GB2387157A - A rotor pitch control system for a rotating wing aircraft - Google Patents

Abstract

A rotor pitch control system comprises an axially and angularly adjustable transmission joint 495 which may include a constant velocity drive ball, and is linked to a follower 490. The follower is attached to a blade 339, to provide pitch controlling output 7D, and the blade may further comprise servo tabs (337 fig. 1). Control computers 455 (and 457 fig. 2) may control the pitch control system, whereby gearing 480, (498 fig. 5) may be utilised to adjust the angle of the transmission joint 495 relative to the follower 490 to provide both cyclic and collective control output, and permit the mixing of these control functions. These functions may be further modified by varying the RPM of the transmission joint relative to the rotor RPM to provide independent control of the pitch adjustment of each blade, either lower or higher than 1/rev. The system may be applied to a radio controlled model helicopter, and helicopters having single or contra-rotating rotors.

Description

t ' ROTOR This invention relates to rotating wing aircraft.

Rotating wing aircraft e.g. helicopters, autogyros and tilt-

rotor aircraft are well known, and typically comprise a fuselage with a rotor assembly, known as a rotor head.

The rotor head controls the pitch of the blades during flight via the rotor hub, to provide control of directional flight and lift, as required.

The helicopter satisfies its specialist market due to its unrivalled fuel efficiency compared to other types of VTOL aircraft, e.g. its nearest competitor the tilt-wing, is less than half as efficient. But the helicopter does suffer a high maintenance requirement, and considerable vibration and this prevents it competing on a wider scale, against comparable fixed wing aircraft.

For it to improve on its position in the world market, the helicopter must therefore reduce the high maintenance requirement, and reduce vibration levels to those of the comparable fixed wing aircraft.

It is known that such performance improvements are achievable, through the use of multi-cycling [greater than the 1/rev] pitch control techniques. However, it has not been possible to incorporate these improved methods of control in the mainstream, due to the poor levels of safety such research systems suffer, caused by their additional complexity, cost and maintenance.

The invention will provide this performance breakthrough, by disposing of the complex existing control systems and their maintenance presently used, and replacing them with a new type of constant velocity pitch control mechanism, to permit

the fullest use of these newer, performance improving control techniques. It will at once provide a system with superior control redundancy to those currently used, whilst providing increased performance and comfort, on a helicopter with reduced cockpit workload and manufacturing cost.

The invention provides this performance breakthrough by permitting the easier and more effective application of Independant Blade Controlling IBC techniques, whilst allowing the pitch control system to be contained within the rotating system; a technique originally proposed in Patent GB 9824075.7; the basis of Patent PCT/GB99/03647.

The invention is contained within the rotor hub to provide individual pitch control to each rotor blade, and dispose of the outmoded pitch control systems currently used, which control pitch via one common tiltdisc, positioned about the rotor mast, these being coupled to both the fuselage and rotating system, and operating substantially perpendicular to the rotating system controlled.

To further clarify the reasons for removal of present control systems, I provide a brief description of existing systems

and their control technique that prevents their practical full time use on IBC systems.

Such systems use a common centrally positioned, tilt disc mechanism, which has been used since the very beginnings of the helicopter over 60yrs ago, they have as their basis a fixed fuselage attachment point, this being coupled to the rotor head via a tilt disc attached to the rotating system.

Through this coupling, they suffer oscillatory feedback proportional to control input and blade reaction.

In addition to the reactively derived vibration levels such systems suffer, these systems are unsatisfactory for the

implementation of the more advanced blade pitch controlling techniques sought, due to their use of the single [common to all blades] centrally positioned, controlling tilt-disc, most typically of awash-plate type, which due to its method of operation can only provide pitch control at 1/rev.

Research areas of the industry have been aware of this failing for many years, and have tried to find a way of modifying the 1/rev output of the helicopter. Such research has generally centred upon the modification of the existing tilt-plate technology, and whilst such research has proven the advantages of multi-cycling pitch control over mono-

cycling, it has also demonstrated that it is not possible to develop a system, capable of competing with the conventional existing helicopter due to its complexity, resulting in unsatisfactory levels of safety, maintenance and cost.

Where used, such technology remains the exclusive domain of the military, as only they can justify the high operating costs, and questionable safety of such aircraft.

Therefore, it is the object of this invention to remove the single tiltplate pitch control mechanism of the current helicopter and replace it with a new type of constant velocity pitch control mechanism, to deliver a helicopter inherently possessing reduced vibration and reduced rotor mast drag, which can fully utilise the many advantages of multi-cycling blade control technology.

