Search results

1 – 10 of 215
Article
Publication date: 31 August 2012

A.S. Roknizadeh, A.S. Nobari, M. Mohagheghi and H. Shahverdi

The purpose of this paper is to analyze the stability of aeroelastic systems using aeroelastic frequency response function (FRF).

Abstract

Purpose

The purpose of this paper is to analyze the stability of aeroelastic systems using aeroelastic frequency response function (FRF).

Design/methodology/approach

The proposed technique determines the instability boundary of an aeroelastic system based on condition number (CN) of aeroelastic FRF matrix or directly from FRFs data.

Findings

Stability margins of typical section and hingeless helicopter rotor blade in the subsonic flow regimes (quasi‐steady and unsteady models) are determined using proposed techniques as two case studies.

Originality/value

The paper introduces a technique which is applicable not only when aerodynamic and structure analytical models are available but also when there are experimental models for structure and/or aerodynamics, such as impulse response functions data or FRFs data. In other words, the main advantage of the proposed method, besides its simplicity and low memory requirement, is its ability to utilize experimental data.

Details

Aircraft Engineering and Aerospace Technology, vol. 84 no. 5
Type: Research Article
ISSN: 0002-2667

Keywords

Article
Publication date: 26 July 2021

Yonghu Wang, Ray C. Chang and Wei Jiang

The purpose of this paper is to present a quick inspection method based on the post-flight data to examine static aeroelastic behavior for transport aircraft subjected to…

Abstract

Purpose

The purpose of this paper is to present a quick inspection method based on the post-flight data to examine static aeroelastic behavior for transport aircraft subjected to instantaneous high g-loads.

Design/methodology/approach

In the present study, the numerical approach of static aeroelasticity and two verified cases will be presented. The non-linear unsteady aerodynamic models are established through flight data mining and the fuzzy-logic modeling of artificial intelligence techniques based on post-flight data. The first and second derivatives of flight dynamic and static aeroelastic behaviors, respectively, are then estimated by using these aerodynamic models.

Findings

The flight dynamic and static aeroelastic behaviors with instantaneous high g-load for the two transports will be analyzed and make a comparison study. The circumstance of turbulence encounter of the new twin-jet is much serious than that of four-jet transport aircraft, but the characteristic of stability and controllability for the new twin-jet is better than those of the four-jet transport aircraft; the new twin-jet transport is also shown to have very small aeroelastic effects. The static aeroelastic behaviors for the two different types can be assessed by using this method.

Practical implications

As the present study uses the flight data stored in a quick access recorder, an intrusive structural inspection of the post-flight can be avoided. A tentative conclusion is to prove that this method can be adapted to examine the static aeroelastic effects for transport aircraft of different weights, different sizes and different service years in tracking static aeroelastic behavior of existing different types of aircraft. In future research, one can consider to have more issues of other types of aircraft with high composite structure weight.

Originality/value

This method can be used to assist airlines to monitor the variations of flight dynamic and static aeroelastic behaviors as a complementary tool for management to improve aviation safety, operation and operational efficiency.

Details

Aircraft Engineering and Aerospace Technology, vol. 94 no. 4
Type: Research Article
ISSN: 1748-8842

Keywords

Article
Publication date: 17 November 2021

Rohollah Dehghani Firouz-Abadi and Mohammad Reza Borhan Panah

The purpose of this paper is to analyze the stability of aeroelastic systems using a novel reduced order aeroelastic model.

Abstract

Purpose

The purpose of this paper is to analyze the stability of aeroelastic systems using a novel reduced order aeroelastic model.

Design/methodology/approach

The proposed aeroelastic model is a reduced-order model constructed based on the aerodynamic model identification using the generalized aerodynamic force response and the unsteady boundary element method in various excitation frequency values. Due to the low computational cost and acceptable accuracy of the boundary element method, this method is selected to determine the unsteady time response of the aerodynamic model. Regarding the structural model, the elastic mode shapes of the shell are used.

Findings

Three case studies are investigated by the proposed model. In the first place, a typical two-dimensional section is introduced as a means of verification by approximating the Theodorsen function. As the second test case, the flutter speed of Advisory Group for Aerospace Research and Development 445.6 wing with 45° sweep angle is determined and compared with the experimental test results in the literature. Finally, a complete aircraft is considered to demonstrate the capability of the proposed model in handling complex configurations.

