Full Flight Envelope Research Articles

Active Model-Based Predictive Control and Experimental Investigation on Unmanned Helicopters in Full Flight Envelope

For the control of unmanned helicopters in full flight envelope, an active model based predictive control scheme is developed in this brief. Dynamics in full envelope is modeled, with uncertainties represented by the system model error and process noise. The model error depends on both helicopter dynamics and flight mode, and the process noise is assumed unknown but bounded. Based on the set-membership filter, an active modeling based stationary increment predictive control, based on the estimated model error and its boundary to optimally compensate the model error, as well as the aerodynamics time delay, is proposed. The proposed method has been implemented on the ServoHeli-40 unmanned helicopter platform and experimentally tested; the results have demonstrated its effectiveness.

Research on Integrated Aircraft Modeling for Enhancing Accuracy of Flight Control System Design and Simulation

In order to realize excellent flight control system design and simulation, it is indispensable to establish a more realistic and more integrated aircraft model, so a scheme for an integrated aircraft model which consists of flight dynamics, propulsion system and flight environment is proposed in this research. Three parts of the integrated aircraft model are closely connected and interacted with each other. Furthermore, the aero engine model can reflect the engine thrust changes in the full real flight envelope and gains great parsimoniousness, which is designed based on K-means clustering algorithm and recursive reduced least squared vector regression. Simulation results confirm the effectiveness and preciseness of the integrated aircraft model. It also indicate to the control system designers that flight control system should be integrated with propulsion control, and the robustness and adaptivity of the flight control system should be paid much more attention to, so as to overcome the adverse effects of the flight environment.

A Complete Control System Design for a Tactical Missile Using Model Predictive Control

During the last decade, a significant research effort has been contributed in the area of non-linear missile autopilot design. But in some kinds of missiles, the classical controllers are still effective and very robust provided that the controller can adapt itself with missile varying parameters and this makes the control implementation and algorithm complicated as it changes with flight conditions. In this paper roll and lateral autopilots are designed for short range surface-to-surface aerodynamically controlled missile using model predictive control with fixed algorithm along the whole flight time. From linearization of missile non-linear model, transfer functions are determined and then discretized. The controllers are designed using Model-based predictive control techniques with fixed algorithm and insure that they canachieve full flight envelope control capability. Finally, the designed controllers are conducted into 6DOF simulation (individually and all-together) which is carried out using the Matlab-Simulink software.

Trajectory Planning and Prediction Guidance Based on the Moon-Earth Return Reentry Dynamics

For the lunar return mission, a concern of the entry guidance requirement is the full flight envelope applicability and landing accuracy control. A concise numeric predictor-corrector (NPC) entry guidance (NPCEG) algorithm is developed for this requirement. It plans a real-time trajectory on-line by modulating the linear parameterized bank profile. To meet the path constraint, we propose an integrated guidance strategy which combines NPC method with an analytical constant drag acceleration method. Monte Carlo analysis shows that the algorithm is sufficiently robust to allow precision landing with a delivery error of less than 2.0 km for the entire between 2,500 km and 10,000 km range.

Autonomous Control of Unmanned Combat Air Vehicles: Design of a Multimodal Control and Flight Planning Framework for Agile Maneuvering

There is a growing demand for unmanned air vehicles (UAVs) with combat capabilities in battlefield scenarios [1]. Whether this capability is for evasive maneuvers or for flying attack patterns, unmanned combat air vehicles (UCAVs) are expected to operate in dense and often threatening environments that require aggressive trajectory planning and controls [1]. These trajectories often require the use of maneuvering capability over the full flight envelope of the aircraft. Examples of such trajectories are high-g turns and high angle-of-attack maneuvers. This article presents the development of a multimodal flight control and flight path planning scheme that allows the vehicle to autonomously perform agile maneuvers over its full flight envelope. The key element of this scheme is the maneuver decomposition methodology, which aims to reduce the complexity of the planning and control problems for UCAVs. This article demonstrates how a parameterized family of maneuver modes for a UCAV can be developed systematically and how an arbitrary agile maneuver can be decomposed into simpler segments. In addition, the article presents a multimodal flight control scheme in which each of the maneuver modes is controlled locally by a sliding mode controller. The overall capability of the system is demonstrated in challenging scenarios such as navigation in dense environments and autonomous execution of aerobatics competition sequences.

