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Article
Publication date: 18 December 2019

Firat Sal

The purpose of this paper presents the effects of actively morphing root chord and taper on the energy of the flight control system (i.e. FCS).

Abstract

Purpose

The purpose of this paper presents the effects of actively morphing root chord and taper on the energy of the flight control system (i.e. FCS).

Design/methodology/approach

Via regarding previously mentioned purposes, sophisticated and realistic helicopter models are benefitted to examine the energy of the FCS.

Findings

Helicopters having actively morphing blade root chord length and blade taper consume less control energy than the ones having one of or any of passively morphing blade root chord length and blade taper.

Practical implications

Actively morphing blade root chord length and blade taper can be used for cheaper helicopter operations.

Originality/value

The main originality of this paper is applying active morphing strategy on helicopter blade root chord and blade taper. In this paper, it is also found that using active morphing strategy on helicopter blade root chord and blade taper reasons less energy consumption than using either passively morphing blade root chord length plus blade taper or not any. This causes also less fuel consumption and green environment.

Details

Aircraft Engineering and Aerospace Technology, vol. 92 no. 2
Type: Research Article
ISSN: 1748-8842

Keywords

Article
Publication date: 8 November 2019

Firat Sal

The purpose of this study is to examine the effect of passive and active morphing of blade root chord length and blade taper on the control effort of the flight control system…

Abstract

Purpose

The purpose of this study is to examine the effect of passive and active morphing of blade root chord length and blade taper on the control effort of the flight control system (FCS) of a helicopter.

Design/methodology/approach

Physics-based helicopter models, which are functions of passive and active morphing, are created and applied in helicopter FCS design to determine the control effort.

Findings

Helicopters, having both passively and actively morphing blade root chord length and blade taper, experience less control effort than the ones having either only passively morphing blade root chord length or only blade taper or only actively morphing blade root chord length and blade taper.

Practical implications

Both passively and actively morphing blade root chord length and blade taper can be implemented for more economical autonomous helicopter flights.

Originality/value

Main novelty of our article is simultaneous application of passive and active morphing ideas on helicopter root chord length and blade taper. It is also proved in this study that using both passive and active morphing ideas on helicopter blade root chord and blade taper causes much less energy consumption than using either only passive morphing idea on helicopter blade root chord and blade taper or only active morphing idea on helicopter blade root chord and blade taper. This also reduces fuel consumption and also makes environment cleaner.

Details

Aircraft Engineering and Aerospace Technology, vol. 92 no. 2
Type: Research Article
ISSN: 1748-8842

Keywords

Article
Publication date: 13 June 2019

Farid Shahmiri, Maryam Sargolzehi and Mohammad Ali Shahi Ashtiani

The effects of rotor blade design variables and their mutual interactions on aerodynamic efficiency of helicopters are investigated. The aerodynamic efficiency is defined based on…

Abstract

Purpose

The effects of rotor blade design variables and their mutual interactions on aerodynamic efficiency of helicopters are investigated. The aerodynamic efficiency is defined based on figure of merit (FM) and lift-to-drag responses developed for hover and forward flight, respectively.

Design/methodology/approach

The approach is to couple a general flight dynamic simulation code, previously validated in the time domain, with design of experiment (DOE) required for the response surface development. DOE includes I-optimality criteria to preselect the data and improve data acquisition process. Desirability approach is also implemented for a better understanding of the optimum rotor blade planform in both hover and forward flight.

Findings

The resulting system provides a systematic manner to examine the rotor blade design variables and their interactions, thus reducing the time and cost of designing rotor blades. The obtained results show that the blade taper ratio of 0.3, the point of taper initiation of about 0.64 R within a SC1095R8 airfoil satisfy the maximum FM of 0.73 and the maximum lift-to-drag ratio of about 5.5 in hover and forward flight.

Practical implications

The work shows the practical possibility to implement the proposed optimization process that can be used for the advanced rotor blade design.

Originality/value

The work presents the rapid and reliable optimization process efficiently used for designing advanced rotor blades in hover and forward flight.

