GB2274634A - Controlling helicopter anti-torque rotor. - Google Patents

Abstract

This invention provides a method and apparatus for reducing the flap movements of the rotor blades of a helicopter anti-torque rotor by introducing a one per rev cyclic pitch change of the rotor blades in addition to the conventional collective pitch change. An actuator 37 acts transversely on one end of a control shaft 25 pivoted about a ball joint 34 and carrying a spider 27 connected by links 28 to pitch control lugs 29 on the blades 19. Sensors 40, 41 for forward speed and sideslip angle of the helicopter send signals to a computer 42 which provide control signals to the actuator 37 in accordance with flight conditions. A conventional collecting pitch actuator 32 moves the shaft 25 axially. <IMAGE>

Description

Title: HELICOPTERS This invention relates to helicopters and is particularly concerned with helicopters having a single main sustaining rotor and an anti-torque tail rotor.

One of the limitations to increasing the performance such as the forward speed and maneouvrability of such helicopters is the anti-torque rotor.

Conventionally, an anti-torque rotor is controlled by means for collectively changing the pitch of all of its rotor blades to provide a lateral thrust force sufficient to provide anti-torque and yaw control of the helicopter. The anti-torque rotor, similarly to a main sustaining rotor, operates edgewise through the air, albeit in a different plane, and is therefore subjected to similar extreme asymmetry of flow conditions over the advancing and retreating sides during forward flight of the helicopter. This results in flapping of each blade in a plane perpendicular to the plane of rotation during each revolution, and this blade flapping movement causes high aerodynamic forces resulting in high loads and consequent stresses in the rotor blades and rotor hub of the anti-torque rotor in the plane of rotation.Such loads are known in respect of both main sustaining rotors and anti-torque rotors as lead/lag loads.

It follows then that any increase in blade flapping such as would be caused by increased forward speed or operation outside a normal side slip limitation as is likely to result from increased manoeuvrability, will result in an increase in the lead/lag loads in the anti-torque rotor.

Conventionally, anti-torque tail rotor design has ensured adequate strength to cater for the lead/lag loads encountered within an allowable flight envelope. To simply strengthen the components to cater for increased lead/lag loads encountered in an extended flight envelope is not an appropriate remedy because it results in an undesirable increase in weight and destroys a sensible balance between adequate strength combined with an acceptable weight and appropriate dynamic characteristics.

It has become common practice in respect of main sustaining rotors to incorporate a lead/lag hinge, either mechanical or in the form of a flexure member, in the joint between the hub and blades to alleviate lead/lag loads. However, it is known from such installations in a main rotor that damping is required across lead/lag hinges in order to prevent a type of instability resulting in a phenomenon known in the art as ground resonance, and the requirement for such dampers on an anti-torque tail rotor to prevent a similar type of instability would be inappropriate due to an inevitable increase in weight, cost, maintenance activities and aerodynamic drag.

An objective of this invention is therefore to provide a helicopter having an anti-torque tail rotor in which blade flapping movement is minimised so as to reduce lead/lag loads.

Accordingly, in one aspect this invention provides a helicopter having a fuselage supporting a main sustaining rotor for rotation about a generally vertical axis, an anti-torque tail rotor located at the end of a tail boom extending rearwardly from the fuselage for rotation about a generally horizontal axis, a power source for rotating said main sustaining and anti-torque tail rotors, and control means for changing the pitch of the blades of the main sustaining rotor both collectively and cyclically and for changing the pitch of the blades of the tail rotor collectively, characterised in that said control means includes means for introducing a one per rev cyclic pitch change of the rotor blades of the anti-torque tail rotor.

In another aspect, the invention provides a method for reducing the flap movements of the rotor blades of a helicopter anti-torque tail rotor having means for changing the pitch of the blades collectively comprises the step of introducing a one per rev cyclic pitch change of each rotor blade.

The invention will now be described by way of example only and with reference to the accompanying drawings in which, Figure 1 is a side view of a helicopter, and Figure 2 is a fragmentary sectioned view taken on lines A-A of Figure 1.

Referring now to Figure 1, a helicopter 10 includes a fuselage 11 supporting a main sustaining rotor 12 comprising a rotor head 13 mounted for rotation about a generally vertical axis 14. Rotor head 13 supports a plurality of radially extending rotor blades 15 and control means (not shown) are provided for changing the pitch of the rotor blades 15 of the main sustaining rotor 12 both collectively and cyclically in a conventional manner.

