Flight Envelope Research Articles

Agile maneuvering with a small fixed-wing unmanned aerial vehicle

This paper presents a general and systematic approach to automating a variety of agile maneuvers with a small fixed-wing unmanned aerial vehicle. The methodology begins by numerically solving optimal control problems off-line to generate a small set of reference trajectories and feedforward control inputs for maneuvers. A dynamic time warping-based interpolation process parametrizes these solutions, adding robustness to the maneuver, whilst allowing the on-board library of state and control time histories to remain compact. To handle errors, inaccuracies, noise, and disturbances that are not accounted for by feedforward control, feedback control laws stabilize about the reference trajectories. The work focuses mainly on one agile maneuver: an aggressive turn-around, in which the aircraft undergoes a rapid 180 degree heading reversal, beginning and ending in straight and level flight. To establish the generality of the methodology, it is also applied to transition maneuvers between straight and level flight, and a nose-up hover. The proposed automation scheme is computationally light during flight, consisting of a simple feedforward/feedback controller coupled to a compact library of maneuvers that are optimized over the full flight envelope. The methodology is validated in simulations and flight tests. The automation scheme is implemented successfully on a small fixed-wing unmanned aerial vehicle.

Wind flow investigation over an aircraft carrier deck by PIV

An aircraft carrier is characterized by a superstructure where different systems are installed. This configuration is similar to a set of bluff bodies with sharp edges where massive flow separation appears in the wake. Depending on the wind-over deck (WOD) conditions, the flow field changes dramatically. Hence, aircraft performances are affected due to the adverse aerodynamic effects generated. Flight envelopes are needed to determine aircraft operating range. In situ measurements are costly and time-consuming. Numerical simulations are providing some solutions to this problem. However, they have to be validated by matching simulation results to those from experimental measurements. In the present paper, the study of the flow configuration at different wind angles has been conducted. Particle image velocimetry technique was used to obtain flow information due to the accurate results provided. Non-dimensional velocity and turbulence maps are presented in order to provide flow field characterization.

CFD-Based Aeroelastic Sensitivity Study of a Low-Speed Flutter Demonstrator

The goal of developing aircraft that are greener, safer and cheaper can only be maintained through significant innovations in aircraft design. An integrated multidisciplinary design approach can lead to an increase in the performance of future derivative aircraft. Advanced aerodynamics and structural design technologies can be achieved by both passive and active suppression of aeroelastic instabilities. To demonstrate the potential of this approach, the EU-funded project Flutter Free Flight Envelope Expansion for Economical Performance Improvement is developing an unmanned aerial vehicle with a high-aspect-ratio-wing and clearly defined flutter characteristics. The aircraft is used as an experimental test platform. The scope of this work is the investigation of the aeroelastic behaviour of the aircraft and the determination of its flutter limits. The modeling of unsteady aerodynamics is performed by means of the small disturbance CFD approach that provides higher fidelity compared to conventional linear-potential-theory-based methods. The CFD-based and the linear-potential-theory-based results are compared and discussed. Furthermore, the sensitivity of the flutter behaviour to the geometric level of detail of the CFD model is evaluated.

Adaptive Backstepping Control for Longitudinal Dynamics of Hypersonic Vehicle Subject to Actuator Saturation and Disturbance

In this paper, an adaptive backstepping control strategy is presented to solve the longitudinal control problem for a hypersonic vehicle (HSV) subject to actuator saturation and disturbances. Small perturbation linearization transforms the dynamics to a seconded-order system at each trimming point, with total disturbance including unmodeled dynamics, parametric uncertainties, and external disturbances. The disturbance can be estimated and compensated for by an extended state observer (ESO), and thus the system is decoupled. To deal with the actuator saturation and wide flight envelope, an adaptive backstepping control strategy is designed. A rigorous proof of finite-time convergence is provided applying Lyapunov method. The effectiveness of the proposed control scheme is verified in simulations.

Determination of the flight dynamic envelope via stable manifold

The flight envelope plays an important role in flight safety. The concept of posing the flight envelope as a region of attraction is explored further, and it is investigated whether the stable manifold for the region of attraction computation is an efficient method for determining envelope. The stable manifold describes the flight dynamic envelope of an aircraft in an explicit representation, which means that the computation needs to be done only on the envelope, not the entire state space. In this paper, the stable manifold is computed by using a fast method which reduces the computation to solving a system of partial differential equation. Then, the stable manifold grows in the way of advancing front mesh generation framework. The stable manifold is then applied to the envelope determination of a nonlinear F-16 model. The result is compared to the results obtained with the level set method, demonstrating that the stable manifold provides a feasible and accurate result to the dynamic envelope. The proposed method is then used to investigate the effect of actuator failure on the flight safety. The proposed method can also be used as a safety assessing tool during the design phase of an aircraft.