Furthermore, the invention, in disposing of the fixed fuselage mounted control link, will at once remove the reaction derived vibration and noise, whilst removing shaft drag power loss.

Removal of the shaft drag will release more power from the engine to improve overall performance, with the additional benefit of improved autorotation performance to increase safety, furthermore, system safety and efficiency can be

increased still further by the use of more than one pitch control mechanism per blade if required.

The invention may be signalled and motivated by any means suitable, and is described using electronic signalling.

The control of the pitch control system may be configured to provide variable pitch control output greater or less than 1/rev, in linear or non linear form, to permit multiple cycling of pitch control, which may be selectable or full-

time in operation. This may be achieved by controlling the rotational speed of the constant velocity pitch control mechanism, relative to rotor RPM, and may be combined with variation of mechanism tilt angle.

Such a system would permit the development of a true rigid rotor rotating wing aircraft, possessing all the attributes of the articulated rotor, without its complexity.

All of the blades required movements, on such a rotor typically flapping, lead and lag, and its associated damping, may be achieved via the simple and safe expedient, of controlling pitch cycling motor RPM, superimposed upon the rotational speed of the rotor to vary blade pitch phasing, this technique may also be used for steering the aircraft, and may be combined or operated in isolation.

Said system may be selectable or of automatic self-adaptive type, typically providing; manocuverability enhancements, blade tracking adjustment, blade-pitch adjustment to blade out of balance conditions, adjustment to gust effects, adjustment to deep stall conditions [to further improve aircraft safety], pressure change or other imperfect airflow variations encountered by the rotor in operation.

A helicopter so equipped, will be more silent in operation, due to reduced aerodynamic noise, and will possess reduced vibration, accompanied by stall free blade operation, with attendant improvements to operating efficiency, this being

further improved by reduced rotor mast drag, resulting in improvements to autorotational safety and efficiency, whilst the reduced component count of the system will permit reduced cost of manufacture, and improved control redundancy accompanied by reduced maintenance and weight.

Present mechanisms conventionally used, are coupled between the fixed component [fuselage], and the rotating component [rotor head], via a awash-plate [tilt-plate] mechanism, or derivatives thereof. Such mechanisms are interactive, and suffer oscillatory feedback, in the form of vibration, this caused by their controlling actions to each blade of the rotor head. Such vibratory action is variable, being proportional to torque and pitch control inputs.

Additionally, present hub control systems, restrict the freer rotation of the rotor main shaft, to waste engine power, and reduce autorotational performance) caused by the mechanical interaction of present mechanisms, between the fixed frame [fuselage] and rotating frame [rotor head].

The invention makes it possible to remove both the reaction derived vibration and shaft drag, by replacing the existing blade pitch control systems used, with a rotor head equipped with self-contained power and controlling means, said rotor head being controlled by external signal, most typically receiving control signalling through the use of fibreoptic signalling techniques via a rotating joint, wherein a signal transmission coupling is provided between transmitter and receiver rings positioned about the rotor main shaft, to provide a signal coupling without mechanical contact; a technique originally proposed in Patent GB 2090214 B [Controlling Helicopter Rotors], and later specified in Patent Application GO 2247089 A [Fibre Optical Rotating Joint]. The invention may obtain power for its integral control system, by utilizing the rotation of said head or heads to motivate an integral system of power generation, such system

may use a variety of mediums, either singly or in combination; ie, electronic, hydraulic, pneumatic, electro/hydraulic. The controlling signal system may use a fibre-optic rotating joint; such joint comprising a first and second part, rotatable with respect to each other about a common axis. A free space path existing between the two parts, each part having a plurality of coupling lenses positioned to form a circle, whilst each lens is connected to an optical coupler, by a respective pigtail optical fibre. The optical coupler will be connected to a wavelength division multiplexer, which is connected to an optical transmitter and an optical receiver. Resulting in a transmission path between each part via the optical rotating joint without mechanical contact.

According to the present invention, there is provided a rotor hub for a rotating wing aircraft including means for controlling blade pitch, wherein the blade pitch controlling means comprises a rotatable member with slaveably attached rotor and follower, having an axis of rotation which is pivotable, being contained within and motivated by its housing body, said follower being attached to at least a part of the blade for controlling the pitch thereof.

A plurality of the invention may be used for the individual control of each blade, and the invention may be positioned anywhere on the rotor head.

