Behavior Of Aircraft Research Articles

Predictive Control with Parameter Adaptation to Achieve α-Protection in the RECONFIGURE Benchmark in the Presence of Icing

Undetected ice accretions on aerodynamic surfaces can deeply change the dynamic behaviour of aircraft, leading to poor performance of the automatic control systems. This is characterised by oscillatory behaviour, and overshoot of setpoints and ight envelope protection values. To mitigate this undesirable behaviour, this paper applies predictive control in combination with online estimation of the first and second order partial derivatives of the lift and pitching moment coeffcients with respect to the angle of attack using an Extended Kalman Filter, to achieve a constrained indirect adaptive ight control law. The design is evaluated on the RECONFIGURE benchmark, which is a nonlinear, high fidelity, industrially validated simulator of a large Airbus aircraft. In icing scenarios at high incidence, the resulting trajectories are shown to be better damped and more compliant with constraints when compared to a predictive control law employing only linear disturbance estimation.

Finite element modeling of wing bird strike

Bird strike events are a potential dangerous problem faced by flight safety nowadays. These events are usually simulated through numerical approaches due mainly to the technical difficulties and high costs associated with experimental tests. The consequences of bird impact can be severe and, therefore, the aircraft components have to be certified for a proven level of bird impact resistance before being put into service. Bird impact poses serious threats to military and civilian aircrafts as they lead to fatal structural damage to critical aircraft components. The exposed aircraft components such as windshields, radomes, leading edges, engine structure, and blades are vulnerable to bird strikes. Among a large number of structural tests an aircraft structure needs a certification requirement for a proven level of impact resistance against bird impacts. Bird strike experiments are very expensive and henceforth explicit numerical modeling techniques have grown importance. Parts of aircraft construction are intensively vulnerable to damage during flight by bird impact. The theoretical approach and results of numerical simulations of dynamic response of the wing loaded by the bird impact are presented. The numerical simulation is carried out using smooth particle hydrodynamics (SPH) method running in the nonlinear explicit finite element code ANSYS AUTODYN. The focus is given to the validation of the stress, strain and deflection of wing on the impact zone. The dependency of given parameters on the variation of ellipsoidal bird aspect ratio, impact velocity and wing design details was discussed. As well, some results of experimental data were given. The behavior of aircraft leading edge against high speed bird impact was successfully simulated and the effect of various parameters on its dynamic response was studied. Numerical model was validated with published experimental results and further employed to investigate the influence of various parameters on dynamic behavior of aircraft leading edge.

A predictive envelope protection system using linear, parameter-varying models

A parameter-varying, model-predictive envelope protection system is developed simplifying the controller structures required to keep the aircraft within a safe angle-of-attack and normal load factor envelope. The idea of a quasi-steady flight condition is used to map the flight envelope limits onto the setpoint values of a single flight control law. Since no mode switching is required, the selected level of automation, i.e., autopilot and flight management functionalities, is independent from the proximity to any angle-of-attack and normal load factor limit. In contrast to previous approaches, the proposed algorithm makes use of a quasi-linear, parameter-varying control loop model to adapt to the true nonlinear aircraft behavior. A variance-based sensitivity analysis highlights the most significant scheduling variables within this control loop model and, therefore, indicates the option of model reduction and improvement in efficiency, respectively. The proposed envelope protection system is evaluated throughout virtual flight tests with the unmanned flight test platform ULTRA-Dimona showing promising overall performance also in the presence of wind and turbulence.

English

Velocity of projectiles plays an important role in determining the response and behaviour of a target under impact. In this project metallic canister with a soft projectile, which is one of the critical dynamic components, is used to launch bird/ice pellet, to study its influence on the behaviour of aircraft and aero-engine parts. This is done by using conventional compressed air launcher which launches the projectile velocities up to 200 m/s. The canisters which house the test projectile collapse against canister arrester located at the delivery end of the launcher. The collapse characteristics of the canisters are dependent on the geometry of the tip of the arresters, material of the canisters, the shape of the canisters and the heat treatment process adapted for canisters. The impact load of the canister on the arrester can be up to several tonnes axially. This force is a function of the collapsibility of the canister at impact and in turn it is determined by the operating pressure of the launcher. It is therefore necessary to develop a velocity measuring system, which determines the dynamic axial loads on the launcher as well as the target plate, which in turn helps in studying the collapsibility parameter of the canister itself. In order to measure the high velocity of the projectile three measuring devices, namely, optical laser sensor, based on programmable logic device, copper wire and graphite rod based devices are developed and used. To provide adequate redundancy for the complex, highly dynamic and costly impact tests, two of these three velocity measuring devices are used simultaneously: (1) copper wire based sensing system and (2) graphite rod based sensing system. The measuring device input parameter is fabricated on a gun barrel with acrylic fixture at two positions at a distance of 1 m before the arrester where the canister strikes and collapses. At initial and final position, optical transmitter and receiver give the delay time when the canister passes through it by obstructing the sensing signal path. The copper wire circuit based device gives open circuit delay between the two points when canister breaks it. The time taken to travel 1m distance by the canister in the gun barrel is obtained from the device and the velocity of the canister or projectile is experimentally evaluated.