An anti-torque tail rotor generally indicated at 16 is located at the end of a tail boom 17 extending rearwardly from the fuselage lib The tail rotor 16 includes a rotor head 18 supporting a plurality of radially extending rotor blades 19 for rotation about a generally horizontal axis 20.

A power source and transmission system (not shown) is provided for rotating the main sustaining rotor 12 and the anti-torque tail rotor 16 in a conventional manner.

As shown in Figure 2, the rotor head 18 of the anti-torque tail rotor 16 is secured at one end of a tubular drive shaft 21 protruding from a gearbox 22 driven by the afore-mentioned transmission system. The rotor blades 19 are attached to the rotor head 18 through a flap hinge 23 and a pitch change hinge 24.

One end of an axially extending actuator shaft 25 protrudes from the drive shaft 21 and carries a pitch control spider 26 having radially extending arms 27 equal in number to the number of rotor blades 19.

The end of each of the arms 27 is connected by a pivotally mounted pitch link 28 to a pitch control lug 29 attached to the root end of each rotor blade 19. The other end of the actuator shaft 25 protrudes from the gearbox 22 and is located in a bearing housing 30 permitting relative rotation. Bearing housing 30 is connected through a pivoted link 31 to a collective input electrohydraulic jack 32 which is operated by a pilot operated control rod 33.

The actuator shaft 25 is mounted for axial sliding movement in the drive shaft 21, such movement being caused during operation by actuation of the collective pitch jack 32 in the directions indicated by arrows 32a by the pilot control rod 33 to change the pitch of all of the rotor blades 19 uniformly. This collective pitch control of the rotor blades 19 of an anti-torque tail rotor 16 is cdnventional in the art and forms no part of the present invention.

In the helicopter of this invention the actuator shaft 25 carries a spherical joint 34 intermediate its ends which is located in a mating part spherical internal surface of a tubular carrier 35 whose external surface carries axial splines (not shown) located in mating axial splines 36 on the internal surface of the tubular drive shaft 21. An electro-hydraulic cyclic pitch input actuator 37 is connected through a pivoted link 38 to the bearing housing 30 at the inner end of the actuator shaft 25, and is controlled by an electric signal 39 generated by a control system as hereinafter described.

Actuation of the jack 37 in the direction indicated by arrows 37a causes tilting of the actuator shaft 25 about the spherical joint 34 which in operation introduces a one per rev cyclic pitch change to the rotor blades 19 of the anti-torque tail rotor 16. Preferably, the jack 37 is located so as to apply the cyclic pitch input at about an upper generally vertical position identified as a 900 azimuth position in Figure 1 so as to have maximum effect at about 1800 and 0 azimuth positions, i.e. the positions of maximum flap deflection due to the blades experiencing a maximum difference in relative rotational speeds at about the 900 and 2700 azimuth positions.

The control signal 39 can be generated either manually or automatically. In its simplest form the signal can be generated manually directly by the pilot, however, in practice this may not be desirable because of the increase in pilot workload. An automatic system for generating control signal 39 is illustrated in Figure 2 and comprises an open loop system including sensors 40 and 41 for sensing forward speed and side slip angle respectively, and for providing signals to a computer 42 for generating control signals 39. A more sophisticated closed loop system may incorporate sensors for sensing lead/lag loads at the rotor head 18 as a basis for calculating the one per rev inputs to reduce the lead/lag loads.

In operation of the helicopter of this invention, the main sustaining rotor 12 is controlled by conventional collective and cyclic pitch control systems operated by collective and cyclic pitch change levers located in a cockpit. Similarly, conventional pilot operated pedals are provided in the cockpit to adjust collectively through control rod 33 and collective input jack 32 the pitch of all of the rotor blades 19 of the anti-torque tail rotor 16 to provide a lateral thrust to counter main rotor torque and to provide yaw control.

When zero collective pitch is effective on the tail rotor blades 19 so that no lateral thrust is produced by the tail rotor 16, no movement of the blades 19 occurs about the flap hinges 23 and the blades 19 rotate in a plane generally perpendicular to the axis of rotation 20 as indicated at X-X in Figure 2.

The application of collective pitch to the rotor blades 19 to provide a lateral thrust from tail rotor 16 when the helicopter 10 is in the hover mode, i.e. with no forward velocity, causes uniform flap movement of all of the rotor blades 19 throughout each revolution to a position indicated at Y-Y in Figure 2. This is known in the art as blade "coning" and the angle between planes X-X and Y-Y as the "coning angle". The actual coning angle in any installation depends on the rotational speed and the amount of collective pitch applied to provide a required thrust.