Experimental method for studying gust effects on micro rotorcraft

Micro rotorcraft have a great potential for both civilian and military applications, but for successful deployment, they must operate robustly in real-world environments including flight in winds. Due to their small size, common wind fields in and around buildings can represent a large percentage of the vehicle’s full flight envelope resulting in highly non-linear motion. An experimental method has been developed to evaluate the capabilities of micro rotorcraft in realistic wind conditions to aid in the design and evaluation of these vehicles. A synthetic wind generation system is used to create repeatable and controllable gust excitations, and a motion capture system is used for precise measurement of vehicle response. Performance metrics associated with the accuracy of position tracking and angular rate excursions are proposed to quantify vehicle capabilities at different excitation levels for design assessment and cross-platform comparison. To demonstrate the experimental method, test results are shown to evaluate the capabilities of a micro coaxial helicopter in realistic winds as a function of gust excitation level and vehicle size. For this example system, light wind gusts degrade hover hold accuracy by a factor of 3 compared to a situation with quiescent winds.

Active Disturbance Rejection Decoupling Control for Aero-Engines

A class of multi-variable decoupling control algorithm for aero-engines is mainly studied In this paper. Firstly, the mechanism of ADRC decoupling control algorithm is elaborated and discussed in detail. Secondly, an ADRC controller for some turbofan engine's nonlinear component-level model is designed to solve two-loop control problem according to the basic principle of ADRC. Compared with ALQR (Augmented Linear Quadratic Regulator) controller in full flight envelope, enough numerical verification tests about ADRC controller have carried out. Finally, simulation results show that the closed-loop system based on the ADRC decoupling controller has better disturbance rejection capability than the ones based on ALQR controller.

Trajectory Control of Miniature Helicopters Using a Unified Nonlinear Optimal Control Technique

It is always a challenge to design a real-time optimal full flight envelope controller for a miniature helicopter due to the nonlinear, underactuated, uncertain, and highly coupled nature of its dynamics. This paper integrates the control of translational, rotational, and flapping motions of a simulated miniature aerobatic helicopter in one unified optimal control framework. In particular, a recently developed real-time nonlinear optimal control method, called the θ-D technique, is employed to solve the resultant challenging problem considering the full nonlinear dynamics without gain scheduling techniques and timescale separations. The uniqueness of the θ-D method is its ability to obtain an approximate analytical solution to the Hamilton–Jacobi–Bellman equation, which leads to a closed-form suboptimal control law. As a result, it can provide a great advantage in real-time implementation without a high computational load. Two complex trajectory tracking scenarios are used to evaluate the control capabilities of the proposed method in full flight envelope. Realistic uncertainties in modeling parameters and the wind gust condition are included in the simulation for the purpose of demonstrating the robustness of the proposed control law.

Model Identification and Active Modeling Control for Rotor Fly-robot: Theory and Experiment

This paper proposes an active modeling control method, as well as a simplified reference model and a modified frequency identification process, to deal with the model mismatch, which brings errors to nominal controller. The hovering model is identified through the proposed semi-decoupled reference model and the modified frequency identification process. Based on this hovering model and adaptive set-membership filter (ASMF), a scheme of compensation for nominal controller is developed due to the unknown but bounded (UBB) model errors. Flight experiments are done on ServoHeli-40 fly-robot platform to verify the effectiveness of the proposed method in full flight envelope.

Estimation of Captive Flight Loads Using Grid-Free Euler Solver

NTEGRATION of a new weapon to a highly agile aircraft is an exigent task, as the weapon and its supporting systems should structurally endure during stern maneuvers of the aircraft. Fighter aircraft has maneuvering capability of about 9 g and, therefore, the support systems of flight vehicles (launcher, launcher rail, launch shoe, and adapter) on the fighter aircraft should be designed to withstand the loads experienced by the vehicle at critical maneuvering conditions. Guidelines for weapon-aircraft integration are provided through military specification MIL-STD-8591 [1]. The specification provides weapon design load requirements for unusual flightconditionsthatsometimeariseonspecificaircraft.Itisbasedon using actual or predicted aircraft flight performance data to examine the full flight envelope and generate consistent set of inertial and aerodynamic loads [2]. In the present work, aerodynamic characterization of the flight vehicle mounted on a fighter aircraft has been carried out using an in-house developed 3-D grid-free Euler solver [3]. Further, the numerical simulations are carried out to estimate the captive loads acting on the flight vehicle integrated with fighteraircraftatspecified flowconditionsandthemaximumloading conditions are identified. The load distribution is extracted at that condition to verify the structural integrity of the flight vehicle and supporting systems. Grid generation around a complex fighter aircraft is the most time consuming task in the numerical simulation and presence of flight vehicle poses a formidable challenge. A grid-free q-LSKUM [3] (entropyvariables-basedleast-squareskineticupwindmethod)Euler solver has been used for the computational fluid dynamics (CFD) analysis. The grid-free solver requires a cloud of points around the configuration and a set of neighbors around each point. The cloud of points was obtained by overlapping grids around fighter aircraft and flightvehicles.CADmodelsof flightvehicleand fighteraircraftwere imported, repaired, and cleaned up to obtain watertight solid models suitableforgridgeneration.Simpleunstructuredgridsaround fighter aircraft and flight vehicles are generated independently and then the grids are overlapped to get the distribution of points around the full configuration. An efficient preprocessor [4] has been developed and applied to generate the connectivity. A parallel version of the 3-D grid-free Euler solver q-LSKUM has been used to obtain the aerodynamic characteristics of the aerospace vehicle configuration. Inviscidsteady flowsimulationfor fighteraircrafthasbeencarried out using q-LSKUM code. The simulations at various Mach numbers, angles of attack, and sideslip angles have been carried out to obtain aerodynamic forces and moments acting on the integrated fighter aircraft with flight vehicles. The captive loads on all the four flight vehicles are computed at specified critical flow conditions.

