Flight Control Law Design Research Articles

Commonalities between robust hybrid incremental nonlinear dynamic inversion and proportional-integral-derivative flight control law design

Incremental Nonlinear Dynamic Inversion (INDI) has received substantial interest in the recent years as a nonlinear flight control law design methodology that features inherent robustness against bare airframe aerodynamic variations. However, systematic studies into the robust design benefits of INDI-based control over the classical divide-and-conquer philosophy have been scarce. To bridge this gap, this paper compares the setup of hybrid INDI with a standard industry benchmark that is based on two-degree-of-freedom gain-scheduled proportional-integral-derivative control. This is done on an architectural basis and in terms of achievable robust stability and performance levels with respect to a common set of design requirements. To this end, a non-smooth, multi-objective H∞-synthesis algorithm is used that incorporates mixed parametric and dynamic uncertainties in the design objective and constraints. It is shown that close similarities exist between hybrid INDI design and gain-scheduled PID control, which leads to virtually equivalent robustness and performance outcomes in both linear time-invariant and linear time-varying contexts. It is therefore concluded that the main benefit of the hybrid INDI does not lie in improved robustness properties per se, but in the opportunity to perform modular robust design in an implicit model-following context. Specifically, this implies that the areas of flying qualities, robustness, and nonlinear implementation are directly visible and accessible in the control law structure.

Sliding Mode Flight Control Law Design Requirements for Oblique Wing Aircraft Based on Perturbation Theory

Sliding Mode Flight Control Law Design Requirements for Oblique Wing Aircraft Based on Perturbation Theory

Piloted Simulation-Based Assessment of Simplified Vehicle Operations for Urban Air Mobility

This work investigates the simplified vehicle operations paradigm, which seeks significant reductions in pilot workload and training requirements through the holistic design of flight control laws, control inceptors, and cockpit displays in the context of vertical takeoff and landing urban air mobility aircraft using piloted flight simulations. Two inceptor layouts, one similar to conventional fixed-wing aircraft and the other similar to conventional rotorcraft, differing in inceptor-to-command mappings, were incorporated into two research flight simulators. A lift-plus-cruise aircraft simulation model, a total energy control system–based fly-by-wire flight control architecture, and identical cockpit displays were used for both setups. Twenty-one participants, divided into three groups based on their prior piloting experience, were used to study pilot performance using representative handling qualities task elements. The piloted simulation data revealed that participants in each group were able to successfully perform the simulation tasks and allowed handling quality ratings to be calculated. They also highlighted instances of control coupling between inceptor axes in each setup for different tasks and the relative ease or difficulty of performing each task with the two inceptor designs.

Suggestions for Criteria to Evaluate Lateral-Directional Nonlinear Pilot-Induced Oscillations Due to Fly-by-Wire Civil Aircraft Landing Configuration Switch

Using a nonlinear pilot-induced oscillation prediction method based on digital virtual flight simulation calculations, digital experiments on predicting lateral-directional nonlinear pilot-induced oscillations due to landing configuration switching of fly-by-wire civil aircraft with different closed-loop dynamic characteristics are carried out. It is proposed that the lateral-directional pilot-induced oscillations due to the landing configuration switch can be evaluated using the changes in dynamic characteristic parameters before and after the configuration switch. The quantitative boundaries of the dynamic characteristic parameters of an example aircraft are determined, and a criterion suggestion is formed to predict the lateral-directional nonlinear pilot-induced oscillations due to landing configuration switching. This study provides a reference for the optimal design of the lateral-directional flight control law of fly-by-wire aircraft during the approach stage and provides suggestions for the formulation of evaluation criteria for other nonlinear pilot-induced oscillation phenomena.