Geometrically Nonlinear Aeroelastic Scaling for Very Flexible Aircraft

High-aspect-ratio wings present in very flexible aircraft can undergo large deformations, which results in significant changes in natural frequencies as well as in static and dynamic aeroelastic response. This geometric nonlinear behavior becomes an integral part of any aeroelastic analysis to be conducted in such class of vehicles. Aeroelastic scaling is an important way to study the aeroelastic behavior of aircraft, and it is an integral part in risk mitigation for aircraft development. However, the current aeroelastic scaling methodologies have focused on geometrically linear structures. This paper demonstrates a methodology for geometrically nonlinear aeroelastic scaling of very flexible aircraft. The known linear scaling factors and similarity rules are extended to address geometrically nonlinear aeroelastic scaling. A high-aspect-ratio flying wing in free flight is taken as an example to verify the new scaling procedure, and numerical studies are conducted using the University of Michigan’s Nonlinear Aeroelastic Simulation Toolbox. Numerical results support the new approach for aeroelastic scaling of very flexible aircraft.

Intelligent Pitch Controller Identification And Design

This paper exhibits a comparative assessment based on time response specification performance between modern and classical controller for a pitch control system of an aircraft system. The dynamic modeling of pitch control system is considered on the design of an autopilot that controls the pitch angle. It starts with a derivation of a suitable mathematical model to describe the dynamics of an aircraft. For getting close to actual conditions the white noise disturbance is applied to the system. In this paper, it is assumed that the model pitch control system is not available. So using the identification system and Box-Jenkins model estimator we identify the pitch control system. System’s identification is a procedure for accurately characterizing the dynamic response behavior of a complete aircraft, of a subsystem, or of an individual component from measured data. To study the effectiveness of the controllers, the LQR Controller and PID Controller and fuzzy controller is developed for controlling the pitch angle of an aircraft system. Simulation results for the response of pitch controller are presented instep’s response. Finally, the performances of pitch control systems are investigated and analyzed based on common criteria of step’s response in order to identify which control strategy delivers better performance with respect to the desired pitch angle. It is found from simulation that the fuzzy controller gives the best performance compared to PID and LQR controller.

Research and Simulation of an Aircraft Behavior during U-Turn Maneuver

U-turn maneuver belongs to the special ground taxiing behavior, which are not accurately demonstrated in current visual simulation system. This paper presents a method to address computational problems of the position and attitude during the special ground maneuvers. From the perspective of simulation of airport tower control, aircraft movement calculation method constrained by nonholonomic kinematics are proposed to describe accurately the U-turn maneuvers. According to actual movement and nonholonomic path following model, an algorithm suitable for simulation-driven calculation is designed. In addition, under the open-loop control, aircrafts U-turn maneuver algorithm constrained by runway geometry and steering angle variation is designed to describe completely the behavior. Compared with dynamics model, this method we proposed have advantages that are easier to be calculated and more suitable for visual simulation. Above all, It is proved that the method is valid through simulation with aircraft geometry data.

Upset Dynamics of an Airliner Model: A Nonlinear Bifurcation Analysis

Despite the significant improvement in safety linked to the fourth generation of airliners, the risk of encountering upset conditions remains an important consideration. Upset, which may arise from faults, external events, or inappropriate pilot inputs, can induce a loss-of-control incident if the pilot does not respond in the correct manner. Any initiative aimed at preventing such events requires an understanding of the fundamental aircraft behavior. This paper presents the use of bifurcation analysis, complemented by time-history simulations, to understand the flight dynamics of the open-loop NASA generic transport model by identifying the attractors of the dynamical system that govern upset behavior. A number of drivers for potential upset conditions have been identified, including nonoscillatory spirals and oscillatory spins. The analysis shows that these spirals and spins are connected in two-parameter space and that, by an inappropriate pilot reaction to the spiral, it is possible to enter the oscillatory spin.

