WO2000021828A1

WO2000021828A1 - Directional control and aerofoil system for aircraft

Info

Publication number: WO2000021828A1
Application number: PCT/AU1999/000874
Authority: WO
Inventors: Paul Vincent Ladd
Original assignee: Paul Vincent Ladd
Priority date: 1998-10-13
Filing date: 1999-10-13
Publication date: 2000-04-20
Also published as: AUPP646098A0

Abstract

Aerofoil surfaces such as wing (10), tailplanes, rudders, etc may be mounted to an aircraft for rotation about axes (12) normal to the longitudinal axis of the aircraft and such that the centre of pressure (p) of the aerofoil surface is rearward of the axis of rotation. Bias means such as springs (15) urge the aerofoil towards a central position while permitting limited rotation of the aerofoil in response to aerodynamic forces. Conventionally controlled surfaces such as canard wings (25) or canard rudders, initiate directional changes.

Description

TITLE: DIRECTION CONTROL AND AEROFOIL SYSTEM FOR AIRCRAFT Field of the Invention

The present invention relates to aircraft, and more particularly to the directional stability and control of aircraft.

The invention has been developed primarily for use in relation to commercial passenger aircraft, and will be described hereinafter with reference to this application. It will be appreciated, however, that the invention is not limited to this particular field of use, being also applicable to a wide variety of other aircraft including military, as well as smaller recreational aircraft.

Background of the Invention

With the increasing importance of air travel as a mode of global transportation, and in an increasingly competitive global market, there is a growing need to optimise the performance potential of commercial aircraft in a number of areas including manoeuvrability, speed, load carrying capacity, fuel efficiency, and comfort. The same

factors in general terms are equally relevant to military and recreational craft.

Most conventional aircraft include wings and a tail assembly which act in conjunction with the elongate fuselage to provide lift, directional stability and control.

More specifically, the wings of an aircraft typically include ailerons to control rotation about the roll axis and flaps to control lift as well as pitch. The tail assembly typically includes a horizontally oriented stabiliser or tail plane with elevators to control rotation about the pitch axis and a vertically oriented fin supporting a movable rudder to control rotation about the yaw axis. The structure and operation of these lift and control surfaces are well known to those skilled in the art, and so will riot be described in further detail. One significant limitation inherent in these conventional aerodynamic structures is that acceleration beyond certain critical limits, for example during sharp turning manoeuvres, can cause one or more of the lift or control surfaces to stall. This can also

occur during relatively low speed manoeuvres, particularly take-offs and landings, where

in order to generate sufficient lift, the control surfaces are presented to the incident air

stream at a relatively steep angle of attack. When a stall condition is induced, the air

flow around the stalled lift or control surface, which is normally smooth and streamlined,

delaminates and breaks into unstable turbulence. This in turn causes the efficiency of the

aerofoil surface to be dramatically reduced. This results in a loss of manoeuvrability,

increased power requirements to maintain momentum, and increased fuel consumption. In some cases, the mimmum degree of lift and control necessary for stable flight cannot

be maintained during a stall. This has potentially catastrophic consequences.

Another disadvantage with conventional control surfaces is that they are not able

to respond sufficiently quickly to changes in air density, pressure, currents and the like to

counteract the turbulent effect on the aircraft as it moves rapidly through these changing

atmospheric conditions. The result is buffeting and discomfort for the passengers and crew within the craft. While the problem can be overcome to some extent using

computer controlled automatic pilots with rapid response times, the effect is not

eliminated entirely. A further disadvantage with conventional lift and control surfaces is that because

of the propensity to stall, the general lack of responsiveness, and the practical limit to structural strength, modern aircraft lack manoeuvrability, particularly at relatively high

speed. Disclosure of the Invention

It is an object of the present invention to overcome or ameliorate one or more of

the disadvantages of the prior art, or at least to provide a useful alternative.

