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This paper aims to propose an improved and computationally efficient motion simulation of a flexible variable sweep aircraft.

The motion simulation is performed on hardware-in-the-loop simulation setup using 6 degree-of-freedom motion platform. The dynamic model of a flexible variable sweep aircraft, Rockwell B-1 Lancer is presented using equations of motions for combined rigid and flexible motions. The peak filter is introduced as a new method to separate flexible motion from aircraft motion data. Standard adaptive washout filter is modified and redesigned for an accurate flexible aircraft flight simulation. The flight data are generated using FlightGear software. Another motion profile with significant oscillations is also tested. The peak filter and the modified adaptive washout filter both are used to process the data according to the motion envelop of motion platform.

The performance of the modified adaptive washout filter is evaluated using hardware-in-the-loop simulation setup and results are compared with the standard adaptive washout filter. Results exhibit that the proposed method is computationally cost-effective and improves the motion simulation of flexible aircraft with close to realistic motion cues.

The proposed work presents motion simulation of a flexible aircraft by introducing a peak filter to extract flexible motion in contrast to the traditional motion separation methods. Also, a modified adaptive washout filter is designed and implemented in place of the traditional washout filters for improved flexible aircraft flight motion simulation.

The purpose of this paper is to identify the flexible aircraft model accurately from the frequency responses.

The frequency domain output error method is used to estimate the aerodynamic (rigid body and elastic body) derivatives, and mode shape parameters in the process of identification of flexible aircraft model. The accurate identification of lightly damped low frequency rigid-body response modes requires a careful selection of the frequency sweep length and the fast Fourier transform (FFT) window size, as the FFT window length cannot be longer than any individual sweep records. To address this issue, an effort is made to derive the FFT window length for the application of frequency domain estimation approach.

The investigations are initially made to select a suitable FFT window size for the accurate identification of the lightly damped low frequency rigid-body response modes of the flexible aircraft. Subsequently, frequency domain estimation approach is applied to simulated data of flexible aircraft. Besides the stability and control derivatives, the structural modes of the flexible aircraft are also estimated as part of state space model identification, and it is shown that all the model parameter estimates are accurate. Identification of such flexible aircraft aerodynamic (rigid body and elastic body) derivatives and structural mode shape parameters will lead to mathematical models of flexible aircraft that are accurate over a wide frequency range. The identified models are validated using the time response of frequency sweep data.

Aircraft system identification is an integral part of aerospace system design and life cycle process. This becomes a complex process when the aircraft has significant effects of flexibility on the flight dynamics, especially as the frequencies of the elastic modes become lower and approach those of the rigid body modes. Thus, an integrated mathematical model of flexible aircraft is required to develop, and it should be valid for a wide frequency range and relevant for the design of flight control system.

This paper focuses on the application of frequency domain approach to identify the valid model of flexible aircraft by estimating the aerodynamic (rigid body and elastic body) derivatives and structural mode shape parameters of flexible aircraft. The unknown frequencies of structural modes are also able to identify accurately in frequency domain. This gives more value addition to analyze the flight data of flexible aircraft, as it is challenging problem in parameter estimation of flexible aircraft.

This paper presents a broad survey of the structural problems associated with variable geometry for aircraft. Variable sweep allows an aircraft to fly throughout a broad regime of speed and altitude efficiently and without excessive power requirements. Tailored lift drag, improved ride quality, lessening of fatigue damage, and reasonable control sensitivities are advantages. Structural problems fall into two general categories: (1) Because of the number of wing positions, the equivalent of many fixed‐wing aircraft must be investigated, analysed and tested; (2) there are unusual problems which have heretofore not been important considerations in design. Category (1) presents the problem of managing and assimilating large amounts of data. Computer pogrammes and a family of cross‐plots assist greatly. Category (2) presents new fail‐safe criteria, a large lumber of possible flutter‐critical configurations, unavoidable free play in mechanisms which affect flutter speeds, dynamic loads, pivot mechanism bearing life, and requires high reliability in materials. Analyses and wind‐tunnel tests have shown that free play in mechanical joints may or may not cause significant service problems depending upon the mechanical arrangement selected and the actual degree of free play under service conditions.

The purpose of this paper is to build a neural model of an aircraft from flight data and online estimation of the aerodynamic derivatives from established neural model.

A neural model capable of predicting generalized force and moment coefficients of an aircraft using measured motion and control variable is used to extract aerodynamic derivatives. The use of neural partial differentiation (NPD) method to the multi-input-multi-output (MIMO) aircraft system for the online estimation of aerodynamic parameters from flight data is extended.

The estimation of aerodynamic derivatives of rigid and flexible aircrafts is treated separately. In the case of rigid aircraft, longitudinal and lateral-directional derivatives are estimated from flight data. Whereas simulated data are used for a flexible aircraft in the absence of its flight data. The unknown frequencies of structural modes of flexible aircraft are also identified as part of estimation problem in addition to the stability and control derivatives. The estimated results are compared with the parameter estimates obtained from output error method. The validity of estimates has been checked by the model validation method, wherein the estimated model response is matched with the flight data that are not used for estimating the derivatives.