Thrust control strategy based on the minimum combustor inlet Mach number to enhance the overall performance of a scramjet engine

A lower combustor inlet Mach number is desirable in order to design a compact, lightweight combustor and boost the overall performance of the scramjet engine. In this study, a thrust control strategy is proposed for a hydrogen-fueled scramjet taking into account the operating limitations, which is called the minimum combustor inlet Mach number rule since the combustor inlet Mach number is used as the control variable. By scheduling the fuel supply and modifying the intake geometry, the combustor inlet Mach number can be minimized while ensuring a certain thrust output within the operation constraints. In this manner, the scramjet engine can be operated with high specific thrust and low fuel consumption throughout the flight envelope. The thrust control strategy is further applied to a hydrogen-fueled scramjet in the hypersonic flight regime. Because the combustor inlet Mach number varies with flight conditions, the thrust strategy can be applied in practice by monitoring the following aerothermodynamic parameters in different flight regimes instead: (1) combustor outlet Mach number, (2) combustor inlet static temperature, and (3) combustor outlet static temperature. Furthermore, the effects of the thrust output on the division of flight regime are investigated, and the overall performance of the hydrogen-fueled scramjet engine obtained from applying the thrust control strategy is discussed in detail.

Two freedom linear parameter varying μ synthesis control for flight environment testbed

To solve the problem of robust servo performance of Flight Environment Testbed (FET) of Altitude Ground Test Facilities (AGTF) over the whole operational envelope, a two-degree-of-freedom μ synthesis method based on Linear Parameter Varying (LPV) schematic is proposed, and meanwhile a new structure frame of μ synthesis control on two degrees of freedom with double integral and weighting functions is presented, which constitutes a core support part of the paper. Aimed at the problem of reference command’s rapid change, one freedom feed forward is adopted, while another freedom output feedback is used to meet good servo tracking as well as disturbance and noise rejection; furthermore, to overcome the overshoot problem and acquire dynamic tuning, the integral is introduced in inner loop, and another integral controller is used in outer loop in order to guarantee steady errors; additionally, two performance weighting functions are designed to achieve robust specialty and control energy limit considering the uncertainties in system. As the schedule parameters change over large flight envelope, the stability of closed-loop LPV system is proved using Lyapunov inequalities. The simulation results show that the relative tracking errors of temperature and pressure are less than 0.5% with LPV μ synthesis controller. Meanwhile, compared with non-LPV μ synthesis controller in large uncertainty range, the proposed approach in this research can ensure robust servo performance of FET over the whole operational envelope.

Motion Equations and Attitude Control in the Vertical Flight of a VTOL Bi-Rotor UAV

This paper gathers the design and implementation of the control system that allows an unmanned Flying-wing to perform a Vertical Take-Off and Landing (VTOL) maneuver using two tilting rotors (Bi-Rotor). Unmanned Aerial Vehicles (UAVs) operating in this configuration are also categorized as Hybrid UAVs due to their ability of having a dual flight envelope: hovering like a multi-rotor and cruising like a traditional fixed-wing, providing the opportunity of facing complex missions in which these two different dynamics are required. This work exhibits the Bi-Rotor nonlinear dynamics, the attitude tracking controller design and also, the results obtained through Hardware-In-the-Loop (HIL) simulation and experimental studies that ensure the controller’s efficiency in hovering operation.

Eye tracking as a debriefing tool in upset prevention and recovery training (UPRT) for general aviation pilots

Upset prevention and recovery training (UPRT) is intended to improve the ability of pilots to recognize and avoid situations that can lead to airplane upsets and to improve their ability to recover control of an airplane that has exceeded the normal flight envelope. To this end, a set of different training contents – from theoretical knowledge of aerodynamics and human factors to practice-based flight training – is necessary. In order to support the debriefing with an objective feedback, and because visual scanning is a core competence, two studies on subjective evaluation of aviation pilots – one conducted in a flight simulator and the other one in-flight – focussed on the practical application of eye tracking as a debriefing tool in UPRT. From a practitioner’s perspective, eye tracking appeared to be a useful method in terms of visualising instrument scanning techniques, supporting the instructor with objective debriefing material and fostering self-awareness in human processes. The discussion recommends adjusted UPRT instructor training and further improvements to eye tracking hardware and software.Practitioner Summary: The article focuses on pilot evaluations of eye tracking as a debriefing tool in UPRT and the identification of critical elements in its use. Eye tracking is a promising debriefing tool for UPRT. The discussion points to desirable improvements of eye tracking hardware and software as well as adjustments to instructor training that are pertinent.

