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Helicopter fuselage vibration prediction is important to keep a safety and comfortable flight process. The helicopter vibration mechanism model is difficult to meet of demand for accurate vibration prediction. Thus, the purpose of this paper is to develop an intelligent algorithm for accurate helicopter fuselage vibration analysis.

In this research, a novel weighted variational mode decomposition (VMD)- extreme gradient boosting (xgboost) helicopter fuselage vibration prediction model is proposed. The vibration data is decomposed and reconstructed by the signal clustering results. The vibration response is predicted by xgboost algorithm based on the reconstructed data. The information transfer order between the controllable flight data and flight attitude are analyzed.

The mean absolute percentage error (MAPE), root mean square error (RMSE) and mean absolute error (MAE) of the proposed weighted VMD-xgboost model are decreased by 6.8%, 31.5% and 32.8% compared with xgboost model. The established weighted VMD-xgboost model has the highest prediction accuracy with the lowest mean MAPE, RMSE and MAE of 4.54%, 0.0162, and 0.0131, respectively. The attitude of horizontal tail and cycle pitch are the key factors to vibration.

A novel weighted VMD-xgboost intelligent prediction methods is proposed. The prediction effect of xgboost model is highly improved by using the signal-weighted reconstruction technique. In addition, the data set used is collected from actual helicopter flight process.

The purpose of this study is to obtain optimum locations, peak deflection and chord of the twin trailing-edge flaps and optimum torsional stiffness of the helicopter rotor blade to minimize the vibration in the rotor hub with minimum requirement of flap control power.

Kriging metamodel with three-level five variable orthogonal array-based data points is used to decouple the optimization problem and actual aeroelastic analysis.

Some very good design solutions are obtained using this model. The best design point in minimizing vibration gives about 81 per cent reduction in the hub vibration with a penalization of increased flap power requirement, at normal cruise speed of rotor-craft flight.

One of the major challenges in the helicopters is the high vibration level in comparison with fixed wing aircraft. The reduction in vibration level in the helicopter improves passenger and crew comfort and reduces maintenance cost.

This paper presents design optimization of the helicopter rotor blade combining five design variables, such as the locations of twin trailing-edge flaps, peak deflection and flap chord and torsional stiffness of the rotor. Also, this study uses kriging metamodel to decouple the complex aeroelastic analysis and optimization problem.

The purpose of this paper is to discuss published research in rotorcraft which has taken place in India during the last ten years. The helicopter research is divided into the following parts: health monitoring, smart rotor, design optimization, control, helicopter rotor dynamics, active control of structural response (ACSR) and helicopter design and development. Aspects of health monitoring and smart rotor are discussed in detail. Further work needed and areas for international collaboration are pointed out.

The archival journal papers on helicopter engineering published from India are obtained from databases and are studied and discussed. The contribution of the basic research to the state‐of‐the‐art in helicopter engineering science is brought out.

It is found that strong research capabilities have developed in rotor system health and usage monitoring, rotor blade design optimization, ACSR, composite rotor blades and smart rotor development. Furthermore, rotorcraft modeling and analysis aspects are highly developed with considerable manpower available and being generated in these areas.

Two helicopter projects leading to the “advanced light helicopter” and “light combat helicopter” have been completed by Hindustan Aeronautics Ltd These helicopter programs have benefited from the basic research and also provide platforms for further basic research and deeper industry academic collaborations. The development of well‐trained helicopter engineers is also attractive for international helicopter design and manufacturing companies. The basic research done needs to be further developed for practical and commercial applications.

This is the first comprehensive research on rotorcraft research in India, an important emerging market, manufacturing and sourcing destination for the industry.

The purpose of this paper is to examine the success of constrained control on reducing motion blur which occurs as a result of helicopter vibration.

Constrained controllers are designed to reduce the motion blur on images taken by helicopter. Helicopter vibrations under tight and soft constrained controllers are modeled and added to images to show the performance of controllers on reducing blur.

The blur caused by vibration can be reduced via constrained control of helicopter.

The motion of camera is modeled and assumed same as the motion of helicopter. In model of exposing image, image noise is neglected, and blur is considered as the only distorting effect on image.

Tighter constrained controllers can be implemented to take higher quality images by helicopters.

Recently, aerial vehicles are widely used for aerial photography. Images taken by helicopters mostly suffer from motion blur. Reducing motion blur can provide users to take higher quality images by helicopters.

Helicopter control is performed to reduce motion blur on image for the first time. A control-oriented and physic-based model of helicopter is benefited. Helicopter vibration which causes motion blur is modeled as blur kernel to see the effect of helicopter vibration on taken images. Tight and soft constrained controllers are designed and compared to denote their performance in reducing motion blur. It is proved that images taken by helicopter can be prevented from motion blur by controlling helicopter tightly.