Ship underwater noise assessment by the acoustic analogy, part III: measurements versus numerical predictions on a full-scale ship

The acoustic analogy represents a powerful and versatile approach, able to numerically predict the noise generated by a body moving in a fluid. It is widely used to provide essential indications about the aeroacoustic behavior of aircraft and helicopters (even at a design stage) and, eventually, to pursue effective strategies aimed at desirable reduction and/or control of noise. Nevertheless, applications in the area of hydroacoustics and in the prediction of ship underwater noise are very rare. In this paper, the potential of the acoustic analogy is directly tested on a large ferry, for which a measurement campaign at sea was performed. In spite of the complexity of the tested configuration [the ship mounts two contracted and loaded tip (CLT) propellers located ahead of two rudders, and its hull is characterized by a rather elongated skeg] and the many variables not taken into account in the numerical simulation (such as the contribution from machinery noise and the probable occurrence of tip vortex cavitation), the agreement between the measured and computed noise spectra is quite satisfactory. The analysis suggests many interesting features of the ship hydroacoustic field: the dominant role played by nonlinear sources far from the body and the relevance of scattering effects from the hull surface. Furthermore, the scattered pressure seems to contribute to alter the frequency content of the resulting signatures with respect to the blade passage frequencies. Finally, an overview of future developments and applications of this numerical approach for marine/maritime problems is presented.

High Level Design of the Flight Control Software for Small VTOL Unmanned Aerial Vehicle

The implementation of an on-board flight control system must be preceded by extensive design activity. The software implementation, test and maintenance effort might be sig- nificantly reduced by using carefully designed software. The design procedure itself is also a complex activity, several requirements, constraints must be considered in order to generate an appropriate, well detailed and easily implementable design. The design task is even more complicated when the aircraft has some special features like the capability of vertical take-off and landing. In this case the software must be able to handle the di- versity of flight modes where the aircraft behaviour in each mode is unique. This article presents the main aspects of the design procedure.

Validation of mathematical models for helicopter flight simulators past, present and future challenges

AbstractAt the heart of a flight simulator resides the mathematical representation of aircraft behaviour in response to control inputs, atmospheric disturbances and system inputs including failures and malfunctions. While this mathematical model can never be wholly accurate, its fidelity, in comparison with real world behaviour, underpins the usefulness of the flight simulator. The present paper examines the state of the art achieved in validating mathematical models for helicopter simulators, addressing the strengths and weaknesses of the present European standard for the qualification of helicopter flight simulators, JAR FSTD-H (previously JAR-STD-1H/2H/3H). Essential questions are examined, such as: What is the required model fidelity to guarantee a simulation is sufficiently representative to be fit for purpose? Are the tolerances set in the current standards fine enough that they lead to only minor changes in handling qualities? What is an acceptable tuning process for the simulation? What is the effect of modelling fidelity on the overall pilot control strategy? What is the relationship between the settings of the simulator cueing environment and the behaviour of the pilot? What is the industrial experience on qualification of flight simulators that might usefully inform developments? Many of these questions were addressed in Europe in a previous GARTEUR Action Group (AG) HC/AG-12 the results of which are documented in this paper. Solutions are proposed for improving the current JAR-FSTD standard with respect to validation of mathematical models.

SIMULATION, PARAMETER IDENTIFICATION AND CONTROL SYSTEM DESIGN OF AN AIRCRAFT USING UNIFIED MATHEMATICAL MODEL

This paper discusses the role of mathematical model of an aircraft dynamics in flight mechanics related studies. A flight dynamics mathematical model represents the motion of a flight vehicle in six-degree-of-freedom. All forces and moments governing the flight motion are specifically represented in the equations, such as inertial force and moment, aerodynamics force and moment, propulsive force and moment, flight control force and moment, and external disturbances. In addition, the mathematical model also includes a set of kinematic and navigation equations, to provide a thorough representation of vehicle motions and movements in the air. Another way to obtain the mathematical representation of an aircraft dynamic is by employing system identification procedure, on which the measured input-output data of the system is analyzed and processed to obtain the parameters of the system. The mathematical model also can be used for simulating and analyzing the dynamic behavior of an aircraft. The simulation is usually conducted by implementing the model into numerical simulation software, such as MATLAB/Simulink. Another implementation of the unified mathematical model presented in this paper is for designing flight control system. In this paper, an aircraft such as Unmanned Aerial Vehicle (UAV) system will be used as a basis for demonstrating how mathematical model an aircraft can be obtained, and how it is further exploited for some analysis and studies in flight mechanics fields.