A control system for an aircraft, said control system including:

an aerofoil surface;

mounting means connecting the aerofoil surface to the aircraft for rotation about an

axis generally normal to a longitudinal axis of the aircraft such that the effective centre

of pressure of the aerofoil surface is spaced rearwardly from its axis of rotation; and

bias means operable to urge the aerofoil surface toward a central rest position

while permitting limited rotational movement of the control surface away from the

central rest position in response to unbalanced pressure loadings whereby unbalanced

aerodynamic reaction pressures acting on the aerofoil surface tend automatically to effect

a corresponding rotation against a restoring force provided by the bias means

By virtue of this arrangement, unbalanced aerodynamic reaction pressures acting

on the aerofoil tend automatically to effect a corresponding rotation against a restoring

force provided by the bias means. It is believed that this increases manoeuvrability and

delays the onset of a stall condition which would otherwise result in a dramatic reduction

in the efficiency and effectiveness of the aerofoil.

The aerofoil according to one aspect of the invention preferably takes the form of a

control surface. The control surface may be oriented vertically, horizontally or at any

intermediate angle, and as such may be configured to operate during manoeuvres

involving roll, pitch or yaw. In one preferred embodiment, the control surface includes a rudder optionally

supported by a fin. In another embodiment, the control surface includes a stabiliser or tail plane, optionally fitted with elevators. A combination of both configurations is also contemplated.

In one preferred embodiment, conventional control surfaces adapted to initiate

directional changes are provided in the form of canard wings and canard rudders,

disposed toward the front of the aircraft, while control surfaces according to the present

invention are integrated into the tail assembly, toward the rear of the aircraft, to

complement the aerodynamic response of the forward control surfaces.

In another form, the invention is embodied to include the primary lifting formation

of the aircraft wherein the aerofoil surface forms part of the primary wing.

In one alternative form, lateral control surfaces according to the invention are

preferably formed integrally with the primary lifting surfaces or wings of the aircraft,

and as such may be disposed substantially further forward on the fuselage from the tail

assembly.

Brief Description of the Drawings

Preferred embodiments of the invention will now be described, by way of example

only, with reference to the accompanying drawings in which:

Figure 1 is a diagrammatic perspective view showing part of the tail section of a

fixed wing aircraft, wherein a conventional fin and rudder assembly has been replaced by

a vertically oriented control surface, according to a first embodiment of the invention;

Figure 1A shows a variation of the arrangement shown in Figure 1; Figure 2 is a diagrammatic plan view of an aircraft according to a second

embodiment of the invention, wherein conventional wings have been replaced by

horizontally oriented control surfaces and integral lifting surfaces according to the

invention for operation in conjunction with forwardly disposed canard wings;

Figure 2 A shows a variation of the arrangement shown in Figure 2;

Figure 3 is a side elevation of the aircraft of Figure 2, incorporating the rear fin and

rudder assembly of Figure 1, operable in conjunction with upper and lower canard

rudder assemblies disposed toward the front of the aircraft;

Figure 4 is an enlarged diagrammatic side elevation of the aircraft of Figures 2 and

3 , showing the pivotal capability of the primary wing formations;

Figure 5 is a cross-sectional front elevation of the aircraft of Figure 2 showing the

bias control mechanism for the primary wing formations in more detail;

Figures 6 to 10 show a series of diagrammatic plan views indicating the air flows

around the vertically oriented control surfaces, during a turning manoeuvre involving

rotation about the yaw axis of the aircraft;

Figures 11 to 17 show a series of diagrammatic side elevation views indicating the

air flows around the horizontally oriented control and lifting surfaces, during

manoeuvres involving changes in pitch of the aircraft;

Figure 18 is a diagrammatic perspective view showing an aircraft according to a

further embodiment of the invention, adapted for large scale passenger transportation;

Figure 18A shows a variation of the arrangement shown in Figure 18;

SIIBSTΓTUTE SHEET (Rule 26) (RO/AU) Figure 19 is a perspective view showing the aircraft of Figure 18 with the

undercarriage deployed, and the lower control surfaces retracted upwardly to provide adequate ground clearance for take-off and landing; and

Figure 20 is an enlarged diagrammatic view of one of the lower control surfaces of the aircraft of Figures 18 and 19, showing one embodiment of a retraction mechanism.