Nonlinear Control of Turbofan Engines: An Active Set-Based Method for Performance Optimization

The inherent nonlinear nature of engine dynamics necessitates advanced design techniques for transient control. Conventional design methodologies are either not ready to apply to large flight envelope control or failing to provide protection over a variety of physical limits. This paper proposes an active set-based method for performance optimization over large envelope while providing limit protection over all sorts of constraints. Detailed design procedures are provided, and extensive numerical investigations are presented with both robustness and implantation issues discussed. Comparisons with both conventional schedule-based control and an advanced nonlinear generalized minimum variance-based (NGMV) control are conducted to illustrate the effectiveness of the proposed method.

Control of Transonic Buffet by Shock Control Bumps on Wing-Body Configuration

The aerodynamic shock buffet phenomenon limits the flight envelope of typical transonic aircraft at high Mach number or incidence. For future aircraft design, its correct prediction by numerical methods and control strategies are of great interest. In the present paper, simplified industrial applicable design guidelines for shock control bumps are summarized and assessed based on numerical (unsteady) Reynolds-averaged Navier–Stokes simulations of a wing-body configuration. The design guidelines cover both design objectives, wave drag reduction at dash flight, and buffet delay/alleviation. Three different shock control bumps have been designed: a two-dimensional bump for wave drag reduction at dash flight, a two-dimensional bump for delay of buffet onset, and a three-dimensional bump array with the intention of combining both design objectives. The two-dimensional bumps mark the maximum potential for each design objective with a drag reduction of 9 drag counts at dash flight and a delay of buffet onset by 4 lift counts. Even though the three-dimensional bump array combines both design objectives, it does not perform as well as expected due to massive crest-flow separation and a (so far) minor role of the vortical wake of three-dimensional bumps in delaying buffet onset.

Flight dynamics modeling of elastic aircraft using signal decomposition methods

To improve the precision and accuracy of the flight dynamic models for elastic aircraft, this paper provides a novel method that extracts observable flight modes from flight test data and uses them in the identification process. For this purpose, a gray-box time-domain method is employed with the nonlinear ARX structure and the Levenberg–Marquardt parameter estimation technique. In the proposed method, the components of the flight parameters are extracted by two signal decomposition techniques, namely the singular spectrum analysis and empirical mode decomposition. These components are inputted into the identification process. Flight test data of the active aeroelastic wing is examined by the proposed method. The results indicate that both the singular spectrum analysis-based and empirical mode decomposition-based identification processes have desired performances in dealing with nontrained flight conditions. Thus, these methods may be utilized as an interpolation method to estimate the aircraft flight dynamics within the flight envelope. However, the empirical mode decomposition outperforms the singular mode decomposition because it is more significant in decomposing flight parameters based on the frequency content. Hence, the empirical mode decomposition may be a better tool to be employed for the aeroelastic aircraft system identification.

Integration of Phase Plane Flight Envelope Protections in Cascaded Incremental Flight Control

Abstract With advances in incremental non-linear flight control, new challenges and possibilities for incorporating efficient flight envelope protections arise. Investigation of a phase plane based command and virtual control limiting pitch angle protection in the reference model of a cascaded incremental backstepping controller is performed for a commercial aircraft model. This provides an almost model free design of a combined flight control law and envelope protection. Verification through simulation shows that accurate predictable response behaviour is obtained close to the limit, which permits a protection that minimizes the interference into the nominal control.

Aero-Engine On-Board Model Based on Batch Normalize Deep Neural Network

A new on-board turbo-fan engine modeling method based on a batch normalize (BN) mini-batch gradient descent (MGD) deep neural network (NN) is proposed. This new method adopts BN algorithm, which accelerates the network training speed and overcomes the gradient vanish problem. Hence, using the BN algorithm, the neural network adopts the deeper structure, which means the network has a stronger representation capacity. This mini-batch gradient descent (MGD-NN) algorithm that consumes much less time to update the NN parameters is adopted. Therefore, it is more suitable for training big dataset and establishing a high-accuracy engine model in a large flight envelope. Finally, to verify whether the proposed method could be applied to larger flight envelope, the conventional NN also adopts MGD (called MGD-NN). The turbo-fan engine models based on these two modeling methods are both conducted within a sub-sonic cruise envelope. The simulation results show that the proposed modeling method has much higher accuracy than the MGD-NN. Moreover, the proposed method has the characteristics of less data storage, low computation complexity, and good real-time performance, which are the most importance indices for model realize on-board.