Nested Modeling of Hazards in the National Airspace System

Highly complex and coupled nature of the National Airspace System makes it challenging to understand the safety implications of changes to its operations. A comprehensive assessment of the hazards associated with the introduction of new technologies must involve modeling of interactions among aircraft behavior, supporting infrastructure, and the operational procedures. The traditional approach to risk assessment does not focus on the relative timing of relevant events. The likelihoods of those events are computed externally, usually by means of physics-based simulations, which, while increasingly realistic in capturing physical phenomena, are limited to describing only few relevant interactions to keep the overall complexity tractable. This paper discusses an intermediate layer of analysis with enough fidelity to capture time-dependent coupling among relevant entities of the system that is compact enough to track a large number of those relevant entities simultaneously. To this end, Stochastic Petri nets are used to analyze the conflict resolution between the merging flows of air traffic for optimized-profile-descent approaches. Stochastic Petri nets are coupled with agent-based simulation, and the efficiency of the merging procedures and their sensitivity to wind conditions and the traffic patterns are analyzed. Finally, compact modeling of two off-nominal scenarios related to future procedures is discussed.

Investigation on the Effect of Airspeed and Altitude to Phugoid Mode of a Small Unmanned Blended Wing-Body Aircraft with Canard as a Longitudinal Control Surface

Phugoid mode is a lowly damped, low-frequency oscillatory motion representing vertical translation usually related to kinetic and potential energy interchange. MIL-F-8785C standard has ruled out qualitative specification requirements on measurement of flying and handling qualities of piloted aircraft. For phugoid mode, these requirements lie in the value of its damping ratio. Small aircraft is sensitive to atmospheric conditions and poor phugoid mode performance is observed in many unmanned aircraft. This paper discusses the effect of airspeed and altitude to phugoid mode of small, unmanned blended wing-body (BWB) aircraft named Baseline-II E-2. Baseline-II is a low subsonic, remotely-piloted UAV used to study the behaviour of a BWB-type aircraft. The case presented here is an E-2 version in which a specifically-designed canard is incorporated as its longitudinal control surface. Five Category B flight cases (airspeeds) per altitude-case, and three altitude cases (low, medium and high) are studied. Model-N dynamic model is introduced here to become the basis of flight simulation. The model is compared to models derived by other authors and approximation equations. The mean of simulating phugoid behaviour is using state-space representation of the aircraft using Matlab SIMULINK. The computations show that Baseline-II E-2 undamped natural frequency of phugoid mode is inversely-proportional to airspeed and reduces as altitude increases. These have adverse effect on its damping ratio that increases near parabolically when the aircraft flies faster, and reduces when it climbs up. The cause of these trends is looked into in detail and issues concerning Baseline-II E-2’s unsatisfactory and unstable phugoid mode oscillation at low speed are addressed.

Stability Analysis of a Light Aircraft Configuration Using Computational Fluid Dynamics

This work aims to simulate and study the flow field around SAFAT-01 aircraft using numerical solution based on solving Reynolds Averaged Navier-Stokes equations coupled with K-ω SST turbulent model. The aerodynamics behavior of SAFAT-01 aircraft developed at SAFAT aviation complex were calculated at different angles of attack and side slip angles. The x,y and z forces and moments were calculated at flight speed 50m/s and at sea level condition. Lift and drag curves for different angles of attack were plotted. The maximum lift coefficient for SAFAT-01 was 1.67 which occurred at angle of attack 16° and Maximum lift to drag ratio (L/D) was 14 which occurred at α=3°, and the zero lift drag coefficient was 0.0342. Also the yawing moment coefficient was plotted for different side slip angles as well as rolling moment. The longitudinal stability derivatives with respect to angle of attack, speed variation (u), rate of pitch (q) and time rate of change of angle of attack were calculated using obtained CFD results. Concerning lateral stability only side slips derivatives were calculated. To validate this numerical simulation USAF Digital DATCOM is used to analyze this aircraft; a comparison between predicted results for this aircraft and Digital DATCOM indicated that this numerical simulation has high ability for predicting the aerodynamics characteristics.