Assuming now that anti-torque tail rotor 16 rotates counter clockwise as indicated by arrow B in Figure 1, and that the helicopter is flying forward resulting in a relative generally longitudinal airflow at the tail rotor 16 in the direction of arrow C. It will be seen that the relative airspeed of the blade 19 in position D equals its rotational speed plus the speed of airflow indicated at C, and the relative airspeed of the blade 19 in position E equals its rotational speed minus the speed of airflow indicated at C.This relative speeding up and slowing down of each rotor blade 19 as it rotates causes the blade 19 when in position D to commence to flap about its flap hinge 23 further outwardly from position Y through angle Dl until a maximum flap deflection is reached at a rotational position about 900 later, i.e., a 1800 azimuth position, as indicated at D2 in Figure 2. The blade 19 at position E commences to flap inwardly from position Y through angle E until a minimum flap deflection is reached at a rotational position about 900 later, i.e., a 0 azimuth position, as indicated at E2 in Figure 2.

As previously explained, in prior art helicopters such flap movements during rotation of the rotor blades 19 result in high in-plane or lead/lag loads especially at the root of the blades 19 and at the blade attachment arms 18 of the rotor head.

In the helicopter of this invention, the sensing means 40 and 41 provide signals to computer 42 representative of forward speed and side slip angle, and computer 42 calculates control signal 39 to operate the cyclic input jack 37 to tilt the actuator shaft 25 about spherical joint 34 to introduce a one per rev cyclic pitch to adjust the pitch of each rotor blade 19 cyclically as it rotates through each revolution.

As previously mentioned, the cyclic input is preferably applied at about the vertical position of the rotor blades 19, i.e. at 900 or 2700 azimuth positions (Figure 1), so as to have maximum effect at about 900 later when a blade is in a 1800 azimuth position and a 0 azimuth position that correspond to the flap positions indicated at D2 and E2 (Figure 2) in order to adjust the pitch of the blades at those azimuth positions to reduce the blade flap movement.

The cyclic pitch input to the anti-torque tail rotor 16 of this invention enables trimming of the anti-torque tail rotor 16 by reducing or eliminating the amount of flap movement of the rotor blades as indicated at D1 and E1 in Figure 2. This reduces the aerodynamic forces active on the rotor blades and the resultant lead/lag loads.

Thus, the helicopter of this invention, introduces cyclic pitch control of the rotor blades of an anti-torque tail rotor to significantly reduce normal lead/lag loads caused by large flap movements. This means that the weight of the anti-torque rotor can be kept at a minimum since strengthening is not necessary and nor are lead/lag hinges and dampers with their attendant increase in cost and aerodynamic drag.

The cyclic pitch actuator 37 is not an exposed component so that it is not subject to environmental damage and it has no detrimental effect on the aerodynamic drag characteristics of the tail rotor 16.

Incorporation of the present invention therefore will minimise the limitation on helicopter performance associated with the conventional anti-torque tail rotor of prior helicopters, enabling a useful improvement in performance in respect of forward speed and maneouvrability to be achieved.

Whilst one embodiment has been described and illustrated it will be understood that many modifications may be made without departing from the scope of the invention. For example spherical joint 34 could be replaced by other suitable means such as a gimbal joint, and the actuator shaft 25 and spider 26 could be replaced by a swash plate mechanism for applying both collective and cyclic pitch inputs. More than one cyclic pitch input actuator 37 can be incorporated in order to increase system sophistication. The mechanical flap hinges 23 and pitch change hinges 24 can be replaced either by elastomeric devices or by a flexible torsionally resilient beam member.

Claims (2)

1. A helicopter having a fuselage supporting a main sustaining rotor for rotation about a generally vertical axis, an anti-torque rotor located at the end of a tail boom extending rearwardly from the fuselage for rotation about a generally horizontal axis, a power source for rotating said main sustaining and anti-torque rotors, and control means for changing the pitch of the blades of the main sustaining rotor both collectively and cyclically and for changing the pitch of the blades of the anti-torque rotor collectively, characterised in that said control means includes means for introducing a one per rev cyclic pitch change of the rotor blades of the anti-torque rotor.

2. A method for reducing the flap movements of the rotor blades of a helicopter antitorque rotor having control means for changing the pitch of the rotor blades collectively comprises the step of introducing a one per rev cyclic pitch change of each rotor blade.