GB2299562A - Actuator for helicopter rotor blade aileron - Google Patents

Abstract

Lateral movement of shaft 100, equipped with spiral splines 105, imparts rotational action to sleeves 140, and thence to aileron 200. The shaft 100, is primarily actuated by fluid pressure acting upon pistons 120, while end caps 20, 21 comprise electronically activated rods 25 to actuate the shaft 100 should the fluid pressure actuation fail. Control by light and/or electronics may be involved. The actuator is of compact profile, possessing the minimum of moving parts, whilst dispensing with the conventional use of cranks, pushrods, gears and the like, with benefits to weight, cost, reliability, manufacture, positional accuracy and speed of response. Patents GB 2090214B, GB 2280412A, GB 2280162A, GB 2276958A and GB 9504355.0 are referred to.

Description

FAST HESPONSE ACTUATOR This invention relates to an actuator controlled by electrical and or optical signalling for the control and actuation of aileron equipped rotor blades, as described in UK PATENTS GB 20902148, GB 2276958A, GO 2280162A, and GO 22B0412A.

Rotating wing aircraft are well known, eg helicopters and comprise a fuselage with overhead driven blades,with means for controlling and varying the angular pitch of those blades.

Existing means for controlling and varying the angular pitch, rely on a mechanical linkage system known as the swash-plate, which when operating, re-acts against the airframe or chassis.

The swash-plate is a reactive mechanism, and therefore receives and transmits forces as a consequence of its pitch controlling inputs to the blades, resulting in the feedback of these re-actions to the airframe in the form of vibration.

Furthermore it also has a poor control of rotor blade aerodynamics, due to control inputs emanating from the blade at its root end, the position of least effectl enforced through use of the swash-plate, and this allows, indeed encourages blade stall and flutter to occur at the tip of the blade, with resultant feed-back to the airframe in the form of further vibration over and above the re-active vibrations.

Blade stall is commonplace on helicopters and would not be tolerated on fixed-wing aircraft.

It is a potentionally destructive action, and is something blade design has to allow for, thereby seriously compromising its performance still further! The poor control of the blades tip, in turn causes poor control of the rotor disc, resulting in poor control of the gyroscopic couple between rotor and airframe, causing yet more vibrationary feedback to the airframe.

The swash-plate and its rotor hub is a complex mechanism, particularly on co-axial rotor assemblies and is aerodynamically unclean, this complexity also results in high production costs, with attendant maintenance requirements. Modern materials have allowed the simplification of the swash-plate, though these types have not demonstrated a useful performance improvement, so much as a reduction to maintenance.

From the foregoing the swash-plate can be summarised, as a reactive control mechanism, with poor control of blade aerodynamics, resulting in increased weight of both hub and blades, accompanied by high vibration levels,(with a high actuating force requirement up to seven times greater than an aileron controlled blade, ref NASA Report 4611) whilst suffering high production and maintenance costs.

Clearly the removal of the swash-plate and its complex mechanics would remove all the faults attributable to its use.

This can be achieved by controlling blade pitch or feathering aerodynamically, through the use of an aileron or servo-tab positioned on the blades to suit design requirements, and which may require a plurality of said ailerons for each blade, these being controlled via an opto-electronic (eg laser) rotating coupling or joint, operating in similar fashion to the commonplace CD player, and described in greater detail in patents GO 20902148 GE 2276958A and G9 2280412A, thereby removing the pitch controlling mechanical link between airframe and rotor head.

The technologies of opto-electronic transmission and aerodynamic blade control are both proven, and their suitable development will provide a breakthrough in rotorcraft performance, whilst reducing manufacturing and maintenance costs.

The key element to the attainment of this performance breakthrough will be the development of an aileron controlling actuator, that is positionally controlled without mechanical contact or reaction with the fixed non-rotating frame of the helicopter, whilst possessing control integrity sufficient to satisfy flight safety redundancy requirements, and this failsafe requirement will be highlighted by the industries stance on what it sees as a high risk technology, despite the evidence.

The fast-response actuator will satisfy these requirements, and its simplicity of design will further assist in this, however no mention of aileron control would be complete without a brief mention of the history of aileron control, as applied to helicopters past and present.

The proposed system of aileron actuation is unique, though aileron control on helicopters is not without precedent, having been used by pioneers in the USA, such as Landgraf, who used ailerons to control twin hingeless rotors in the 1940's, and Kaman in the 1950's, who developed his servo-tab system of blade feathering, and later formed the company which takes his name, and which he runs to this day.

These Kaman aircraft have been so successful, that until recently their use was restricted exclusively to the US military, reinforced by a trade embargo preventing their sale outside the USA.

This trade restriction was lifted 4yrs ago, though the type still remains largely unknown, as does its unique control system.

