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1 – 10 of 11Mahdi Fatehi, Majid Moghaddam and Mohammad Rahim
The purpose of this paper is to present a novel approach in aeroservoelastic analysis and robust control of a wing section with two control surfaces in leading‐edge and…
Abstract
Purpose
The purpose of this paper is to present a novel approach in aeroservoelastic analysis and robust control of a wing section with two control surfaces in leading‐edge and trailing‐edge. The method demonstrates how the number of model uncertainties can affect the flutter margin.
Design/methodology/approach
The proposed method effectively incorporates the structural model of a wing section with two degrees of freedom of pitch and plunge with two control surfaces on trailing and leading edges. A quasi‐steady aerodynamics assumption is made for the aerodynamic modeling. Basically, perturbations are considered for the dynamic pressure models and uncertainty parameters are associated with structural stiffness and structural damping and are accounted for in the model by a Linear Fractional Transformation (LFT) model. The control commands are applied to a first and second order electro‐mechanical actuator.
Findings
Dynamic performance of aeroelastic/aeroservoelastic system including time responses, system modal specifications, critical flutter speeds, and stability margins are extracted and compared with each other. Simulation results are validated through experiments and are compared to other existing methods available to the authors. Results of simulations with four structural uncertainties and first order controllers have a good agreement with experimental test results. Furthermore, it is shown that by using a high gain second order controller, the aeroservoelastic (ASE) system does not have any coupling nature in frequency response.
Originality/value
In this study, modeling, simulation, and robust control of a wing section have been investigated utilizing the μ‐Analysis method and the wing flutter phenomenon is predicted in the presence of multiple uncertainties. The proposed approach is an advanced method compared to conventional flutter analysis methods (such as V‐g or p‐k) for calculating stability margin of aeroelastic/aeroservoelastic systems.
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Mustafa Tolga Tolga Yavuz and İbrahim Özkol
This study aims to develop the governing differential equation and to analyze the free vibration of a rotating non-uniform beam having a flexible root and setting angle for…
Abstract
Purpose
This study aims to develop the governing differential equation and to analyze the free vibration of a rotating non-uniform beam having a flexible root and setting angle for variations in operating conditions and structural design parameters.
Design/methodology/approach
Hamiltonian principle is used to derive the flapwise bending motion of the structure, and the governing differential equations are solved numerically by using differential quadrature with satisfactory accuracy and computation time.
Findings
The results obtained by using the differential quadrature method (DQM) are compared to results of previous studies in the open literature to show the power of the used method. Important results affecting the dynamics characteristics of a rotating beam are tabulated and illustrated in concerned figures to show the effect of investigated design parameters and operating conditions.
Originality/value
The principal novelty of this paper arises from the application of the DQM to a rotating non-uniform beam with flexible root and deriving new governing differential equation including various parameters such as rotary inertia, setting angle, taper ratios, root flexibility, hub radius and rotational speed. Also, the application of the used numerical method is expressed clearly step by step with the algorithm scheme.
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L. Djayapertapa and C.B. Allen
Transonic flutter and active flap control, in two dimensions, are simulated by coupling independent structural dynamic and inviscid aerodynamic models, in the time domain. A…
Abstract
Transonic flutter and active flap control, in two dimensions, are simulated by coupling independent structural dynamic and inviscid aerodynamic models, in the time domain. A flight control system, to actively control the trailing edge flap motion, has also been incorporated and, since this requires perfect synchronisation of fluid, structure and control signal, the “strong” coupling approach is adopted. The computational method developed is used to perform transonic aeroelastic and aeroservoelastic calculations in the time domain, and used to compute stability (flutter) boundaries of 2D wing sections. Open and closed loop simulations show that active control can successfully suppress flutter and results in a significant increase in the allowable speed index in the transonic regime. It is also shown that active control is still effective when there is free‐play in the control surface hinge. Flowfield analysis is used to investigate the nature of flutter and active control, and the fundamental importance of shock wave motion in the vicinity of the flap is demonstrated.
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Fathi Jegarkandi Mohsen, Salezadeh Nobari Ali, Sabzehparvar Mahdi, Haddadpour Hassan and Tavakkoli Farhad
The purpose of this paper is to investigate the aeroelastic behavior of a supersonic flight vehicle flying at moderate angles of attack using global analytic nonlinear aerodynamic…
Abstract
Purpose
The purpose of this paper is to investigate the aeroelastic behavior of a supersonic flight vehicle flying at moderate angles of attack using global analytic nonlinear aerodynamic model.