Analytical Volterra-based models for nonlinear low order flight dynamics approximation systems

AbstractAnalytical methodology is presented to conduct dynamical assembly of simple low order nonlinear responses for system synthesis and prediction using Volterra theory. The procedure is set forth generically and then applied to several atmospheric flight examples. A two-term truncated Volterra series, which is enough to capture the quadratic and bilinear nonlinearities, is developed for first and second order generalised nonlinear single degree of freedom systems. The resultant models are given in the form of first and second kernels. A parametric study of the influence of each linear and nonlinear term on kernel structures is investigated. A step input is then employed to quantify and qualify the nonlinear response characteristics. Uniaxial surge and pitch motions are presented as examples of the low order flight dynamic systems. These examples show the ability of the proposed analytical Volterra-based models to predict, understand, and analyse the nonlinear aircraft behaviour beyond that attainable by linear-based models. The proposed analytical Volterra-based model offers an efficient nonlinear preliminary design tool in qualifying the aircraft responses before computer simulation is available or invoked.

FDR data as a source for flying events visualisation and aircraft parameter estimation

The hundreds of parameters that are recorded by Flight Data Recorders (FDR) during flight facilitates the investigation of incidents and also provides the opportunity to use the recorded data to predict future aircraft behaviour. Current researches are oriented in the direction of data detection responsible for a typical behavior of flight, data reconstruction and flight simulation. This paper presents a mathematical model for aircraft used in flight data interpretation, and an application that can give the graphical representation of any flight data parameter. Also, using a preprocessed data set, application produces a flash animation of the aircraft flight.

Cruising Autopilot for a Flexible Aircraft with an Internal Loop of Model Following

AbstractA cruising autopilot is developed for a nonlinear flexible aircraft based on the model-following method. The autopilot controller, called the outer loop, is designed for a rigid body approximation of the flexible aircraft. The model-following control, named the inner loop, approximates the input-output behavior of the flexible aircraft to a rigid body. The autopilot is applied to the flexible aircraft with a model-following inner loop. This inner loop controls the flexible modes and the outer loop, which are the rigid body dynamics. The model-following control is a suboptimal linear quadratic formulation with output feedback. This technique can stabilize the flexible modes and keep their amplitudes acceptable. The autopilot is a speed and altitude tracker with a wing leveler. It is designed via a robust H∞ with static output feedback. An aircraft model with nine flexible modes is used to evaluate the method. It is a conceptual medium size jet, like an Embraer 190/190 or Boeing 737–200/300. Simulat...

9.2.2 Observations from applying SE principles on a large research project developing Processes, Methods & Tools for Modelling & Simulation of Aircraft Behaviour

AbstractThis paper describes experiences gained by applying System Engineering (SE) management practices within a particular EU research project (CRESCENDO). Additionally, we attempt to answer the underlying question: to what extent is it practical to apply SE management practices within the context of large‐scale, collaborative research projects? A common characteristic of these projects is that their structure is quite complex and requires a systematic management approach for their successful outcome. The benefits of managing industrial projects using SE management practices are well‐known. However, difficulties arise not only from the maturity of the SE processes, methods & tools available, but also from selecting the right toolsets from the plethora on offer. Therefore, significant investments are needed to develop methodologies to navigate the complexity of processes and products in the aerospace sector. In this paper, we highlight the advantages of the SE approach, but also point out some drawbacks and provide some recommendations for future adopters of SE management practices in similar project contexts.

Structural and Aeroelastic Design of a Joined-Wing UAV

The Italian Aerospace Research Center (CIRA) is currently designing an unmanned aerial research system that is lightweight and has high-structural flexibility, code named the High Altitude Performance Demonstrator (HAPD).The project is framed within the Italian Aerospace Research Program, under the Unmanned Aerial Vehicle (UAV) Chapter. This unmanned aerial system is mainly aimed at developing and validating advanced modeling methodologies for flexible aircrafts. A compendium of the system is provided in this paper, together with a deeper discussion of how CIRA developed the structural and aeroelastic design of HAPD. Some experimental tests performed to validate the main concepts are also presented. The vehicle has an unconventional joined-wing configuration that mitigates undesired extreme flexibility, but that results in a more complicated design. First, aeroelasticity has been taken into account from the preliminary stages of design because flexibility significantly affects aircraft behavior. Second, the HAPD structure is redundant with regard to constraints (because of its joined wing), thus making the internal forces dependent on the stiffness distribution. For these reasons, the availability of an integrated methodology that can support the structural design is mandatory. The output of such a methodology consists of primary structure stiffness distributions (fuselage, wings, and vertical tail), compatibly with the absence of any aeroelastic instability, and structural failure under operative loads.