Ample testimony to the advantages of controlling rotor blades aerodynamically, is provided by the rated blade life of the Kaman Seasprite, at 10 000hrs, whilst the type is known to provide smoother flight, and does not require the use of powered assistance for control authority,(due to reduced blade pitch controlling inputs) such is the efficiency of aerodynamic blade control by aileron.

This efficiency has been further acknowledged, by the recent National Aeronautics and Space Administration (NASA) Report 4611 of June 1994, and I make no apology for repeating, that pitch controlling forces are 7 times lower than swash-plate hub control systems, or to put it another way, 7 times more efficient! The high system redundancy of the fast-response actuator, with its alternative fail-safe paths of motivation, combined with its ability to withstand the harsh operating environment of the rotor blade, will enable it to achieve redundancy factors of the high orders necessary for such an application, and permitting its use with suitably adapted 'fixed-wing' fly-by-wire technology, to encourage its fastest and most economical development.

The improved aerodynamic performance offered by aileron blade control, with its fast control response, will encourage its use with the new programmeable control technologies being developed worldwide, such as 'multi-cyclic'.

This form of control enables the aerodynamic performance of the blades to be improved, by varying the angular pitch of the blades several times per revolution, as compared to the single graduated pitch change per revolution, for a monocyclic system.

Interestingly the development of multi-cyclic control, has arisen due to the deficiencies of present blade control techniques using hub based control inputs , via the swash-plate mechanism, and the aforementioned NASA Report proves that aileron controlled rotors operating in mono-cyclic mode, are compareable in performance to the complex swash-plate systems operating in multi-cyclic.' whilst the complexity and unacceptable sytem safety of these multi-cyclic swashplate systems, has so far prevented the introduction of this technology.

It is to remove this complexity, and provide satisfactory response action combined with alternative fail-safe paths of control motivation, that the fast-response actuator has been invented.

It will dispense with present complex hub mechanics, by using light in the form of transmitted signals, for positional control of the fast-response actuator passed from the fixed non-rotating frame of the helicopter chassis, via an optical coupling to a suitably controlled (eg switched) controlling unit in the hub or blades.

Further redundancy safety, with the beneficial effect of even greater simplification of the actuator, may be obtained by using the centrifugal loads present in helicopter rotors or similar rotating mechanisms, to provide motivational means to the sliding action of the actuators sliding shaft, thus increasing operational safety by allowing the use of controlled motivational means to one end of said actuator.

The unique design of the fast-response actuator, will allow its use on applications other than helicopters, eg fixedwing aircraft and marine vessels of powered or sailing type, where a compact actuator may be used to advantage to control controlling surfaces such as rudders,fin(trim), wing-sails etc.

Further use for the actuator will be in the field of model or small radio control equipment, its simplicity of design permitting easy miniaturisation.

Alternative types of motivation may be used to suit design requirements, in both full size and model use, eg fluidic, electric, mechanical, or a combination of these.

According to the present invention, there is provided a fastresponse actuator, attached to and positionally controlling an aileron or servo-tab control surface. Characterised in that it derives its rotational surface controlling action, from the lateral movement of a central shaft said shaft being motivationally controlled at either end,both ends or from the centre.

A specific embodiment of the invention will now be described by way of example with reference to the accompanying drawings in which: Figure 1 shows in third angle orthographic projection, the actuator with scale references; Figure 2 shows in perspective, the components of the actuator exploded for clarity; Figure 3 shows in perspective, a typical aileron installation on the tip of a helicopter rotor blade; Figure 4 shows a typical section through the fast-response actuator installation; Figure 5 shows in perspective an actuator with aileron, the aileron having angled ends for use on a swept rotor blade tip.

Figuer 6 shows a helicopter with co-axial rotors, equipped and controlled by tip mounted ailerons.

Referring to the drawings, the fast-response actuator is shown and described in Figs 1, 2, 3 and 4, and shown operationally installed in Fig 5 as a positional controller for an aileron assy 250, of integral construction with actuator assy 600, to form aileron actuator assy 700.

As shown in Fig 2, the actuator comprises actuator body 10, containing shaft 100 equipped with pistons 120 and seals 125, and housed in bores 15, positioned at either end of actuator body 10, these pistons being motivated and positionally controlled fluidically, via ports 400 and 410.