A multi-model and multi-objective approach to the design of helicopter flight control laws

A multi-model and multi-objective approach to the design of helicopter flight control laws

Improved Adaptive NDI Flight Control Law Design Based on Real-Time Aerodynamic Identification in Frequency Domain

The traditional aircraft controller design is usually based on the off-line aerodynamic model. Due to the deviation of the off-line aerodynamic model, the flight quality is difficult to meet the requirements when the aircraft is flying in the real atmosphere. To solve this problem, this paper proposes a frequency domain identification-based improved adaptive nonlinear dynamic inversion (NDI) control method (FDI-ANDI). In this paper, an online recursive aerodynamic parameter identification method in the frequency domain is first designed, and then an adaptive dynamic inversion control method based on the online aerodynamic parameter identification results is established. Finally, aiming at the problem of the slow response speed of the NDI controller, an improved adaptive dynamic inversion control law is designed by using the method of series lead correction. Compared with the traditional control method, the adaptive dynamic inversion method based on online aerodynamic identification has stronger robustness and a faster response speed in the face of model uncertainty. The final simulation analysis shows that the method has a better control effect than the traditional control method.

Suggestions for flying qualities requirements of autonomous control unmanned combat aerial vehicles

There is a lack of quantitative flying qualities assessment requirements for unmanned combat aerial vehicles. The mission-oriented flying qualities evaluation approach can make up for the deficiencies of existing flying qualities specifications. Considering the control characteristics and mission requirements of autonomous control unmanned combat aerial vehicles, flying qualities assessment tasks are designed and performance standards are developed on the basis of manned aircraft flying qualities assessment tasks. Multiple sets of mathematical simulations are performed by varying the control law parameters to investigate the relationship between the control law parameter values, closed-loop aircraft system characteristics and flying qualities levels. The simulation results and closed-loop frequency domain analysis show that the existing flying qualities guidelines for manned aircraft are not fully applicable to the flying qualities assessment of autonomous control unmanned combat aerial vehicles. It is found that the combination of the bandwidth and the phase at the bandwidth frequency can define the flying qualities requirements of autonomous control unmanned aerial vehicles. The criterion boundaries of different levels are given, and the physical reasons for the formation of boundary are analysed. Our results can be applied to flying qualities assessment and design of flight control laws for autonomous control unmanned combat aerial vehicles.

Research on Dynamic Characteristics Analysis and Control Law Design Method of Model Aircraft in Wind Tunnel–Based Virtual Flight Testing

In the three-degree-of-freedom virtual flight test, the scaled model has no translational motion, which will result in the model’s dynamics being different from that of free flight, and this difference will directly affect its effectiveness in simulating free flight and the design of the flight control law. In this paper, the longitudinal and lateral-directional dynamic characteristics of virtual flight are studied from the three aspects of modal characteristics, disturbance response characteristics, and control response characteristics. The dynamic characteristic difference between the virtual flight test and free flight is analyzed. In addition, the scaling ratio of the model will also lead to differences in the flight control law design methods of virtual flight and full-size aircraft. The control law design method of the virtual flight test is established by studying the similarity relation of control law parameters between full-size aircraft and the scaled model, and the correctness of similar design of closed-loop control law parameters of the scaled model is verified by a virtual flight test. Finally, the attitude control law parameters obtained from the virtual flight test are applied to the six-degree-of-freedom flight simulation of a full-size aircraft, and the simulation results show that the three-axis attitude control law designed based on the virtual flight test also has good performance in the six-degree-of-freedom flight of the full-size aircraft, and the response of attitude angle and angular rate are approximately the same for both.

Aerodynamic Identification and Control Law Design of a Missile Using Machine Learning

The new generation of air vehicle is confronted with a more intricate environment and challenging missions, which puts forward higher requirements for the flight control system. In this study, the aerodynamic identification and control law design based on machine learning for a missile configuration is investigated through numerical simulations. The missile pitch and elevator deflection are realized via the combination of a rigid dynamic grid method and an overlapping grid technology, while the computational fluid dynamics/rigid body dynamics (CFD/RBD) strong coupling method is implemented to simulate the unsteady flows associated with the motion of the missile. Firstly, the aerodynamic data of the missile are gathered through forced pitching motion involving elevator deflection, and an aerodynamic model is constructed using a deep neural network to identify the aerodynamic moment with only a small number of unsteady aerodynamic data. Then, the accuracy and fidelity of the model are checked with the open-loop control law. Afterward, a missile pitch control law is generated through deep reinforcement learning based on the aerodynamic model, which enables the realization of a robust and exact angle-of-attack control process. Finally, the control law is transferred to a numerical environment and numerical virtual flight based on CFD is conducted, which demonstrates that stable control can be maintained even in continuous maneuvering. This study verifies the possibility of applying a deep neural network to air vehicle aerodynamic identification and deep reinforcement learning to a complicated flight control law design with excellent generalization ability. Machine learning is expected to play a significant role in the design and research for the novel generation of air vehicle.