Originality/value

The paper introduces an algorithm to construct an aeroelastic model applicable to any unsteady aerodynamic model including experimental models and modal structural models in the implicit and reduced order form. In other words, the main advantage of the proposed method, further to its simplicity and low computational effort, which can be used as a means of real-time aeroelastic simulation, is its ability to provide aerodynamic and structural models in implicit and reduced order forms.

Details

Aircraft Engineering and Aerospace Technology, vol. 94 no. 3
Type: Research Article
ISSN: 1748-8842

Keywords

Article
Publication date: 12 October 2012

Mahdi 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.

Article
Publication date: 1 June 2021

Wang Jianhong

The purpose of this paper considers optimal input signal design for flutter model parameters identification, as input signal is the first step during the whole…

Abstract

Purpose

The purpose of this paper considers optimal input signal design for flutter model parameters identification, as input signal is the first step during the whole identification process. According to the constructed flutter stochastic model with observed noises, separable least squares identification and set membership identification are proposed to identify those unknown model parameters for statistical noise and unknown but bounded noise, respectively. The common trace operation with respect to the asymptotic variance matrix is minimized to solve the power spectral for the optimal input signal in the framework of statistical noise. Moreover, for the unknown bout bounded noise, the radius of information, corresponding to the established parameter uncertainty interval, is minimized to give the optimal input signal.

Design/methodology/approach

First, model identification for aircraft flutter is reviewed as one problem of parameter identification and this aircraft flutter model corresponds to one stochastic model, whose input signal and output are corrupted by external noises. Second, for aircraft flutter statistical model with statistical noise, separable least squares identification is proposed to identify the unknown model parameters, then the optimal input signal is designed to satisfy one given performance function. Third, for aircraft flutter model with unknown but bounded noise, set membership identification is proposed to solve the parameter set for each unknown model parameter. Then, the optimal input signal is designed by applying the idea of the radius of information with unknown but bounded noise.

Findings

This aircraft flutter model corresponds to one stochastic model, whose input signal and output are corrupted by external noises. Then identification strategy and optimal input signal design are studied for aircraft flutter model parameter identification with statistical noise and unknown but bounded noise, respectively.

Originality/value

To the best knowledge of the authors, this problem of the model parameter identification for aircraft flutter was proposed by their previous work, and they proposed many identification strategies to identify these model parameters. This paper proposes two novel identification strategies and opens a new subject about optimal input signal design for statistical noise and unknown noise, respectively.

Details

Aircraft Engineering and Aerospace Technology, vol. 93 no. 10
Type: Research Article
ISSN: 1748-8842

Keywords

Article
Publication date: 28 May 2019

Mohammadreza Amoozgar and Hossein Shahverdi

This paper aims to develop a new approach for aeroelastic analysis of hingeless rotor blades.

Abstract

Purpose

This paper aims to develop a new approach for aeroelastic analysis of hingeless rotor blades.

Design/methodology/approach

The aeroelastic approach developed here is based on the geometrically exact fully intrinsic beam equations and three-dimensional unsteady aerodynamics.

Findings

The developed approach is accurate, fast and very useful in rotorcraft aeroelastic analysis.

Originality/value

This beam formulation has been never combined with three-dimensional aerodynamic model to be used for aeroelastic analysis of blades. In addition, it is possible to handle the composite blades, as well as blades with initial curvatures and twist with this proposed formulation.

Details

Aircraft Engineering and Aerospace Technology, vol. 91 no. 8
Type: Research Article
ISSN: 1748-8842

Keywords

Article
Publication date: 23 January 2009

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…

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.

Details

Aircraft Engineering and Aerospace Technology, vol. 81 no. 2
Type: Research Article
ISSN: 0002-2667

Keywords

Article
Publication date: 30 January 2007

Altan Kayran

To provide a general review of the flight flutter test techniques utilized in aeroelastic stability flight testing of aircraft, and to highlight the key items involved in…

1711

Abstract

Purpose

To provide a general review of the flight flutter test techniques utilized in aeroelastic stability flight testing of aircraft, and to highlight the key items involved in flight flutter testing of aircraft, by emphasizing all the main information processed during the flutter stability verification based on flight test data.

Design/methodology/approach

Flight flutter test requirements are first reviewed by referencing the relevant civil and military specifications. Excitation systems utilized in flight flutter testing are overviewed by stating the relative advantages and disadvantages of each technique. Flight test procedures followed in a typical flutter flight testing is described for different air speed regimes. Modal estimation methods, both in frequency and time domain, used in flutter prediction are surveyed. Most common flight flutter prediction methods are reviewed. Finally, key considerations for successful flight flutter testing are noted by referencing the related literature.

