions in the effective detection range correspond to the five bright lines of bistatic RCS one by one. When the baseline of bistatic radar is 20 km, the effective detection range is the largest. The research provides data support for the layout optimization of bistatic radar station and the trajectory optimization of flying wing UAV.

Analysis of Stability Margin of Dynamic Inverse Control Law for Flying Wing UAV

The longitudinal and directional stability of flying wing UAV is poor, the coupling of triaxial force and torque is serious, and the control surface is special, so the flight control is difficult and the control law is complex. Taking a flying wing UAV as an example, the control laws of track tracking based on dynamic inverse control method and control allocation are designed. The applicability of traditional stability margin excitation method based on signal generator and pilot’s manual control is analyzed, and a stability margin excitation method based on three-a2xis control decoupling is proposed, which is verified by simulation. The results show that the stability margin of the designed three-axis dynamic inverse control law system meets the requirements, and the stability margin excitation method based on three-axis decoupling is suitable for the stability margin analysis of three-axis coupled UAV.

Fault tolerant control of uav with wing layout based on control allocation

As the flying wing layout unmanned aerial vehicle (uav) extensive research and task environment increasingly complex, Yu Feiyi layout unmanned aerial vehicle (uav) for fault tolerant control gradually become the main technical means of the flight control, using the established mathematical model of the flying wing uav longitudinal layout setting the actuator failure effect, is in the nature of adaptive control allocation fault-tolerant algorithm is given, and MATLAB/simulink simulation is carried out for uav longitudinal motion, realize the rapid and stable, the control command and response to complete the nonlinear fault-tolerant control of flying wing uavs.

전익기형 무인기의 비행 안정성 향상을 위한 형상 최적화 연구

전익기형 무인기의 비행 안정성 향상을 위한 형상 최적화 연구

Investigation and Improvement of Pusher-Propeller Installation Effect for Flying Wing UAV

This paper presents a collaborative experimental and numerical research to investigate the influences of the installation of propeller on aircraft aerodynamic performance, and two approaches to mitigate the propeller installation effect are also analyzed. The configuration under investigation is based on a real flying wing unmanned aerial vehicle with a two-bladed pusher-propeller mounted on the aft part of its associated airframe. Reynolds-averaged Navier Stokes simulations coupled with a structured finite-volume cell-vertex based solver are applied to propeller uninstalled and installed configurations, and the numerical results are shown to be in good agreement with the experimental data. The comparative results obtained through prop-uninstalled and prop-installed cases show that the installation of the pusher-propeller leads to the decline of the maximum lift-drag ratio due to aerodynamic interactions between the propeller and airframe. Detailed analysis of the numerical results highlighted that the low pressure areas on the aft part of the airframe generated by the installation of the pusher-propeller will result in an increase of the base drag and hence a decline of the maximum lift-drag ratio. The feasibilities of center shaft extension and the modification of airframe geometry to mitigate the propeller installation effect are verified by numerical simulations, and results showed that proper propeller–airframe spacing and good designs of airframe geometry can effectively improve the flow characteristic and reduce the drag induced by propeller installation effect.

Gust alleviation control for flying-wing UAV by control surface based on limited parameter variation rate

Gust alleviation control for flying-wing UAV by control surface based on limited parameter variation rate

Attitude Control of Flying Wing UAV Based On Advanced ADRC

The main points of the problem are that the flying wing UAV has strong coupling, obvious nonlinear characteristics and significant changes in steering efficiency during the flight, a combination of ADRC (active disturbance rejection control) and PID (proportional-integral-derivative) was proposed. The adaptive control method is designed with the angle and angular velocity as the controlled variables. The UAV flight attitude control system was designed, where the ADRC control law is used as the outer ring to control the angle, and the inner ring uses the PID to control the angular velocity. In order to solve the problem of high frequency jitter in traditional ADRC, a new nonlinear function is used to optimize the design. Simulink simulation shows that the attitude control overshoot is less than 1% and the adjustment time is less than 2s, which satisfies the control of the flying wing UAV. The flight test was finally carried out to verify the feasibility of the algorithm at the request.