Details

Aircraft Engineering and Aerospace Technology, vol. 91 no. 9
Type: Research Article
ISSN: 1748-8842

Keywords

Article
Publication date: 1 August 1954

P.R. Payne

BECAUSE of the complexity of rotor theory which has been developed for tapered and twisted blades (for example, Sissingh's well known papers) there is a great tendency to use and…

Abstract

BECAUSE of the complexity of rotor theory which has been developed for tapered and twisted blades (for example, Sissingh's well known papers) there is a great tendency to use and draw conclusions from the simple theory developed for untwisted untapered blades. In the United States this practice appears to be almost universal, and much use is made of the highly questionable ‘equivalent chord’ when dealing with tapered blades. It cannot be too much emphasized that there is no such thing as an ‘equivalent chord’, and that its use not only masks the true effect of taper, but leads to solutions which in some cases are in considerable error.

Details

Aircraft Engineering and Aerospace Technology, vol. 26 no. 8
Type: Research Article
ISSN: 0002-2667

Article
Publication date: 1 June 1955

P.R. Payne

In‐plane vibration of a balanced helicopter rotor is caused by variations with azimuth of the in‐plane forces acting on individual blades. These forces may be summarized under…

80

Abstract

In‐plane vibration of a balanced helicopter rotor is caused by variations with azimuth of the in‐plane forces acting on individual blades. These forces may be summarized under three headings: ‘Induced forces’ caused by the inclination of elemental lift vectors relative to the axis of rotation. ‘Profile drag forces’: variations are caused by changes with azimuth angle of the angle and airspeed of the individual blade elements. ‘Coriolis forces’, which are caused by blade flapping, which brings about a variation of blade moment of inertia about the axis of rotation. Equations are developed in this paper for the resultant hub force due to each of these forces, on the assumptions of small flapping hinge offset. It is assumed that blades are linearly twisted and tapered, an assumption which in practice can be applied to any normal rotor. It is shown that by suitably inclining the mechanical axis it is possible to balance out the worst induced and profile drag vibrations by the coriolis one, which can be made to have opposite sign. If the mechanical axis is fixed in the fuselage, this suppression is fully effective for one flight condition only. In multi‐rotor helicopters, vibration suppression can be extended over a much wider range by varying the fuselage attitude. The logical result of this analysis is, for single rotor helicopters, a floating mechanical axis which can be adjusted or trimmed by the pilot. This would be quite simple to do on a tip‐driven rotor, and has already been achieved with a mechanical drive on the Doman helicopter. The more important causes of vibration from an unbalanced rotor are next con‐sidered, attention here being confined principally to fully articulated rotors, which are the most difficult to balance because the drag hinges tend to magnify all in‐accuracies in finish and balance. From a brief discussion of the vertical vibration of an imperfect rotor it is shown that some contemporary methods of ‘tracking’ are fundamentally wrong. Finally the vibration due to tip‐mounted power units is described. In discussing the effect of a vibratory force on a helicopter a simple response chart is developed, and it is thought that its use could well be accepted as a simple standard for general assessment purposes. In the development of equations for vibration the following points of general technical interest are put forward: An equation for induced torque is developed which includes a number of hitherto neglected parameters. A new form of equation for mean lift coefficient of a blade is suggested. The simple Hafner criterion for flight envelopes is shown to give rise to considerable error, and the use of Eq. (28) is suggested in its place. The variation of profile torque with forward speed is given, and the increase due to ? varying round the disk is expressed as an explicit equation, thus allowing considerable improvement in the present methods of allowing for this effect.

Details

Aircraft Engineering and Aerospace Technology, vol. 27 no. 6
Type: Research Article
ISSN: 0002-2667

Article
Publication date: 1 February 1954

P.R. Payne

In Part 1 it is shown that the equation for blade equilibrium about the flapping pin is

Abstract

In Part 1 it is shown that the equation for blade equilibrium about the flapping pin is

Details

Aircraft Engineering and Aerospace Technology, vol. 26 no. 2
Type: Research Article
ISSN: 0002-2667

Article
Publication date: 1 November 1954

P.R. Payne

The theory of rotor dynamics given in Ref. 1 is extended to include the effects of coupling between feathering and flapping (δ3 angle) and flapping hinge offset. Both introduce…