Gain Scheduled Control: Switched Polytopic System Approach

Gain Scheduled Control: Switched Polytopic System Approach

Turbofan aero-engine full flight envelope control using H<SUB align=right>∞ loop-shaping control technique

A full flight envelope turbofan aero-engine control is developed by H ∞ loop-shaping control concept in this paper. First, H∞ loop-shaping control concept which is carried out in frequency domain is introduced and used to aero-engine control design. In order to meet the maximum robustness, multi-objective genetic algorithm is proposed creatively to obtain the global optimum loop-shaping weights. Second, considering the finite robustness of H ∞ loop-shaping controller while implementing the task of flight envelope turbofan engine control, the full flight envelope is divided to eight sub-regions and gain-scheduled control system where the sub-region controller could be switched with the change of the engine Mach number and altitude is constructed as the original work. The simulation results show the ability of the proposed control system in tracking the flight command, it is also verified that switching process has only slight effects on the engine condition under the control of the designed flight envelope controller.

Observations Regarding Algorithms Required for Robust CFD Codes

Over the last three decades Computational Fluid Dynamics (CFD) has gradually joined the wind tunnel and flight test as a primary flow analysis tool for aerodynamic designers. CFD has had its most favorable impact on the aerodynamic design of the high-speed cruise configuration of a transport. This success has raised expectations among aerodynamicists that the applicability of CFD can be extended to the full flight envelope. However, the complex nature of the flows and geometries involved places substantially increased demands on the solution methodology and resources required. Currently most simulations involve Reynolds-Averaged Navier-Stokes (RANS) codes although Large Eddy Simulation (LES) and Detached Eddy Suimulation (DES) codes are occasionally used for component analysis or theoretical studies. Despite simplified underlying assumptions, current RANS turbulence models have been spectacularly successful for analyzing attached, transonic flows. Whether or not these same models are applicable to complex flows with smooth surface separation is an open question. A prerequisite for answering this question is absolute confidence that the CFD codes employed reliably solve the continuous equations involved. Too often, failure to agree with experiment is mistakenly ascribed to the turbulence model rather than inadequate numerics. Grid convergence in three dimensions is rarely achieved. Even residual convergence on a given grid is often inadequate. This paper discusses issues involved in residual and especially grid convergence.

Linear, parameter-varying control of a supercavitating vehicle

A systematic approach to parameter-dependent control synthesis of a high-speed supercavitation vehicle (HSSV) is presented. The aim of the control design is to provide robust reference tracking across a large flight envelope, while directly accounting for the interaction of liquid and gas phases with the vehicle. A nonlinear dynamic HSSV model is presented and discussed relative to the actual vehicle. A linear, parameter-varying (LPV) controller is synthesized for angle rate tracking in the presence of model uncertainty. The control design takes advantage of coupling in the governing equations to achieve improved performance. Multiple LPV controllers synthesized for smaller overlapping regions of the parameter space are blended together, providing a single controller for the full flight envelope. Time-domain simulations implemented on high-fidelity simulations, provide insight into the performance and robustness of the proposed scheme.