Figure 20 A shows a variation of the arrangement shown in Figure 20;

Figure 21 is a side elevation of the aircraft of Figure 3 showing a linear retraction

mechanism for the lower rearward control surface.

Preferred Embodiments of the Invention

Referring initially to Figure 1, the invention provides a directional control system

for an aircraft 1. The control system includes at least one aerofoil in the form of a

directional control surface 2. In Figure 1, the control surface 2 is a vertically oriented

aerofoil aligned with the longitudinal axis of the aircraft fuselage 4. The control surface 2, in this case taking the place of a conventional fm and rudder assembly, is supported by

means of bearings 5 for rotation about a vertical axis 6. The rotational axis 6 is generally

perpendicular to the longitudinal axis of the aircraft.

Resilient bias means 7, represented diagrammatically as spring box 8, urges the control surface 2 toward a central rest position which is aligned with the longitudinal

axis of the aircraft, as shown. The spring bias mechanism is, however, adapted to permit

a limited degree of rotational movement of the control surface 2 away from the central

rest position, in response to unbalanced aerodynamic pressure loadings, as will be

described in more detail below. Although the bias mechanism is represented diagrammatically as an arrangement of springs, it will be appreciated that alternative

biasing means such as hydraulic, pneumatic, or electric actuators, could also be used in

conjunction with computerised controllers to produce the desired aerodynamic or system

response.

It should be noted that the effective centre of pressure of the control surface,

indicated as point P, is spaced rearwardly from the axis of rotation 6. In this

configuration, unbalanced reaction pressures acting on the control surface tend

automatically to effect a corresponding rotation against a restoring force provided by the

spring bias mechanism.

Turning to Figures 2 and 3, it will be seen that the inventive principle has been

applied on the primary wing formations as well as the tail fin of an aircraft. In this case,

control surfaces 10 according to the invention are formed integrally with the primary

lifting surfaces or wings of the aircraft. Accordingly, they are disposed substantially further forward on the fuselage than the vertical tail control surface 2 of Figure 1. The

horizontally oriented wing formations are rotatably supported by means of respective

static transverse support shafts 12, which are located by support shaft bearings 14. Each

static shaft 12, which is fixed to the fuselage, rotatably supports a respective axle shaft

12 A, which is fixed with respect to the corresponding wing formation. In this way, the

wing formations are able to rotate independently about a common horizontally oriented

transverse pivot axis 13. This mounting arrangement enables different sizes or shapes of

wings to be interchanged, to optimise performance having regard to variable factors such

as payload, intended cruising speed, weather conditions, desired comfort levels, and the

like. It is conceivable that an exchange of wings or other control surfaces could be performed as a routine operation, possibly even during a refuelling stop, even acting as a fuel cell which is already fully laden with fuel.

Resilient bias means, in this case shown diagrammatically in the form of spring

mechanisms 15, independently urge the respective wing formations toward a central

generally horizontal rest position, in a common plane, while accommodating a limited

degree of rotational displacement in response to unbalanced pressure loadings. Once

again, the centre of pressure P of each of the main wing surfaces is disposed rearwardly

of the axis of rotation 13.

It will be noted that the vertically oriented control surface 2 of the aircraft replaces

a conventional fin and rudder assembly. However, the aircraft includes a pair of

forwardly disposed vertically oriented canard rudders 21 (see Figure 3), disposed respectively above and below the fuselage. Similarly, the conventional tail plane and

elevator assembly as shown in Figure 1 is replaced by a pair of forwardly disposed,

relatively smaller, canard wings 25. Each canard wing is independently supported on a

shaft 27 for rotation about an horizontal axis 28 by means of respective separate motor

and bearing assemblies 29. The canard rudders are mounted in the same manner.