The invention may be contained within an externally signalled rotor hub control system, which may have self contained motivational means, for the control of a helicopter rotor, wherein the pitch of each blade may be controlled independently, or synchronously with the other blades in the system, or in variable synchronicity with the other blades in the system, and which may be controlled without reactive controlling input, between fuselage and rotor.

The invention is shown with two blades configured for blade pitch control, using aerodynamic blade pitch controlling means, though the invention may be used with any quantity of blades, and in any position or various positions, or in any combination on the hub or blade.

Furthermore the invention may be applied to the more commonly used blade root twisting, pitch control system.

An embodiment of the invention will now be described by way of example with reference to the accompanying drawings, in which: Figure 1 shows a perspective view of the system applied to a model helicopter; Figure 2 shows schematically, the system applied to a model helicopter; Figure 3 shows in exploded perspective, the invention in a helicopter rotor head application; Figure 4 shows a perspective view of a pitch control assembly installed in a hub arm; Figure 5 shows in cut away perspective, the pitch control mechanism; Figure 6 shows in schematic perspective, the layout of the pitch control system; Figure 7 shows in schematic perspective, a pitch control mechanism; Figure 8 shows in schematic perspective, a pitch control mechanism in collective position; Figure 9 shows in schematic perspective, a pitch control mechanism in cycling mode;

Figure 10 shows in schematic perspective, a pitch control mechanism in tilted cycling position; Figure 11 shows in perspective, a pitch control actuator in collective pitch position, with cyclic pitch superimposed in negative output position connected to a crank mechanism; Figure 12 shows in perspective, a detail view of a pitch control actuator with positive cycled pitch and its relative position to its housing; Figure 13 shows in perspective, a detail view of a pitch control actuator with negative cycled pitch and its relative position to its housing; Figure 14 shows a perspective view of the system applied to a helicopter; Figure 15 shows in schematic perspective the system applied to a helicopter; Figure 16 shows a perspective view of a rotor hub applied to a helicopter; Figure 17 shows a perspective view of the system applied to a helicopter of co- axial type.

Figure 18 shows in schematic perspective, the system applied to a helicopter of co-axial type.

Referring to the drawings, the helicopter shown in Figure 1 includes a non-rotating component in the form of a fuselage 2 with rotating component in the form of an overhead rotor assembly 3, attached to rotatably driven rotor mast 10

containing fixed shaft 470 mounted within fuselage 2; fixed shaft 470 being non rotatably secured to fuselage 2.

Rotor mast 10 is driven by engine 30 by way of a belt 20. A rotor hub 150 is mounted at the upper end of rotor mast 10, and is provided with electrical power by means of an annular generator 110, such as a brushless induction generator, mounted around rotor mast 10. As shown in Figure 2, power from generator 110 is supplied via power supply line 111 and 112; power supply line 111 contained within rotor mast 10, supplies electrical power to accumulator 300 within rotor hub 150, whilst power supply line 112 supplies power to fuselage installed accumulator 300.

Control signals for the helicopter originate from a radio transmitter 50 which transmits control signals 51 and 52 respectively, to control computers 455 and 457, installed within rotor head 150 and fuselage 3 respectively. In response to signals from the radio transmitter 50, control signals are generated by control computers 455 and 457 which send control signals to pitch control unit 460 via signal lines 456, and engine control actuator 280 and tail rotor control actuator 290 via signal lines 2351 260 and 270 respectively. As can be seen from Figure 3, the rotor hub body comprises hub casing 420 attached to rotor mast 10, and two side covers 440; attached to each hub casing 420 are hub arms 430, to which are attached blade assemblies 340, each comprising blade grip 338, rotor blade 339, pushrod 336, servo tab 337 and pivot block 335. Blade grip 338 is pivotally attached to pivot block 335 on its pitch rotation axis to permit pitch action 452,and is of fully articulated type.

Contained within rotor hub 150, are batteries 300, control computer 455 retained by cover plate 450, signal lines 456, controller 460, fixed gear wheel 465 and driven gear wheels 480. Just visible is one of two control arms 485, attached to controller 460, this also being shown in Figure 4 which shows

in more detail the location of pitch control assembly 500 in hub arm 430, and pushrod 336 attached to servo-tab 337 on the rotor blade 339.

As can be seen from Figures 2, 3 and 4, contained within the rotor hub 150 are rotatably driven pitch control assemblies 500, contained within each hub arm 430, these being connected to pitch control unit 460 via control arms 485. Each pitch control assembly 500 is controllably adjusted by pitch control unit 460, in turn responding to signal input from receiver control computer 455.