Findings

Online, real time monitoring of flutter stability during flight testing is very crucial, if the flutter character is not known a priori. Techniques such as modal filtering can be used to uncouple response measurements to produce simplified single degree of freedom responses, which could then be analyzed with less sophisticated algorithms that are more able to run in real time. Frequency domain subspace identification methods combined with time‐frequency multiscale wavelet techniques are considered as the most promising modal estimation algorithms to be used in flight flutter testing.

Practical implications

This study gives concise but relevant information on the flight flutter stability verification of aircraft to the practicing engineer. The three important steps used in flight flutter testing; structural excitation, structural response measurement and stability prediction are introduced by presenting different techniques for each of the three important steps. Emphasis has been given to the practical advantages and disadvantages of each technique.

Originality/value

This paper offers a brief practical guide to all key items involved in flight flutter stability verification of aircraft.

Details

Aircraft Engineering and Aerospace Technology, vol. 79 no. 2
Type: Research Article
ISSN: 0002-2667

Keywords

Article
Publication date: 12 October 2012

Emanuele Piccione, Giovanni Bernardini and Massimo Gennaretti

The purpose of this paper is to present the development and application of a numerical formulation for the structural dynamics and aeroelastic analysis of new generation…

3772

Abstract

Purpose

The purpose of this paper is to present the development and application of a numerical formulation for the structural dynamics and aeroelastic analysis of new generation helicopter and tiltrotor rotor blades. These are characterized by a curvilinear elastic axis, typically with the presence of tip sweep and anhedral angles.

Design/methodology/approach

The structural dynamics model implemented is based on nonlinear, flap‐lag‐torsion, rotating beam equations that are valid for slender, homogeneous, isotropic, non‐uniform, twisted blades undergoing moderate displacements. A second‐order approximation scheme for strain‐displacement is adopted. Aerodynamic contributions for aeroelastic applications are derived from sectional theories, with inclusion of wake inflow models to take into account three‐dimensional effects. The numerical integration is obtained through implementation within the COMSOL Multiphysics Finite‐Element‐Method (FEM) software code, considering the elastic axis of arbitrary curvilinear shape.

Findings

The computational tool developed is validated by comparisons with results available in the literature. These demonstrate the capability of the tool to accurately predict structural dynamics and aeroelastic behavior of curved‐axis rotor blades. In particular, the influence of sweep and anhedral angles at the blade tip is successfully captured.

Research limitations/implications

The numerical tool developed is limited to the analysis of isotropic blades, with a simple sectional aerodynamic modeling for aeroelastic applications. However, the flexibility of the process through which the proposed tool has been developed is such that a moderate effort is required for its extension to composite blades and more accurate aerodynamic loads predictions.

Practical implications

The proposed computational solver is a reliable tool for preliminary design and optimal design processes of helicopter and tiltrotor rotor blades.

Originality/value

Computational tools for rotors with advanced‐geometry blades are not commonly available. Therefore, the presentation of a successful way to implement structural dynamics/aeroelastic mathematical formulations for rotor blades with curvilinear elastic axis in highly flexible, multiphysics, FEM‐based, commercial software may be of interest for designers and researchers.

Details

Aircraft Engineering and Aerospace Technology, vol. 84 no. 6
Type: Research Article
ISSN: 0002-2667

Keywords

Article
Publication date: 1 April 2004

R. Srivastava, M.A. Bakhle, K T.G. and D. Hoyniak

Part II of the two‐part paper describes an aeroelastic analysis program and its application for stability computations of turbomachinery blade rows. Unsteady Euler or…

Abstract

Part II of the two‐part paper describes an aeroelastic analysis program and its application for stability computations of turbomachinery blade rows. Unsteady Euler or Navier‐Stokes equations are solved on dynamically deforming, body fitted, and grid to obtain the aeroelastic characteristics. Blade structural response is modeled using a modal representation of the blade and the work‐per‐cycle method is used to evaluate the stability characteristics. Non‐zero inter‐blade phase angle is modeled using phase‐lagged boundary conditions. Results are presented for a flat plate helical fan, a turbine cascade and a high‐speed fan, to highlight the aeroelastic analysis method, and its capability and accuracy. Obtained results showed good correlation with existing experimental, analytical and numerical results. Numerical analysis also showed that given the computational resources available currently, engineering solutions with good accuracy are possible using higher fidelity analyses.

Details

International Journal of Numerical Methods for Heat & Fluid Flow, vol. 14 no. 3
Type: Research Article
ISSN: 0961-5539

Keywords

1 – 10 of 215