Development and Evaluation of a Full Flight Envelope Integrated Flight and Vibration Controller

Development and Evaluation of a Full Flight Envelope Integrated Flight and Vibration Controller

Attitude control of tiltwing aircraft using a wing-fixed coordinate system and incremental nonlinear dynamic inversion

In this paper, we present a novel concept for robustly controlling the attitude of tiltwing aircraft. Our main contribution is the introduction of a wing-fixed coordinate system for angular acceleration control, which forms the basis of a simple and robust attitude controller. Using the wing-fixed coordinate system allows us to describe the actuator effectivity using simple approximations based on the current operating conditions of the aircraft. Coupled with a robust angular rate control concept, which does not rely on an accurate aerodynamic model, we present a controller stabilizing the entire flight envelope of a tiltwing aircraft. The underlying angular acceleration controller uses the concept of Incremental Nonlinear Dynamic Inversion (INDI) to achieve robustness against aerodynamic uncertainties. The resulting controller is evaluated in both simulation studies and flight tests.

Active flutter suppression: non-structured and structured H∞ design

Abstract In this article a comparison of two controllers for the active suppression of flutter of a flexible remotely-piloted aerial vehicle is presented. The H∞ technique is used for both, but the first relies on the standard (non-structured) approach and the second on the structured synthesis. High-fidelity models derived from finite element modeling are used to verify the performance of the obtained controllers. Reduced-order models are used for the control synthesis, which are obtained from applying a combination of balanced and modal reduction techniques on the high-fidelity models. The results show that both controllers are similar in performance and robustness, and both are capable of suppressing flutter and extending the flight envelope well beyond the open-loop flutter speed.

Minimization of Ground Vibration Test Configurations for F‐16 Aircraft by Subtractive Modification

The certification process of external loads designed for aircraft needs to satisfy various criteria where compatibility with existing systems is one of the essential requirements. Flight flutter testing is a critical part of a certification process that requires many preliminary studies. Computational flutter analysis must precede actual flutter test to determine an approximately safe flight envelope to ensure the safety of the personnel and aircraft. To be able to perform flutter analysis of an aircraft, an accurate structural model such as finite element (FE) model is required. An accurate FE model can be obtained from a coarse model using ground vibration test (GVT) which is also the primary test campaign for certification of a new external load, new aircraft design, or modification on existing aircraft. On the other hand, performing GVT for each configuration of an aircraft is both time consuming and costly. It would be more practical to determine the critical configurations for an aircraft using computational tools and perform actual GVT for those configurations. The objective of this study is to simulate GVT characteristics for downloading and fuel configurations of F‐16 aircraft. A novel methodology is proposed where various loading configurations can be simulated by subtractive modification from loaded GVT data so that joint stiffnesses between stores and aircraft need not be identified. The proposed technique decreases the number of necessary physical GVT testing campaigns.

Optimal civil aircraft missions exploiting free routing possibilities

A method is presented for optimizing burnt fuel and flight time of civil aircraft missions, under the scope of a full-scale free route airspace implementation as well as usage of actual environmental data, focusing on ambient temperature and pressure. A computational analysis model, using online available aircraft data is considered, for flight envelopes defined by the user, is used as the base tool. The model is validated over real flight data, before it is used as the main tool for optimization of flight missions. Flight paths are analyzed by considering both their vertical and horizontal aspect. Optimal paths are derived, for each chosen mission, through an optimization process that takes advantage of the flexibilities that become available by implementation of Free Route Airspace. Vertical flight path is shown to play very important role in achieving optimal flights, while horizontal paths also offer optimality possibilities, with a strong dependence on weather conditions.

LPV/PI Control for Nonlinear Aeroengine System Based on Guardian Maps Theory

The requirement of fighter’s maneuverability puts forward higher indexes for the control performance of aeroengine. It is generally recognized that the proportion integration (PI) control is the most commonly used control method of aeroengine. However, due to the high nonlinearity and performance requirements of aeroengine, the selection of PI parameters needs more accurate methods to satisfy higher performance requirements. In this paper, the linear parameter varying (LPV) model is established according to a nonlinear mathematical model of a turbofan engine. Then, the LPV/PI control based on guardian maps theory is proposed to solve the highly nonlinear aeroengine control problem which improves the response characteristics of aeroengine without changing the control method in the range of the scheduling parameters. In the process of design, a set of controller parameters satisfying the performance requirements is automatically generated by a given initial controller parameter, so designing the controllers at multiple equilibrium points is avoided. Finally, the simulations are performed by the integral separation of PI control at different points in the flight envelope for the nonlinear model. The results illustrate that the control method based on guardian maps theory can contribute significantly to solving the nonlinear problems of aeroengine control system.