Detailed in Figure 5 is pitch control assembly 500, installed in hub casing 420; comprising gear wheel-driven 480, control arm 485, housing 487, bearings 488, follower rotor 489, follower 490, spring 491, pushrod 336, follower rotor bearing 493, screw rotor securing 494, constantvelocity drive ball 495, ball bearings 496, rotor finger 497, gear wheel drive 498, motor 492 and sliding housing 499.

The system is shown schematically in Figure 6, wherein the c/v drive balls 495 are shown attached to pitch control unit 460 via control arms 485. Pitch control unit 460 provides controlling output by three separate and/or mixable control adjustments. Vertical adjustments 6A and 6B, rotational adjustment 6C, and lateral adjustment 6H, this to radially position control arm 485 relative to the hub axis via attachment points 900. The rotational direction ED of each c/v drive ball 495, and their phased output positions 6E and 6F are shown, these being diametrically opposite at 180 , for the two blade system shown, whereas for example, on a three blade system this position would be phased at 120 , and a four blade system at 90 and so on.

Importantly this phase angle may be variably adjusted also or independently by adjustment of the rotational speed of c/v drive ball 495, such adjustment may be synchronous or non synchronous and may be implemented to control each blade

individually or in variable synchronicity with rotor [speed of rotation] (RPM).

Figure 7 shows in greater detail c/v drive ball 495 in neutral position 6E. Drive ball 495 is driven in the direction 6D, which in turn controllably drives enslavably linked rotor finger 497 and follower rotor 489, attached to follower 490, which imparts pitch controlling output 7D to pushrod 336 (reference fig 5) attached to servo-tab 337 (reference fig 1). Output 7D may be adjusted by controllably moving axis 7G in directions 7C, to increase or reduce collective pitch setting.

Referring to Figure 8, c/v drive ball 495 is shown in COLLECTIVE output position, with control arm 485 aligned with axis 7E. COLLECTIVE pitch adjustment being obtained by adjusting axis 7G in directions 7C, as described in Figure 7, such adjustment varying the position of pitch controlling follower 490, connected to each blade via pushrod 336.

Figure 9 shows CYCLED pitch output +0 to - resulting from adjustment of control arm 485 on axis Z. resulting in phased output positions 6L and 6F per revolution of c/v drive ball 495, operating at substantially the same RPM as rotor 3.

Figure 10 shows the CYCLED output axis 7G tilted from the horizontal, by adjustably positioning control arm 485 between axis' Z and Y. with c/v drive ball 495 operating at substantially the same RPM as rotor 3, thus causing a change in blade pitch phasing, to alter the vector angle of the rotor disc, to suit control requirements typically directional flight control.

Thus COLLECTIVE pitch control is achieved by sliding adjustment of c/v drive ball 495 on datum 7E, whilst CYCLED pitch output for any given pitch setting is achieved by adjustment of control arm 485, and/or adjustment to the rotational speed of c/v drive ball 495 relative to rotor 3

RPM, to VARY THE PHASED POSITION OF EACH BLADE INDIVIDUALLY, in variable synchronicity, synchronously or independently.

Figure 11 shows control assembly 500, providing negative cycled output at phased output position 6F (reference Fig 9), to pitch controlling crank 905, said output being superimposed over collective pitch setting 7C.

Figures 12 and 13 show the cycled action of follower rotor 489 at phased position 6L and 6F respectively, and its positive and negative movements 7M and 7N relative to sliding housing 499 adjusted collective position.

The helicopter shown in Figure 14 includes a fixed component in the form of a fuselage 2 with rotating component in the form of an overhead rotor blade assembly 3. A rotor mast 10 is mounted within the fuselage 2, and is driven by combined engine and transmission unit 30, by way of belt drive 20. A rotor hub 100 is mounted at the upper end of rotor mast 10, and is provided with electrical power, by means of an annular generator 110, such as a brushless induction generator, mounted around the rotor mast 10.

As shown in Figure 15, power from generator 110 is fed by way of power supply line 111, extending within rotor mast 10 to battery 300 within rotor hub 100. Control signals for the helicopter originate from a cockpit control unit shown diagrammatically as 230 and pass by way of signal line 235 to a processor or computer 220.

In response to signals from the cockpit control unit 230, control signals are generated by the processor 220 and are transmitted to the rotor head 100 along signal lines 235 and 240, control ring 210 and signal line 320, within the rotor mast 10. Further control signals are transmitted by processor 220 through signal lines 260 and 270, to engine control actuator 280, and tail rotor control actuator 290 respectively.