In this way, it will be appreciated that rotation of the aircraft, in flight, around the

pitch axis is regulated primarily by the canard wings 25. Rotation about the yaw axis is

controlled primarily by the dual canard rudders 21. Rotation about the roll axis is regulated by rotation of the primary or main wing formations, trailing in opposite

directions. Alternatively or additionally, ailerons may be fitted to the wing formations

for this purpose if required. In the embodiment of Figures 2 to 5, the wing mounted engines are shown as fixed

with respect to the fuselage, such that the engine thrust line remains substantially parallel to the longitudinal axis of the aircraft independent of the rotational displacement of the

main wing formations. Alternatively, however, in other embodiments the same engines affixed to the fuselage or extensions thereof, could be designed to rotate in concert with

the primary wing formations, such that the thrust lines remain substantially parallel to

the respective control surfaces. The same engine rotation principle could be applied to

other control surfaces embodying the invention, for example aligning with the tail fin.

Turning now to consider the operation of the control surfaces during various

manoeuvres, Figure 6 is a diagrammatic plan view showing an aircraft according to the invention flying straight ahead, at normal cruising speed, in a steady state of dynamic

equilibrium. To initiate a turn to the right, the forward canard rudders 21 are deflected to

the left shown in Figure 7. During this initial phase of the turning manoeuvre, as the aircraft rotates about the vertical yaw axis, the incident air stream causes the vertical

control surface 2 to be rotationally displaced in an anticlockwise direction (when

viewing the drawings) relative to the longitudinal axis of the fuselage, as shown in

Figures 8 and 9. The spring bias force tending to restore the control surface 2 to its

central rest position is calibrated to maintain an optimum angle of attack with respect to

the incident air stream, thereby maintaining the efficiency of the control surface and

delaying the onset of a stall condition. Once the new direction of the aircraft begins to

be established, the fin 2 also begins to return to the central rest position as shown in

Figure 9. Finally, Figure 10 shows the aircraft stabilised after the turn, with the canard

rudders straight and the control surface 2 realigned in its rest position.

SUBSΉTUTE SHEET (Rule 26) (RO/AU) Figures 11 to 17 show a similar sequence in side elevation, indicating the

movement of the wing formations in response to rotation of the aircraft about the pitch

axis, during initial take-off, steady climb, and levelling off manoeuvres. Initially, Figure

11 shows the aircraft with the wing formations 10 aligned with the longitudinal axis of

the fuselage, at the commencement of taking-off. With take-off speed attained, as shown in Figure 12, the forward canard wings 25 are rotated and hence deflected downwardly

to lift the nose. It will be noted, however, that as the nose begins to lift, the differential

pressure on the wing surfaces causes the wings to rotate in an anti-clockwise direction

(when viewing the drawings) relative to the fuselage, against the bias force provided by

the internal spring mechanisms 15. Once again, the bias force is calibrated to ensure that

the wing formations are presented to the incident air stream at an optimum angle of attack. This maximises lifting efficiency and at the same time, helps to prevent the onset

of a stall condition. As the angle of climb begins to stabilise, as shown in Figures 13 and

14, the pressure differential diminishes and the wings progressively return toward the

normal rest position, close to the neutral orientation before take-off. As the desired

cruising altitude is approached, as shown in Figures 15 and 16, the canard wings are

deflected upwardly, to cause the aircraft to begin to roll out of the climb, into level

flight. During this phase, the wings are temporarily deflected in the opposite direction to

possibly cause an argumentation of lift. Finally, as stable, level flight is established, the wings once again assume the neutral or rest position, generally parallel with the axis of

the fuselage, as shown in Figure 17.

Although the various control surfaces are shown as incorporating a fixed

traditional section and a moveable section, this is essentially for illustrative purposes only. It is envisaged that a fixed section is not required and that in practice this would be integrated with the moveable section, as shown in the more detailed drawings.