Performance enhancement of a bioinspired micro air vehicle by integrating a smart composite in its morphing wing

The purpose of this paper is to show the advantages of using a smart composite in a micro air vehicle (MAV) equipped with morphing wing technology. A Macro Fiber Composite (MFC) actuator is attached to the wing’s bottom surface to modify the wing camber during the mission. This material allows the MAV to be optimized according to each flight, thus making it more versatile and attractive to the market. The elongation of the lower surface when a positive voltage is applied to the actuator is translated to an increment in camber, which results in an increment in the maximum lift coefficient, thus enabling the vehicle to fly slower to adapt to any payload. Besides, a reduction in camber results in an increase in aerodynamic efficiency, which improves range and endurance. Several tests of the MAV at prototype level have been carried out at INTA, so as to demonstrate the feasibility of implementing MFC actuators to control and manoeuvre these vehicles. The use of this material in aerospace industry opens up various fields of research in aerospace engineering, such as new features in flight mechanics and aerodynamic performance and new strategies in the design of flight stability and control laws.

Trajectory Optimization of Multi-UAVs for Marine Target Tracking during Approaching Stage

This paper focuses on a coordinated tracking planning method of multiple unmanned aerial vehicles (UAVs), which was deployed in the best positions to better fulfill the marine target localization tasks when approaching the target. The optimal planning of multi-UAVs was implemented using an online centralized nonlinear model predictive control (NMPC) based on the target state’s uncertainty criteria. The penalty function is used to solve UAV platform dynamic performance in the model predictive control method to consider the more realistic situation. The coordinated planning problems of multi-UAVs are numerically simulated and compared with the Lyapunov vector field guidance (LVFG) method under classical mission scenarios. Simulation results demonstrate that the algorithm can maintain the optimal observation configuration of multi-UAVs to improve the marine target positioning accuracy, verifying the feasibility and superiority of this method. Furthermore, the simulation results can provide a useful reference for the flight control law design of multi-UAVs with optimal observation configuration.

Design of adaptive flight control law for center of gravity and mass variation and HIL real-time simulation verification

Considering the influence of the center of gravity (CG) and mass variations on flight dynamics, a nonlinear L1 adaptive control approach is presented in this paper and applied to design a flight controller to achieve the aim of improving the robustness of the flight system against sudden changes in the CG. Based on L1 adaptive control, this approach introduces a feedback linearization strategy to eliminate the known nonlinear dynamics of the aircraft and realize the rapid decoupling of the aircraft states. The nonlinear L1 adaptive control also addresses the problem of insufficient robustness in conventional L1 adaptive control due to overcoming its own known nonlinearities, and then improves the robustness to external disturbances. Additionally, the modified piecewise constant is employed to design the adaptive law to improve the estimation accuracy. Finally, the control performance and robustness of the designed flight control law are verified in the hardware in the loop (HIL) flight control semi-physical experimental platform. Experimental results indicate that the nonlinear L1 adaptive flight control law has strong robustness, which can effectively overcome the disturbance of the CG and mass sudden variations. Moreover, the designed controller takes into account the transient performance while ensuring the steady-state performance of the aircraft with CG change.

