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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.

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.

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.

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.

The purpose of this paper is to analyze the active suppression of the nonlinear aeroelastic vibrations of ailerons caused by freeplay by robust H_∞ and linear quadratic Gauss (LQG) methods of control in case of incomplete measurements of the state of the system.

The flexible wing with nonlinear aileron with freeplay is treated as a plant-controller system with H_∞ and LQG controllers used to suppress the aeroelastic vibrations. The simulation approach was used for analyzing the impact of completeness of measurements on the efficiency and robustness of the controllers.

The analysis shows that the H_∞ method can be effectively used for suppression of nonlinear aeroelastic vibrations of aileron, although its efficiency depends essentially on completeness and types of measurements. The LQG method is less effective, but it is also able to prevent aileron vibrations by reducing their amplitudes to acceptable, safe level.

Only numerical analysis was carried out for the problem described; thus, the proposed solution is of theoretical value at this stage of analysis, and its application to the real suppression of aeroelastic vibrations requires further research.

The work presents a potentially useful solution to the problem of interest and results are a theoretical basis for further research.

This work may lead to a hot debate on the advantages and drawbacks of the active suppression of vibrations in the aeroelasticians community.

The work raises the important questions of practical stabilizability of the nonlinear aeroelastic systems, their dependence on completeness and types of measurements and robustness of the controllers.

Aircraft wings, one of the most important parts of an aircraft, have seen changes in its topological and design arrangement of both the internal structures and external shape during the past decades. This study, a numerical, aims to minimize the weight of multilaminate composite aerospace structures using multiobjective optimization.

The methodology started with the determination of the requirements, both imposed by the certifying authority and those inherent to the light, aerobatic, simple, economic and robust (LASER) project. After defining the requirements, the loads that the aircraft would be subjected to during its operation were defined from the flight envelope considering finite element analysis. The design vector consists of material choice for each laminate of the structure (20 in total), ply number and lay-up sequence (respecting the manufacturing rules) and main spar position to obtain a lightweight and cheap structure, respecting the restrictions of stress, margins of safety, displacements and buckling.

The results obtained indicated a predominance of the use of carbon fiber. The predominant orientation found on the main spar flange was 0° with its location at 28% of the local chord, in the secondary and main web were ±45°, the skins also had the main orientation at ±45°.

The key innovations in this paper include the evaluation, development and optimization of a laminated composite structure applied to a LASER aircraft wings considering both structural performance and manufacturing costs in multiobjetive optimization. This paper is one of the most advanced investigations performed to composite LASER aircraft.

The purpose of this study is to develop an active controller of both leading-edge (LE) and trailing-edge (TE) control surfaces for an unmanned air vehicle (UAV) with a composite morphing wing.

Instead of conventional hinged control surfaces, both LE and TE seamless control surfaces were integrated with the wing. Based on the longitudinal state space equation, the root locus plot of the morphing wing aircraft, with a stability augmented system, was constructed. Using the pole placement, the feedback gain matrix for an active control was obtained.

The aerodynamic benefits of a morphing wing section are compared with a wing of a rigid control surface. However, the 3D morphing wing with a large sweptback angle produces a washout negative aeroelastic effect, which causes a significant reduction of the control effectiveness. The results show that the stability augmentation system can significantly improve the longitudinal controllability of an aircraft with a morphing wing.

This study is necessary to analyse the effect of a morphing wing on an UAV and perform a comparison with the rigid model.

The control surfaces assignment plan for trim, pitch and roll control was obtained. An active control algorism for the morphing wing was created to satisfy the required stability and control effectiveness by operating the LE and TE control surfaces according to flight conditions. The aeroelastic effect of control derivatives on the morphing aircraft was considered.

This paper aims to develop an efficient evaluation method to more intuitively and effectively investigate the influence of the wing fuel mass variations because of fuel burn on transonic aeroelasticity.

The proposed efficient aeroelastic evaluation method is developed by extending the standard computational fluid dynamics (CFD)-based proper orthogonal decomposition (POD)/reduced order model (ROM).

The results of this paper show that the proposed aeroelastic efficient evaluation method can accurately and efficiently predict the aeroelastic response and flutter boundary when the wing fuel mass vary because of fuel burn. It also shows that the wing fuel mass variations have a significant effect on transonic aeroelasticity; the flutter speed increases as the wing fuel mass decreases. Without rebuilding an expensive, time-consuming CFD-based POD/ROM for each wing fuel mass variation, the computational cost of the proposed method is reduced obviously. It also shows that the computational efficiency improvement grows linearly with the number of model cases.

The paper presents a potentially powerful tool to more intuitively and effectively investigate the influence of the wing fuel mass variation on transonic aeroelasticity, and the results form a theoretical and methodological basis for further research.

The proposed evaluation method makes it a reality to apply the efficient standard CFD-based POD/ROM to investigate the influence of the wing fuel mass variation because of fuel burn on transonic aeroelasticity. The proposed efficient aeroelastic evaluation method, therefore, is ideally suited to deal with the investigation of the influence of wing fuel mass variations on transonic aeroelasticity and may have the potential to reduce the overall cost of aircraft design.

THIS Note discusses the types of failure to be expected of some main engine components and the problems of avoiding such failure in new designs. Examples are given of the shortcomings of some design practices and in the current presentation of material property data. The need is stressed for more research in these areas to enable true safety factors to be predicted from a knowledge of the state of stress and the appropriate material property.

AT a time when the glamour of space travel is capturing the imagination of the aeronautical world it seems somewhat mundane to be asked to report on the progress of the past years‘ achievements. The first reaction is to question if there has been any progress, or is this just ’another' year.

THERE have been two previous James Forrest Lectures dealing with aeronautics. In 1912, Mr. Mallock addressed this Institution on “Aerial Flight,” and in 1914, Dr. Lanchcster took as his subject “The Flying‐Machine from an Engineering Standpoint.”

IT was in 1947 that we first had the privilege of seeing the SE‐2010 Armagnac under construction and we were then greatly impressed by the courage and determination of all concerned, who were embarking on the ambitious project of building a production series of 70‐ton air liners before testing a prototype. Since our previous article—which dealt in some detail with the whole of the manufacturing methods being employed at Toulouse, Marseilles and Paris, where, because of accommodation difficulties, the various components were being built—the work has gone steadily ahead until a production line exists for the first fifteen airframes.

A collaborative design environment is needed for multidisciplinary design optimization (MDO) process, based on all the modules those for different design/analysis disciplines, and a systematic coupling should be made to carry out aerodynamic shape optimization (ASO), which is an important part of MDO.

Computerized environment for aircraft synthesis and integrated optimization methods (CEASIOM)-ASO is developed based on loosely coupling all the existing modules of CEASIOM by MATLAB scripts. The optimization problem is broken down into small sub-problems, which is called “sequential design approach”, allowing the engineer in the loop.

CEASIOM-ASO shows excellent design abilities on the test case of designing a blended wing body flying in transonic speed, with around 45 per cent drag reduction and all the constraints fulfilled.

Authors built a complete and systematic technique for aerodynamic wing shape optimization based on the existing computational design framework CEASIOM, from geometry parametrization, meshing to optimization.

CEASIOM-ASO provides an optimization technique with loosely coupled modules in CEASIOM design framework, allowing engineer in the loop to follow the “sequential approach” of the design, which is less “myopic” than sticking to gradient-based optimization for the whole process. Meanwhile, it is easily to be parallelized.