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Helicopter Rotor Vibration in the Tip Path Plane

Helicopter Rotor Vibration in the Tip Path Plane Inplane vibration of a balanced helicopter rotor is caused by variations with azimuth of the inplane 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 multirotor 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 tipdriven 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 considered, attention here being confined principally to fully articulated rotors, which are the most difficult to balance because the drag hinges tend to magnify all inaccuracies 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 tipmounted 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. http://www.deepdyve.com/assets/images/DeepDyve-Logo-lg.png Aircraft Engineering and Aerospace Technology Emerald Publishing

Helicopter Rotor Vibration in the Tip Path Plane

Aircraft Engineering and Aerospace Technology , Volume 27 (6): 10 – Jun 1, 1955

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Publisher
Emerald Publishing
Copyright
Copyright © Emerald Group Publishing Limited
ISSN
0002-2667
DOI
10.1108/eb032563
Publisher site
See Article on Publisher Site

Abstract

Inplane vibration of a balanced helicopter rotor is caused by variations with azimuth of the inplane 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 multirotor 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 tipdriven 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 considered, attention here being confined principally to fully articulated rotors, which are the most difficult to balance because the drag hinges tend to magnify all inaccuracies 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 tipmounted 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.

Journal

Aircraft Engineering and Aerospace TechnologyEmerald Publishing

Published: Jun 1, 1955

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