Automatic weighting filter tuning for robust flight control law design

This article presents a practical framework for the co-design of weighting filter parameters and controller gains in the context of multi-objective/multi-model tuning of flight control laws. A non-smooth optimization-based procedure is proposed aiming to reduce the complexity of the control design process, evolving from a rather classical manual filter selection all the way to a fully automatic one. A normal acceleration Control Augmentation System (CAS), designed by making use of non-smooth optimization software, for the linearized dynamics of an agile autonomous aerospace vehicle is used as an illustrative example. Several performance indexes are compared in the context of the developed framework as well as against more classical tuning methods such as pole placement, demonstrating that the proposed approach may lead to improved results and potentially benefit the flight control engineer via an easier handling of conflicting design requirements, model uncertainty and application-tailored control structures.

Aircraft Configuration Parameter Boundaries Based on Closed-Loop Flying Qualities Requirements

For aircraft employing the fly-by-wire technique, the closed-loop dynamic characteristics are determined by both the configuration design and the flight control system. As the capacity of the control system has certain limitations, the configuration parameters are also constrained by the requirements of the closed-loop flying qualities. This paper presents an aircraft configuration parameter boundaries determination method based on closed-loop flying qualities requirements independent of the actual flight control law design, mainly aiming at the parameters that affect the stability and control characteristics. First, a nonlinear dynamic inversion-based flight control law is adopted to decouple the control law gains from the configuration parameters and to study the relationship between the configuration parameters and closed-loop flying qualities. Second, a flying qualities evaluation scheme is established by selecting the most severe flight conditions and the evaluation criteria that are most sensitive to changes in the parameters. Finally, the parameter boundaries according to the requirements of Level 1 flying qualities are determined by searching for the critical values that lead to degradation of the flying qualities. The proposed method is verified by an application example of the design ranges of a sample aircraft’s wing position, horizontal tail area, center of gravity, vertical tail area and vertical tail position.

Stability characteristics and airworthiness requirements of blended wing body aircraft with podded engines

Blended-Wing-Body (BWB) aircraft have a relatively short fuselage and no horizontal tail, and they usually adopt podded engines and a V tail instead of a vertical tail. Generally, BWB aircraft have decreased longitudinal and directional static stability and damping. In this paper, the three-axis static and dynamic stability characteristics of an example BWB aircraft with podded engines are studied. According to the differences in flight characteristics of BWB aircraft and conventional aircraft, the different airworthiness requirements for BWB aircraft are analyzed: first, based on current airworthiness regulations and transport aircraft flying quality specification, the relaxation requirement of longitudinal static stability for BWB aircraft is studied; second, the influences of podded engines on longitudinal trim, attitude and trajectory responses and maximum directional control power requirement of BWB aircraft are analyzed; third, the changes in the proportional relationships between the takeoff characteristic speeds of the example BWB aircraft are analyzed. In view of these variations in directional control power requirement and relationships between takeoff characteristic speeds, new recommendations for airworthiness evaluation are proposed for BWB aircraft. The conclusions of this paper are helpful references for configuration design, flight control law design and airworthiness evaluation of BWB aircraft.

Design of sliding mode flight control system for a flexible aircraft

The evolution of large transport aircraft is characterized by longer fuselages and larger wingspans, while efforts to decrease the structural weight reduce the structural stiffness. Both effects lead to more flexible aircraft structures with significant aeroelastic coupling between flight mechanics and structural dynamics, especially at high speed, high altitude cruise. The lesser frequency separation between rigid body and flexible modes of flexible aircraft results in a stronger interaction between the flight control system and its structural modes, with higher flexibility effects on aircraft dynamics. Therefore, the design of a flight control law based on the assumption that the aircraft dynamics are rigid is no longer valid for the flexible aircraft. This paper focuses on the design of a flight control system for flexible aircraft described in terms of a rigid body mode and four flexible body modes and whose parameters are assumed to be varying. In this paper, a conditional integral based sliding mode control (SMC) is used for robust tracking control of the pitch angle of the flexible aircraft. The performance of the proposed nonlinear flight control system has been shown through the numerical simulations of the flexible aircraft. Good transient and steady-state performance of a control system are also ensured without suffering from the drawback of control chattering in SMC.