Design/methodology/approach
Aeroelastic behavior of a supersonic flight vehicle flying at moderate angles of attack is considered, using nonlinear aerodynamics and linear elastodynamics and structural models. Normal force distribution coefficient over the length of the vehicle and pitching moment coefficient are the main aerodynamic parameters used in the aeroelastic modeling. It is very important to have closed form analytical relations for these coefficients in the model. They are generated using global nonlinear multivariate orthogonal modeling functions in this work. Angle of attack and length of the vehicle are selected as independent variables in the first step. Local variation of angle of attack is applied to the analytical model and due to its variation along the body of the vehicle, equations of motion are finalized. Mach number is added to the independent variables to investigate its role on instability of the vehicle and the modified model is compared with the previous one in the next step. Thrust effect on the aeroelastic stability of the vehicle is analyzed at final stage.
Findings
It is shown that for the vehicles having simple configurations and low length to diameter ratios flying at low angles of attack, assuming normal force distribution coefficient linear relative to α is reasonable. It is concluded that vehicle's velocity and thrust has not great effect on its divergence dynamic pressure.
Originality/value
Based on the constructed model, a simulation code is generated to investigate the aeroelastic behavior of the vehicle. The resultant code is verified by investigating the static aeroelastic stability margin of the vehicle presented in the references. Mach number effect on the aeroelastic behavior of the vehicle is considered using modified aerodynamic model and is compared with the results. Data base for identifying aerodynamic coefficients is conducted using CFD code.
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Saeed Shamaghdari and S.K.Y. Nikravesh
The purpose of this paper is to present a nonlinear model along with stability analysis of a flexible supersonic flight vehicle system.
Abstract
Purpose
The purpose of this paper is to present a nonlinear model along with stability analysis of a flexible supersonic flight vehicle system.
Design/methodology/approach
The mathematical state space nonlinear model of the system is derived using Lagrangian approach such that the applied force, moment, and generalized force are all assumed to be nonlinear functions of the system states. The condition under which the system would be unstable is derived and when the system is stable, the region of attraction of the system equilibrium state is determined using the Lyapunov theory and sum of squares optimization method. The method is applied to a slender flexible body vehicle, which is referenced by the other researchers in the literature.
Findings
It is demonstrated that neglecting the nonlinearity in external force, moment and generalized force, as it was assumed by other researchers, can cause significant variations in stability conditions. Moreover, when the system is stable, it is shown analytically here that a reduction in dynamic pressure can make a larger region of attraction, and thus instability will occur in a larger angle of attack, greater angular velocity and elastic displacement.
Practical implications
In order to carefully study the behavior of aeroelastic flight vehicle, a nonlinear model and analysis is definitely necessary. Moreover, for the design of the airframe and/or control purposes, it is essential to investigate region of attraction of equilibrium state of the stable flight vehicle.
Originality/value
Current stability analysis methods for nonlinear elastic flight vehicles are unable to determine the state space region where the system is stable. Nonlinear modeling affects the determination of the stability region and instability condition. This paper presents a new approach to stability analysis of the nonlinear flexible flight vehicle. By determining the region of attraction when the system is stable, it is demonstrated analytically, in this research, that decreasing the dynamic pressure can produce larger region of attraction.
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Promio Charles F., Raja Samikkannu, Niranjan K. Sura and Shanwaz Mulla
Ground vibration testing (GVT) results can be used as system parameters for predicting flutter, which is essential for aeroelastic clearance. This paper aims to compute GVT-based…
Abstract
Purpose
Ground vibration testing (GVT) results can be used as system parameters for predicting flutter, which is essential for aeroelastic clearance. This paper aims to compute GVT-based flutter in time domain, using unsteady air loads by matrix polynomial approximations.
Design/methodology/approach
The experimental parameters, namely, frequencies and mode shapes are interpolated to build an equivalent finite element model. The unsteady aerodynamic forces extracted from MSC NASTRAN are approximated using matrix polynomial approximations. The system matrices are condensed to the required shaker location points to build an aeroelastic reduced order state space model in SIMULINK.
Findings
The computed aerodynamic forces are successfully reduced to few input locations (optimal) for flutter simulation on unknown structural system (where stiffness and mass are not known) through a case study. It is demonstrated that GVT data and the computed unsteady aerodynamic forces of a system are adequate to represent its aeroelastic behaviour.
Practical implications
Airforce of every nation continuously upgrades its fleet with advanced weapon systems (stores), which demands aeroelastic flutter clearance. As the original equipment manufacturers does not provide the design data (stiffness and mass) to its customers, a new methodology to build an aeroelastic system of unknown aircraft is devised.
Originality/value
A hybrid approach is proposed, involving GVT data to build an aeroelastic state space system, using rationally approximated air loads (matrix polynomial approximations) computed on a virtual FE model for ground flutter simulation.