Abstract

The theory of rotor dynamics given in Ref. 1 is extended to include the effects of coupling between feathering and flapping (δ3 angle) and flapping hinge offset. Both introduce considerable modification to the classic equations, and instead of simple explicit equations for flapping amplitudes, coning angle, collective pitch and inflow angles, five simultaneous equations have now to be solved. Data sheets have been constructed which enable this to be done quickly and accurately for any design of linearly tapered and twisted blade. It is suggested that the intelligent use of such data sheets is of great assistance in a design office, not only because of the very considerable time savings achieved, but also because they eliminate the most fruitful sources of error in numerical calculation. It is shown that a high offset rotor enables much higher speeds to be achieved with a conventional helicopter—an effect which has already been fairly well publicized. A penalty is paid for this in the form of hub pitching moments which have to be balanced out externally; either by the use of two rotors, offset C.G., aerodynamic surfaces, or inclination of the mechanical axis. These effects will be considered in detail in a further article. Finally, equations are developed for a convenient method of calculating blade elemental angle of attack which is claimed to be superior to classic methods for design office purposes.

Details

Aircraft Engineering and Aerospace Technology, vol. 26 no. 11
Type: Research Article
ISSN: 0002-2667

Article
Publication date: 1 May 1957

P.R. Payne

A NUMBER of approaches to the calculation of rotor downwash have already been discussed. Broadly spsaking, the methods of Castles and DeLeeuw and Squire and Mangler are the same…

Abstract

A NUMBER of approaches to the calculation of rotor downwash have already been discussed. Broadly spsaking, the methods of Castles and DeLeeuw and Squire and Mangler are the same. In both methods the downwash at the rotor disk is assumed to be perpetrated in a helical downwash sheet which, as the slipstream, extends below the rotor to infinity. The downwash in the disk due to the bound vortices, and the additional downwash in the disk which is induced by the helical sheets in the slipstream (Castles and DeLeeuw substitute downwash rings for helices, in the interest of mathematical simplicity) is calculated, on the assumption of an infinite number of lightly loaded blades. The final results of Castles and DeLeeuw on the one hand, and Squire and Mangier on the other, are in very wide disagresment. This disagreement is principally due to the fact that, whereas the first investigation assumes constant circulation along the blade (ideal twist and taper), Mangier and Squire assume a ‘practical’ variation of the form likely to be encountered on an untwisted untapered blade. We conclude that the radial distribution of lift on a helicopter blade will have a profound effect on the downwash pattern: which in turn will affect the calculated lift.

Details

Aircraft Engineering and Aerospace Technology, vol. 29 no. 5
Type: Research Article
ISSN: 0002-2667

Article
Publication date: 1 November 1958

J. Lockwood Taylor

HELICOPTER rotor blades, as is well known, are subjected to alternating bending stresses at various multiples of the rotor r.p.m. when the aircraft is flying forward. One of the…

Abstract

HELICOPTER rotor blades, as is well known, are subjected to alternating bending stresses at various multiples of the rotor r.p.m. when the aircraft is flying forward. One of the factors determining the magnitude of the fluctuating stresses at the various frequencies is the approach to resonance between the harmonic components of the periodically varying aerodynamic loads and the natural blade frequencies. Most rotor blades are flexible enough to allow the frequencies of at any rate the lower modes, which are probably the most important, to be estimated by applying a correction for blade bending stiffness to the natural frequency of the blade if it were perfectly flexible, and subject only to inertia and centrifugal forces. It is on this latter aspect of the problem that the present article is concentrated.

Details

Aircraft Engineering and Aerospace Technology, vol. 30 no. 11
Type: Research Article
ISSN: 0002-2667

Article
Publication date: 1 July 1969

W.N. Twelvetrees

THE helicopter presents a classic case of a device for which there is an insatiable world demand that cannot be met because the current state of engineering development prohibits…

Abstract

THE helicopter presents a classic case of a device for which there is an insatiable world demand that cannot be met because the current state of engineering development prohibits an economic product. The overriding factor in any airborne vehicle is the necessity of achieving a viable payload and performance at an acceptable cost with dependable standards of safety and reliability. Sensible figures are difficult to quote but it would not be far wrong to indicate that existing helicopters carry an all in cost penalty of 5 to 1 compared with fixed wing aircraft of the same all up weight. It is not surprising therefore that, apart from prodigal military usage and small specialist operations as with offshore drilling rigs, progress over the last decade has been slow. Indeed it would be more realistic to say that in terms of basic design there has been what is more like stagnation, as at least three major projects in the U.K. alone have been dropped after many millions of pounds had been spent on them.

Details

Aircraft Engineering and Aerospace Technology, vol. 41 no. 7
Type: Research Article
ISSN: 0002-2667

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