Design of Stability Augmentation Systems for Flexible Aircraft Using Projective Control

There are many challenges related to the design and operation of flexible aircraft. Flight control law design for improving handling qualities is one of them because the major problem in the design of controllers then concerns aeroservoelastic stability. To deal with this difficulty, methodologies for flight control law design considering the aeroelastic dynamics of the aircraft are being pursued. In this paper, a static output-feedback-based stability augmentation system is proposed and designed to improve the handling qualities and the structural dynamics decoupling of a flexible aircraft. The design is based on the projective control technique, which allows preserving in the closed-loop system the eigenstructure of certain modes of interest whose dynamic characteristics stem from an optimal state feedback solution. An experimental prototype of a flexible aircraft that mimics a high-altitude long-endurance airplane, called X-HALE, is considered in the case studies. Robustness analysis based on classical and disk-based stability margins, and nonlinear simulations of gusts and turbulent flight conditions evaluated and confirmed the effectiveness of the proposed controller.

Flying qualities evaluation based nonlinear flight control law design method for aircraft

This paper is focused on the parameter tuning problem for the design of nonlinear flight controllers. Taking into account the preferences and requirements of the pilot, a novel flying qualities evaluation framework and a tuning method are proposed for the design of nonlinear flight controllers. Considering that the existing flying qualities criteria are not applicable to evaluating the closed-loop aircraft system, a novel evaluation framework consisting of three modules namely ideal reference model, controller and aircraft model, is proposed. Based on this evaluation framework, an overall evaluation model is established, the flying qualities based rating scores for this model will reflect the level of the controller performance. Subsequently, a schematic evaluation method, consisting of three sub-procedures namely linearization, low-order equivalent fitting and flying qualities criteria based rating, is proposed for the controller design. To validate the evaluation framework and the method, a nonlinear F/A-18 aircraft model and the incremental backstepping control law are chosen to build a demonstration example. Theoretical analysis and the simulation results reveal that the proposed evaluation framework and schematic method are effective for the design of nonlinear flight control laws.

UAV's Precise Trajectory Tracking Control Based on Nonlinear Dynamic Inverse and Quaternions

The precise tracking 3D trajectory that changes dramatically within the boundary of flight performance is the essential capability of the UAV applied to air combat, which requires that flight control law design be adapted to nonlinear characteristics of flight dynamics and solve the singularities problem during the flight. Continuous trajectory control on the basis of nonlinear dynamic inversion (NDI) is studied. The singularities problem is caused by continuous trajectory expressed by flight velocity and flight-path angles. A sudden change of flight-path axis system happens when aircraft flies along a continuous trajectory which results in singularities. A method for amending flight-path axis system is presented to solve the sudden change problem of flight-path axis system. Finally, the combination of the maneuver generator based on quaternions and amending flight-path axis system realize the precise trajectory tracking control without singularities and obtain excellent tracking characteristics.

Post-stall flight dynamics of commercial transport aircraft configuration: A nonlinear bifurcation analysis and validation

Loss-of-control has become the largest fatal accident category for worldwide commercial jet accidents, and any initiative aimed at preventing such events requires an understanding of the fundamental aircraft behavior, especially the flight dynamics at post-stall region at which loss-of-control usually occurred. A series of low-speed static and dynamic wind tunnel tests of the Common Research Model over a large angle of attack/sideslip envelope was conducted and a non-linear aerodynamic model was developed. The bifurcation analysis, complemented by time-history simulation was used to understand the post-stall flight dynamics and the numerical analysis results were preliminary validated by wind tunnel virtual flight test. Several representative post-stall behaviors for the transport aircraft have been identified, including departure, periodic oscillation, post-stall gyration and steep spiral, etc. Furthermore, the predicted periodic oscillation in pitch motion has been perfectly duplicated in wind tunnel virtual flight test. The approach used in this work shows a promising way to uncover the flight dynamics of transport aircraft at extreme and loss-of-control flight conditions, as well as to apply to nonlinear unsteady aerodynamics modeling and validation, flight accident investigation, advanced flight control law design or studying initiative for loss-of-control prevention or mitigation.