The purpose of this study is to use the individual blade pitch control (IBC), reduce actively both the rotor hub vibratory loads and airframe vibration responses for the lift-offset compound helicopter at a high-speed flight condition.

The Sikorsky X2 technology demonstrator (X2TD) is used as the lift-offset compound helicopter. The X2TD lift-offset rotor is modelled and its rotor hub vibratory loads at a flight speed of 250 knots are predicted using a rotorcraft comprehensive analysis code, CAMRAD II, and the airframe structural dynamics is represented with a finite element analysis code, MSC.NASTRAN. When the propulsive trim methodology is applied for rotor trim, the best input condition for IBC using multiple harmonic inputs is searched to reduce the rotor vibration, while the rotor aerodynamic performance (the rotor effective lift-to-drag ratio) is improved or maintained at least. Finally, the reduction in airframe vibration responses is investigated when the best input condition for IBC with multiple harmonics is applied to the lift-offset rotor.

When the IBC with the single harmonic input using the 2/rev actuation frequency, amplitude of 2° and control phase angle of 120° (2P/2°/120°) is considered for X2TD rotor, the rotor vibration is reduced by about 26.37% only and the rotor effective lift-to-drag ratio increases slightly by 0.98%. When X2TD rotor uses the IBC with multiple harmonic inputs (2P/2°/45° + 5P/1°/90°), the rotor hub vibratory loads and airframe vibration responses are reduced by 44.69% and from 0.48 to 79.10%, respectively, while rotor effective lift-to-drag ratio is improved by 0.77%, as compared to the baseline without IBC.

This study is the first study to use the 2/rev actuation for IBC to the four-bladed lift-offset coaxial rotor and to investigate to obtain simultaneously the rotor vibration reduction, rotor performance improvement and airframe vibration reduction, using IBC with multiple harmonic inputs.

A theoretical approach is presented with the object of determining the natural frequencies and hub characteristics of a single‐rotor helicopter in free ground vibration for use in the analysis of self‐excited mechanical oscillations of hinged rotor blades. The method is intended to obviate the experimental determination of these parameters, a technique which necessitates the construction of the helicopter before its vibrational features can be explored.

Describes the Westland active control of structural response (ACSR) technique for reducing and alleviating vibration. Examines a refined coupled rotor/flexible fuselage model based on the ACSR approach. Reports on vibration health monitoring (VHM) which has been developed by Westland and details further work which is continuing in this field.

To investigate the use of centre of gravity location on reducing cyclic pitch control for helicopter UAV's (unmanned air vehicles) and MAV's (micro air vehicles). Low cyclic pitch is a necessity to implement the swashplateless rotor concept using trailing edge flaps or active twist using current generation low authority piezoceramic actuators.

An aeroelastic analysis of the helicopter rotor with elastic blades is used to perform parametric and sensitivity studies of the effects of longitudinal and lateral center of gravity (cg) movements on the main rotor cyclic pitch. An optimization approach is then used to find cg locations which reduce the cyclic pitch at a given forward speed.

It is found that the longitudinal cyclic pitch and lateral cyclic pitch can be driven to zero at a given forward speed by shifting the cg forward and to the port side, respectively. There also exist pairs of numbers for the longitudinal and lateral cg locations which drive both the cyclic pitch components to zero at a given forward speed. Based on these results, a compromise optimal cg location is obtained such that the cyclic pitch is bounded within ±5° for a BO105 helicopter rotor.

The reduction in the cyclic pitch due to helicopter cg location is found to significantly reduce the maximum magnitudes of the control angles in flight, facilitating the swashplateless rotor concept. In addition, the existence of cg locations which drive the cyclic pitches to zero allows for the use of active cg movement as a way to replace the cyclic pitch control for helicopter MAV's.

A variational procedure is developed, in the form of an extension of the Rayleigh‐Ritz method, leading to a rapid estimation of the flapwise vibration modes and frequencies of a helicopter rotor blade. The initial data required are the blade mass and stiffness distribution and the angular velocity of the rotor blade. The normal modes and frequencies are subsequently used to determine blade shapes in flight. The aerodynamic forces only enter at a late stage of the analysis, and the effect of differing flight conditions is readily assessed. The method makes extensive use of matrix formulation and particularly lends itself to electronic computation techniques. A numerical example is given for the special case of constant spanwise blade mass distribution, although the method may readily be extended to cover this restriction. The bending moment distribution is also worked out, and flexible and rigid blades are compared.