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I. Ursu, G. Tecuceanu, F. Ursu, T. Sireteanu and M. Vladimirescu
A method of designing the controller to solve the robust servomechanism problem is applied in the case of an electrohydraulic servo actuating primary flight control. This method…
Abstract
A method of designing the controller to solve the robust servomechanism problem is applied in the case of an electrohydraulic servo actuating primary flight control. This method is based on the well‐known solution consisting of two separate devices: a servocompensator, in fact an internal model of the exogenous dynamics, including the reference commands and disturbance signals; and a stabilizing compensator. The proof is made involving the servocompensator structure which is close to the one designed for step signals. The stabilizing compensator is assured by way of a linear quadratic optimal procedure. An antiwindup compensation is added to deal with the adverse effects caused by actuator saturation.
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Ioan Ursu, Felicia Ursu and Lucian Iorga
Presents a switching type neuro‐fuzzy control synthesis. The control algorithm supposes as a component part a neurocontrol designed to optimize a performance index. Whenever the…
Abstract
Presents a switching type neuro‐fuzzy control synthesis. The control algorithm supposes as a component part a neurocontrol designed to optimize a performance index. Whenever the neurocontrol saturates or a certain performance parameter of the system decreases, the scheme of control switches to a feasible and reliable fuzzy logic control. Describes the procedure of return to the optimizing neurocontrol which is essential. This methodology of control synthesis ensures antisaturating, antichattering and robustness properties of the controlling system, as illustrated by numerical simulation in the case of a primary flight controls electrohydraulic servo actuator
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Rossana Fernandes, Benyang Hu, Zhichao Wang, Zheng Zhang and Ali Y. Tamijani
This paper aims to assess the feasibility of additively manufactured wind tunnel models. The additively manufactured model was used to validate a computational framework allowing…
Abstract
Purpose
This paper aims to assess the feasibility of additively manufactured wind tunnel models. The additively manufactured model was used to validate a computational framework allowing the evaluation of the performance of five wing models.
Design/methodology/approach
An optimized fighter wing was additively manufactured and tested in a low-speed wind tunnel to obtain the aerodynamic coefficients and deflections at different speeds and angles of attack. The flexible wing model with optimized curvilinear spars and ribs was used to validate a finite element framework that was used to study the aeroelastic performance of five wing models. As a computationally efficient optimization method, homogenization-based topology optimization was used to generate four different lattice internal structures for the wing in this study. The efficiency of the spline-based optimization used for the spar-rib model and the lattice-based optimization used for the other four wings were compared.
Findings
The aerodynamic loads and displacements obtained experimentally and computationally were in good agreement, proving that additive manufacture can be used to create complex accurate models. The study also shows the efficiency of the homogenization-based topology optimization framework in generating designs with superior stiffness.
Originality/value
To the best of the authors’ knowledge, this is the first time a wing model with curvilinear spars and ribs was additively manufactured as a single piece and tested in a wind tunnel. This research also demonstrates the efficiency of homogenization-based topology optimization in generating enhanced models of different complexity.
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Jianhang Xu, Peng Li and Yiren Yang
The paper aims to develop an efficient data-driven modeling approach for the hydroelastic analysis of a semi-circular pipe conveying fluid with elastic end supports. Besides the…
Abstract
Purpose
The paper aims to develop an efficient data-driven modeling approach for the hydroelastic analysis of a semi-circular pipe conveying fluid with elastic end supports. Besides the structural displacement-dependent unsteady fluid force, the steady one related to structural initial configuration and the variable structural parameters (i.e. the variable support stiffness) are considered in the modeling.
Design/methodology/approach
The steady fluid force is treated as a pipe preload, and the elastically supported pipe-fluid model is dealt with as a prestressed hydroelastic system with variable parameters. To avoid repeated numerical simulations caused by parameter variation, structural and hydrodynamic reduced-order models (ROMs) instead of conventional computational structural dynamics (CSD) and computational fluid dynamics (CFD) solvers are utilized to produce data for the update of the structural, hydrodynamic and hydroelastic state-space equations. Radial basis function neural network (RBFNN), autoregressive with exogenous input (ARX) model as well as proper orthogonal decomposition (POD) algorithm are applied to modeling these two ROMs, and a hybrid framework is proposed to incorporate them.
Findings
The proposed approach is validated by comparing its predictions with theoretical solutions. When the steady fluid force is absent, the predictions agree well with the “inextensible theory”. The pipe always loses its stability via out-of-plane divergence first, regardless of the support stiffness. However, when steady fluid force is considered, the pipe remains stable throughout as flow speed increases, consistent with the “extensible theory”. These results not only verify the accuracy of the present modeling method but also indicate that the steady fluid force, rather than the extensibility of the pipe, is the leading factor for the differences between the in- and extensible theories.
Originality/value
The steady fluid force and the variable structural parameters are considered in the data-driven modeling of a hydroelastic system. Since there are no special restrictions on structural configuration, steady flow pattern and variable structural parameters, the proposed approach has strong portability and great potential application for other hydroelastic problems.
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