Computational Simulation of Flight Behavior of Flying Wing UAV

Abstract — The aim of this work is to model and simulate a Flying Wing Aircraft and compare acquired flight data from two different simulations. The framework proposed constitutes of the development of a graphical model as well as the use of a mathematical model of the aircraft. In order to implement the Software-in-the-Loop (SIL), we used a commercial flight simulation environment, X-plane 10, which simulates the dynamics of the aircraft through the Blade Element Theory, and the software MATLAB/Simulink, which simulates and implement the control laws. We collected the flight data from the simulator, as well as the responses of the mathematical model based on equations of motion and the aerodynamic derivatives, which were computed via the Digital DATCOM software. Both analytical simulation and the SIL simulation were performed with the same flight conditions, validating the longitudinal dynamic of the computational model and allowing further studies of this type of aircraft in future projects.

Nonlinear robust control method for maneuver flight of flying wing UAV

Nonlinear robust control method for maneuver flight of flying wing UAV

Control system design of flying-wing UAV based on nonlinear methodology

Abstract In this paper, A fluid vector rudder flying-wing UAV is employed as the design object, so as to study the nonlinear design method and flight validation. For the maneuvering flight control, this paper presents a control structure. This control structure included the inner loop linearization decoupling methods to eliminate the known negative coupling and the outer loop backstepping methods for trajectory tracking control. The stability of the control structure has been proved in this paper. Compared with the traditional backstepping control method, this controller increases the inner loop decoupling structure and retains the aerodynamic damping term which makes the linearized system a weak nonlinear system. This structure can not only reduce the conservatism of the outer loop controller design, but also is convenient for engineering implementation. Simulation and flight validation results show that the proposed control scheme is effective.

AN H∞ DESIGN METHOD FOR THE PITCH ATTITUDE HOLD AUTOPILOT OF A FLYING UAV

The paper presents an H∞ loop shaping method for the design of the automatic flight control system of an unmanned air vehicle's (UAV's) longitudinal dynamics. The design objectives include robust stabilization with respect to modeling uncertainties, pitch angle tracking of an ideal model and reduced sensitivity with respect to low frequency measurement errors. The design is illustrate by a case study for a flying wing UAV. The applications based on Unmanned Air Vehicle's (UAV's) received much attention over the last decades. Among them, the flying wing configuration of the UAVs was intensively analyzed due to their advantages concerning the reduced drag and the fuel consumption. A disadvantage of this configuration with respect to the classical one is its instability which fact requires an automatic control system able to stabilize the aircraft and to provide acceptable maneuverability qualities. The design problem of such a system has been previously considered in the control literature (see e.g. (1), (2), (3)). The proposed design methods include both classical techniques based for instance on the well- known PID (Proportional Integral Derivative) control laws, optimal approaches based on the linear quadratic problem, the H∞ norm minimization and nonlinear methods including nonlinear inversion, etc ((4), (5), (6), (7), (8)). The methods mentioned above have advantages and drawbacks; the main difficulty in choosing the control design method is to ensure a trade-off between the complexity of the automatic flight control system and its performances. The actual applications require a wide spectrum of performances including not only the stabilization of the aircraft but also robustness with respect to modeling uncertainties and flying conditions changes, time response performances and reduced sensitivity to disturbances and measurement errors. The aim of this paper is to present a design methodology for the Pitch Attitude Hold (PAH) autopilot of a flying wing configuration. The design procedure is illustrated and investigated for the Hirrus UAV designed and manufactured by a private Romanian company in collaboration with academics from Faculty of Aerospace Engineering of University Politehnica of Bucharest. The design approach is based on a modified version of the so-called two degrees of freedom (2 DOF) H∞ loop-shaping able to accomplish simultaneously several objectives as robustness stability, model tracking and disturbances attenuation requirements. The original 2 DOF H∞ method presented in (9) have been previously used for the autopilot design of the aircraft longitudinal and lateral SCIENTIFIC RESEARCH AND EDUCATION IN THE AIR FORCE-AFASES 2016