Multi-Timescale Nonlinear Robust Control for a Miniature Helicopter

A new nonlinear control approach, which is applied to a miniature aerobatic helicopter through a multi-timescale structure, is proposed. Because of the highly nonlinear, unstable, and underactuated nature of a miniature helicopter, it is a challenge to design an autonomous flight control system that is capable of operating in the full flight envelope. To deal with unstable internal dynamics, the translational, rotational, and flapping dynamics of the helicopter (eleven degrees of freedom) are organized into a three-timescale, nonlinear model. The concepts of dynamic inversion and sliding manifold are combined together such that 1) the controller proposed is robust with respect to functional and parametric uncertainties, and 2) the settling time in faster modes is guaranteed to be less than the fixed step size of slower modes. A time-varying feedback gain, derived according to global stability and sliding manifold variations, is proved to be uniquely solvable based on the Perron-Frobenius Theorem. Partial uncertainties are explicitly taken into account in the nonlinear robust control design, and Monte Carlo simulations are used for validations under other sensor noises, model uncertainties, and a Federal Aviation Administration suggested gust condition.

Development of a Full Flight Envelope Helicopter Simulation Using System Identification

Development of a Full Flight Envelope Helicopter Simulation Using System Identification

Stability Analysis of Switched Linear Systems with Locally Overlapped Switching Law

Combining the common Lyapunov function method and the dwell time method, we propose stability analysis tools for switched linear systems evolving on the locally overlapped switching law. These systems exist naturally in many aerospace applications, such as separation dynamical systems involving rapid and discrete variations that are specified by the predetermined separation sequence. We show that switched linear systems with locally overlapped switching law are globally asymptotically stable provided that the dwell time on some subsystems is no smaller than a fixed positive constant. This result is also extended to the case of combination of the common Lyapunov function method and the average dwell time method. We validate the proposed stability analysis tools by an example and illustrate an application to full-envelope flight controller design.

Design and implementation of a hardware-in-the-loop simulation system for small-scale UAV helicopters

We present in the paper the design of a hardware-in-the-loop simulation framework and its actual implementation on our custom constructed unmanned-aerial-vehicle (UAV) helicopter systems. Real-time hardware-in-the-loop simulation is one of the most effective methods for the verification of the overall control performance and safety of the UAVs before conducting actual flight tests. In our proposed framework, four modules, which include onboard hardware, flight control, ground station and software, are integrated together to realize the hardware-in-the-loop simulation. This design is successfully utilized for simulating several flight tests including basic flight motions, full-envelope flight and multiple UAV formation flight. Results obtained show that the constructed hardware-in-the-loop simulation system is highly effective and useful.

Design of Flight Control System for a Small Unmanned Tilt Rotor Aircraft

A tilt rotor is an aircraft of a special kind, which possesses the characteristics of a helicopter and a fixed-wing airplane. However, there are a great number of important technical problems waiting for settlements. Of them, the flight control system might be a critical one. This article presents the progresses of the research work on the design of flight control system at Nanjing University of Aeronautics and Astronautics (NUAA). The flight control law of the tilt rotor aircraft is designed with the help of an inner/outer loop control structure and an eigenstructure assignment algorithm on the basis of a proper mathematical model already verified by the wind tunnel tests. The proposed control law has been born out through the construction of the flight control system and the flight tests. Now, the flight tests are still underway on a prototype of small unmanned tilt rotor aircraft. The results have evidenced the credibility of the aircraft design and the effectiveness of the flight control system for the tilt rotor working in the helicopter mode. A full envelope flight test is planned to carry out further researches on the flight control law.

Flight control of a rotary wing UAV using backstepping

This paper presents a novel application of backstepping controller for autonomous landing of a rotary wing UAV (RUAV). This application, which holds good for the full flight envelope control, is an extension of a backstepping algorithm for general rigid body velocity control. The nonlinear RUAV model used in this paper includes the flapping and servo dynamics. The backstepping-based controller takes advantage of the ‘decoupling’ of the translation and rotation dynamics of the rigid body, resulting in a two-step procedure to obtain the RUAV control inputs. The first step is to compute desired thrusts and flapping angles to achieve the commanded position and the second step is to compute control inputs, which achieve the desired thrusts and flapping angles. This paper presents a detailed analysis of the inclusion of a flapping angle correction term in control. The performance of the proposed algorithm is tested using a high-fidelity RUAV simulation model. The RUAV simulation model is based on miniature rotorcraft parameters. The closed-loop response of the rotorcraft indicates that the desired position is achieved after a short transient. The Eagle RUAV control inputs, obtained using high-fidelity simulation results, clearly demonstrate that this algorithm can be implemented on practical RUAVs. Copyright © 2009 John Wiley & Sons, Ltd.