The outer ends of bores 15 are sealed with actuating caps 20 and 21, these being equipped with an electronically motivated actuating rod 25, said rods being operably extendible for emergency use, to maintain the sliding action of shaft 100 in the event of failure of primary motivational source, supplied through ports 400 and 410, electrical supply is delivered by wires 23. The actuating caps 20 and 21 also serve as pivotal shafts for the outer halves of the aileron assy 250, the outer circumference of the actuator rod 25 housing 22, serving this dual purpose, with aileron pivot bearings 40, and aileron thrust bearings 130 mounted thereon. The electric supply wires pass through holes 60, and are attached to suitable connectors adjacent to ports 400 and 410, and engage with matching connectors 24, 401 and 411 shown in Fig 3.These connections being equipped with seals 3 and 4, to provide waterproofing and maintain system integrity of fluidic and electric means.

Fig 4 is a section of the rotor blade, showing the arrangement of such supply delivered through blade spar 500, and showing the electric supply wires 27 with fluidic supply tube 401, these being encapsulated in a suitable substance, to prevent chafe during operation.

Referring to Figs 3 and 4 the actuator assy 700, is securably located on the rotor blade, in aileron actuator housing 11, the actuator aileron assy 700 being a push fit into said housing, and engaging a series of matching ports in housing block 800, provided with suitable recesses for seals 3 and 4. The aileron actuator assy is secured by screws 810.

The aileron derives its twisting (rotating) action from the lateral movement of sliding shaft 100, this shaft being provided with flutes or splines of spiral or helix forml05.

Said shaft is prevented from rotational movement by pins 110, so that its lateral sliding movement 98 and 99, will cause the shaft splines to impart a twisting action to the sleeves 140, in which it is housed. Each sleeve 140 has an internal groove of helix or spiral form, to match and engage the flutes of spiral form on shaft 100.

Preventing sideways movement to the sleeves 140, causes them to rotate when the helix splines of shaft 100, with which they are engaged, are moved sideways, such movement being provided by controllable motivation in the form of hydraulic supply, reacting against pistons 120 at each end of said shaft.

Shown also in ghosted form are spacers 150, these being provided to prevent the lateral movement of sleeves 140.

Sleeves 140 are shown in three groups of three, to optimise the frictional versus thrust movements of said sleeves, their quantities being adjustable to assist this optimisation, and to simplify production, by enabling the easier manufacture of the internal helix.

Sleeves 140 are locked radially by splines or grooves on their outer circumferance, these engaging matching grooves 230, inside and integral with aileron body 200.

Each end of the helix formed length of shaft 100, is supported in bearing assys 115, providing friction free lateral and rotational movement through the use of two sets of balls, one set bearing on the helix of said shaft, and the other bearing on the periphery of housing 12 in body 10.

To prevent endwise frictional loss between aileron body 200, and face 10a on body 10, thrust bearings 135 are used.

Lubrication to all moving parts is provided by oil, taken from the hydraulic supply, via internal oilways in shaft 100 supplied by holes 160 and 170.

The ends of aileron 200, are pivotally supported by bearings 140,whilst lateral movement of said aileron is assisted by thrust bearing 130, these bearings being housed and contained respectively by aileron tip 220, this being secured by screw 215 and suitable 'keying arrangement'. End cap 210 is a snap fit and is provided to prevent the ingress of dirt and moisture.

Claims (15)

CLAIlIS

1 A fast-response actuator, in which controlling output actions are derived from the motivation and control of sliding shaft inputs.

2 A fast-response actuator, in which controlling output actions are obtained from lateral inputs or rotary inputs, said inputs providing respectively, rotary or linear outputs.

3 A fast-response actuator, in which rotational output action is derived from a controllably motivated, sliding shaft.

4 A fast-response actuator, as claimed in Claim 1, Claim 2, and Claim 3, which may obtain motivation partially or wholely, by fluidic means.

5 A fast-response actuator, as claimed in any preceding claim, which may obtain motivation partially or wholely, by electronic means.

6 A fast-response actuator, as claimed in any preceding claim, which may obtain motivation partially or wholely, by mechanical means.

7 A fast-response actuator, as claimed in any preceding claim, which may obtain motivation partially or wholely, by photonic means.

8 A fast-response actuator, as claimed in any preceding claim, which may obtain motivation partially or wholely, by photonic-electronic means.

9 A fast-response actuator, as claimed in any preceding claim, which may obtain motivation partially or wholely, by electronic and -mechanical means.

10 A fast-response actuator, as claimed in any preceding claim, which may obtain motivation partially or wholely, by electronic, photonic, mechanical and fluidic means.

11 A fast-response actuator, substantially as described with reference to Fig 1 of the accompanying drawings.

12 A fast-response actuator substantially as described with reference to Fig 2 of the accompanying drawings.

13 A fast-response actuator, substantially as described with reference to Figs 3 and 4 of the accompanying drawings.

14 A fast-response actuator, substantially as described with reference to Fig 5 of the accompanying drawings.

15 A fast-response actuator substantially as described with reference to Fig 6 of the accompanying drawings.