Compared to the Delta and Zero methods of neural networks for parameter estimation, the NPD method has an additional advantage of providing the direct theoretical insight into the statistical information (standard deviation and relative standard deviation) of estimates from noisy data. The NPD method does not require the initial value of estimates, but it requires a priori information about the model structure of aircraft dynamics to extract the flight stability and control parameters. In the case of aircraft with a high degree of flexibility, aircraft dynamics may contain many parameters that are required to be estimated. Thus, NPD seems to be a more appropriate method for the flexible aircraft parameter estimation, as it has potential to estimate most of the parameters without having the issue of convergence.

This paper highlights the application of NPD for MIMO aircraft system; previously it was used only for multi-input and single-output system for extraction of parameters. The neural modeling and application of NPD approach to the MIMO aircraft system facilitate to the design of neural network-based adaptive flight control system. Some interesting results of parameter estimation of flexible aircraft are also presented from established neural model using simulated data as a novelty. This gives more value addition to analyzing the flight data of flexible aircraft as it is a challenging problem in parameter estimation of flexible aircraft.

Currently undergoing NASA flight testing is the Grumman X‐29 forward swept wing (FSW) demonstrator aircraft which was built under a contract sponsored by the Defence Advanced Research Projects Agency (DARPA) and funded through the Air Force. The FSW aircraft first took to the air in December, 1984 and after four flights with the manufacturer was handed over to NASA for a verification programme involving all the benefits of this design. Advanced technology features which will be demonstrated are the forward swept wing for improved aerodynamic efficiency and good control at high angles of attack; tailored composite wing structure that resists the tendency towards structural divergence inherent in this configuration; thin supercritical aerofoil for improved transonic performance at high lift coefficients; variable camber trailing edge to reduce drag at all lift coefficients and avoid supersonic drag associated with aerofoil camber; canard longitudinal control for efficient trimming of variable camber pitching moments and favourable canard‐wing interactions; highly relaxed static stability for very low trim drag at all Mach numbers; and a digital fly‐by‐wire flight control system.

The purpose of this paper is to present an overview of the aircraft design problems which can be efficiently solved using a special solid finite‐element model of variable density.

Optimization algorithms based on fully‐stressed design philosophy, sensitivity coefficients, and employing material density as a design variable provide means to generate optimal topology layouts, subject to a wide range of design constraints. A novel non‐dimensional criterion is used for assessment of load‐carrying efficiency of structures and knowledge accumulation.

Variable density model, together with non‐dimensional criterion of structural efficiency, yields a new versatile approach to a structural weight estimation at early design stages. New weight equations are used. The approach is a powerful tool for addressing complex multidisciplinary design optimization (MDO) problems such as aerodynamic load prediction taking aeroelastic deformations into account and aerodynamic‐structural design optimization of unconventional aircraft configurations.

For accurate estimation of wing weight and deflections, the method should be tuned by regression analysis of existing aircraft to properly account for secondary structural weight.

The developed software tools for aeroelastic behaviour prediction and coupled aerodynamic‐structural design optimization are ready for integration into the complex MDO framework.

The variable density model is shown to have broad predictive opportunities for design problems at early stages of a product development.

EACH September the eyes of the aeronautical World turn towards the S.B.A.C. Air Display and Exhibition with interest unequalled by any other event. It is fitting that the Display is now held each year at the airfield of the Royal Aircraft Establishment, one of the world's most prominent aeronautical research centres. This interest becomes increasingly keen too, as the preview day comes closer, because new prototypes of unorthodox designs often appear a short time before the Show to illustrate the results of years of careful planning, development and research of the particular company. These designs often mould the path of progress for smaller countries without the economic resources to forge the way ahead alone. Most British citizens are very proud of their country's place in aviation today, both in the military and civil fields. This is understood by most foreigners because it is clear that Britain has won a place in aeronautical development second to none.

This paper discusses a large range of layouts that have been studied in relation to both strike and transport V/S.T.O.L. requirements. The performance of aircraft with different powerplant arrangements is considered, with the effects on structure and aerodynamics being based on extensive detail design work.

A Summary by Dr. Alexander Klemin of the Papers Presented Before the Fourteenth Meeting of the Institute held at Columbia University, New York, on January 29–31, 1946. AERODYNAMICS IN spite of increased wing loadings, the use of full span wing flaps has been delayed, because of inability to find a suitable aileron. The Development of a Lateral‐Control System for use with Large‐Span Flaps by I. L. Ashkenas (Northrop Aircraft), outlines the various steps in the aerodynamic development of a retractable aileron system well adapted to the full span flap and successfully employed on the Northrop P‐61. Included is a discussion of the basic data used, the design calculations made, and the effect of structural and mechanical considerations. Changes made as a result of preliminary flight tests are discussed and the final flight‐test results are presented. It is concluded that the use of this retractable aileron system has, in addition to the basic advantage of increased flap span, the following desirable control characteristics: (a) favourable yawing moments, (b) low wing‐torsional loads, (c) small pilot forces, even at high speed.

There are three lateral dynamic attitudes, delineated by rolling, yawing, and sideslipping. It is possible to solve for the pressures on the rolling wing by quasi‐steady analysis. This approach is, however, inapplicable for the yawing or sideslipping wing, and it is with the latter two cases that this paper deals.