Figures 18 to 20 show a third embodiment of the invention, adapted for large scale

passenger transportation, wherein corresponding features are denoted by corresponding

reference numerals. The principles of operation of the various control surfaces are

essentially the same as those described above. In this case, however, it will be seen that

the aircraft incorporates respective upper and lower pairs of forward canard rudders 21

and respective upper and lower pairs of rearward primary, vertical control surfaces 2.

This multiple arrangement including additional control surface placements enables the

vertical extent of these control surfaces to be reduced, while maintaining adequate surface area for efficient and effective directional control. Importantly, these four pairs

of vertically oriented control surfaces define respective control quadrants, within which the individual control surfaces are independently moveable in a complementary manner,

to optimise manoeuvrability and turning efficiency. As with the embodiments

previously described, the passive control surfaces 2 may be governed by spring bias

means, or alternatively by computer controlled hydraulics or other suitable means.

Additional supplementary canard wings, rudders and other control surfaces may also be provided to optimise these particular performance characteristics.

In another embodiment, the primary wing formations 10, the canard wings 25, and

the rearward control surfaces 2 each incorporate a central aperture or opening 30. These

apertures are believed to minimise turbulence, reduce response time to control inputs,

and generally enhance manoeuvrability. It will be appreciated that the ratio of the area

of the aperture to the overall area of the control surface can be varied so as to optimise

SUBSTΓΠJΓΈ SHEET (Rule 26) (RO/AU) particular performance criteria. For example, in passenger aircraft it may be desirable to

optimise lift at the expense of manoeuvrability, whereas in military aircraft, it may be

preferable to optimise speed and manoeuvrability.

Figure 19 shows the aircraft of Figure 18 with undercarriage 32 deployed, ready

for landing. In landing mode, the lower vertical control surfaces (both forward and

rearward) are retracted upwardly toward the fuselage, to provide adequate ground

clearance for landing. Once the aircraft is airborne, the landing gear is retracted in

conventional manner, and the lower control surfaces are operatively redeployed to the optimum orientation for high speed flight.

Figure 20 is an enlarged detail showing one possible mechanism for retraction of

the lower vertically oriented control surfaces. In this case, the control surface 21 is

mounted on a shaft 35 adapted for pivotal movement about transverse axis 36. A tie rod

37 (shown in several positions indicating its locus of movement) connects the support shaft 35 to a driven gear 40 by means of a crank pin. The driven gear 40 is engaged and

activated by a driving pinion 41, connected to a suitable drive motor. As the driven gear

is rotated in a clockwise direction (when viewing the drawing) the tie rod moves

upwardly and in turn draws the control surface 21 upwardly toward the fuselage, as

indicated by arrow 45. At the same time, the control surface is pivoted through 90° so as to lie against the underside of the fuselage, as shown. Deployment of the control surface

is essentially the reverse of this procedure. A variation of this embodiment is shown in

Figure 20 A. In an alternative embodiment, as shown in Figure 21, the retraction

mechanism makes use of linear hydraulic actuators 50 in place of the rotary crank

mechanism. It is believed that the control surfaces embodying the present invention operate

during turning manoeuvres to increase manoeuvrability, augment lift, πύnirnise drag,

and optimise turning efficiency. This in turn improves fuel consumption, reduces power

consumption, preserves air speed, and conserves the momentum of the aircraft. Perhaps most significantly, the invention delays the onset of stall conditions which could

otherwise result in a dramatic reduction in the efficiency and effectiveness of the control

surface, as well as associated lifting surfaces. In these respects, the invention represents

a functional and commercially significant improvement over the prior art.

Although the invention has been described with reference to specific examples, it

will be appreciated by those skilled in the art that the invention may be embodied in many other forms. In particular, it should be noted that the principle may be applied to

any control or lifting surface of an aircraft including fins, rudders, tail planes, elevators,

wings, ailerons and flaps, whether disposed forwardly or rearwardly, and whether

oriented vertically, horizontally or at any intermediate angle. It should also be

appreciated that the desired biasing forces need not be provided by resilient means in the

conventional sense, but could be regulated by more complex hydraulic, pneumatic,

electrical, or mechanical actuators, with control algorithms tailored to the desired

performance characteristics of the aircraft.