; The rotor hub shown in Figure 16 is a modification of that shown in Figures 1, 2 and 3, wherein the fixed gear wheel 465 has been supplanted by drive motor assemblies 4g2, and the same references are retained to denote the same or similar parts. Drive motor assemblies 492, are linked to and drive gear wheels 480 via gearwheels 498, these being electronically signalled and motivated via lines 451 from pitch control unit 460, which in turn receives control signalling via signal line 320, this being provided with electronic supply via accumulator 300.

Power supply to accumulator 300 is delivered through line 111, within rotor mast 10, and is generated by means of annular generator 110, mounted around the rotor mast 10.

Figure 17 shows the invention in co-axial, contra-rotating form applied to a helicopter including a fixed component in the form of a fuselage 1, without a tail rotor, with two co-

axial overhead rotor blade assemblies 3 and 4.

Rotor mast 10 is mounted within fuselage 1, and is attached to rotor hub 180, being driven by combined engine and transmission unit 35, by way of belt drive 20, in anti-

clockwise direction.

Rotor mast 15 passes through rotor mast 10, and is also mounted within fuselage 1, passing through rotor hub 180, its upper end being attached to rotor hub 160, and is driven in clockwise direction by belt drive 25. Both rotor hubs 160 and 180 are provided with electrical power, by means of annular generators 110 and 112 respectively, mounted around rotor masts 10 and 15.

As shown in Figure 18, power from generators 110 and 112 is fed by way of power supply lines 111 and 815, extending within rotor masts 10 and 15 respectively to batteries 300, within rotor hubs 180 and 160.

( Control signals for the helicopter originate from cockpit control unit 230, and pass by way of signal line 235 to a processor or computer 220.

In response to signals from the cockpit control unit 230, control signals are generated by processor 220 and are transmitted to rotor heads 160 and 180, along signal line 235 to processor 220, and thence to signal lines 240 and 246 to control rings 210 and 211, which transmit control signals via signal lines 320 and 865 to rotor heads 180 and 160 respectively, by way of signal lines 260 and 27G. Further control signals are transmitted by processor 220 through signal line 260 to engine control actuator 280.

The rotor hubs 160 and 180 are a modifications of that shown in Figures 16, and the same references are retained to denote the same or similar parts.

Rotor hub 180 is modified to allow rotor mast IS to pass through it, and control unit 461 is a modification of control unit 460.

Rotor hub 160 is modified to suit its clockwise rotation, by suitable control signalling to drive motor assemblies 492, whilst rotor blade assembly 4, incorporates rotor blades 341, for clockwise [opposite hand] rotation, to rotor assembly 4.

Claims (11)

-- <o CLAIMS

1. A rotor for a rotating wing aircraft, including means for controlling blade pitch, characterized in that the blade pitch controlling means comprises an axially and angularly adjustable transmission joint [495], enslavably linked to a follower [490], connected to at least a part of the blade [339] r [340] for controlling the pitch thereof.

2. A rotor as claimed in claim 1, characterized in that the axis of rotation of a transmission joint [495], is generally perpendicular to the axis of rotation of a rotor hub [100], [150], [160], [130].

3. A rotor as claimed in claim 1 or 2, characterized in that a transmission joint [495], is contained within a housing [487], rotated in variable synchronism with a rotor hub [100], [150], t160], [180].

A rotor as claimed in any preceding claim, characterized in that a transmission joint [495], is contained within a housing [487], motivated by fixed wheel [465] and driven wheel [480].

5. A rotor as claimed in any preceding claim, characterized in that lateral adjustment [6H] is provided for moving a transmission joint [495], in an axial direction on axis [7E].

6. A rotor as claimed in claim 5, characterized in that the lateral adjustment [6H] for axial movement on axis [7E], is provided by a control unit [460].

7. A rotor as claimed in any preceding claim in that a transmission joint [495], is pivotable in a plane which includes the axis of rotation of a rotor hub [100], [150], [160], [180].

8. A rotor as claimed in any preceding claim in that pivoting of a transmission joint [495], is effected by vertical adjustments [6A] and [6B], provided by a control unit [460].

9. A rotor as claimed in any preceding claim in that pivoting of a transmission joint [495], is effected by rotational adjustment [6C], provided by a control unit [46G].

1Q. A rotor as claimed in any preceding claims, characterized in that pivoting of a transmission joint [495], is effected by controlling movement of a control arm [485] extending axially therefrom.

11. A rotor substantially as herein described and illustrated in the accompanying drawings.