Claims

CLAIMS:-

1. A control system for an aircraft, said control system including:

an aerofoil surface;

bias means operable to urge the aerofoil surface toward a central rest position

aerodynamic reaction pressures acting on the aerofoil surface tend automatically to effect a corresponding rotation against a restoring force provided by the bias means

2. A control system for an aircraft according to claim 1 wherein the aerofoil surface takes the form of a directional control surface.

3. A control system for an aircraft according to claim 2 wherein the control surface is

oriented vertically, horizontally or at any intermediate angle, and is configured to operate

during manoeuvres involving roll, pitch or yaw.

4. A control system for an aircraft according to claim 2 or 3 wherein the control

surface includes a rudder.

5. A control system for an aircraft according to claim 4 wherein the rudder is

supported by a fin.

6. A control system for an aircraft according to claim 2 or 3 wherein the control

surface includes a stabiliser or tail plane.

7. A control system for an aircraft according to claim 6 wherein the stabiliser or tail plane includes elevators.

8. A control system for an aircraft according to any one of claims 2 to 7 wherein the control surface is integrated into a tail assembly toward the rear of the aircraft, and is

complemented by conventional control surfaces in the form of canard wings and canard

rudders, disposed toward the front of the aircraft and adapted to initiate directional

changes.

9. A control system for an aircraft according to any one of claims 2 to 8 including

respective upper and lower pairs of forward canard rudders and respective upper and

lower pairs of rearward vertical control surfaces.

10. A control system for an aircraft according to claim 9 wherein the respective upper and lower pairs of forward canard rudders and the respective upper and lower pairs of

rearward vertical control surfaces are independently powered, while being controllable to

operate in a complementary manner, in conjunction with or independent of the canard

wings and primary wing surfaces.

11. A control system for an aircraft according to claim 10 wherein at least one of the

lower vertical control surfaces includes retraction means adapted for selectable retraction

upwardly toward the fuselage, to provide adequate ground clearance for landing.

12. A control system for an aircraft according to claim 11 wherein said retraction means include a shaft disposed pivotably to support the control surface, the shaft being supported for movement about a transverse axis, a tie rod hingedly connected at one end

to the shaft and hingedly connected at the other end to an actuating crank integral with a

driven gear.

13. A control system for an aircraft according to any one of the preceding claims

wherein the control surface incorporates a central aperture or opening.

14. A control system for an aircraft according to claim 1 wherein the aerofoil surface takes the form of a primary wing surface.

15. A control system for an aircraft according to any one of the preceding claims wherein the mounting means include a bearing.

16. A control system for an aircraft according to any one of the preceding claims

wherein the bias means include a spring box adapted to permit a limited degree of

rotational movement away from the central rest position in response to unbalanced aerodynamic pressure loadings.

17. A control system for an aircraft according to any one of claims 14 to 16 wherein

each wing surface is disposed to rotate independently about common horizontally

oriented transverse pivot axis.

18. A control system for an aircraft according to any one of claims 14 to 17 wherein

the primary wing surfaces each include a central aperture.

19. A control system for an aircraft according to any one of claims 14 to 18 wherein

wing mounted engines are moveable with respect to the wing, such that the engine thrust

line remains substantially parallel to the longitudinal axis of the aircraft independent of

the rotational displacement of the main wing formations.

20. A control system for an aircraft according to any one of claims 14 to 18 wherein engines are adapted to rotate in concert with the primary wing formations, such that the

thrust lines remain parallel to the respective control surfaces.

21. A control system for an aircraft substantially as herein described with reference to any one of the embodiments of the invention illustrated in the accompanying drawings

and/or examples.

SUBSΉTUTE SHEET (Rule 26) (RO/AU)