Vehicle

The present invention provides a vehicle comprising: a rotor and a stator; at least one planar control surface coupled to the rotor, wherein the rotor is configured to rotate relative to the stator such that, in use, the at least one planar control surface moves from a first position to a second position, and wherein in the first position the planar control surface is controllable to affect substantially only the pitch of the vehicle and in the second position the planar control surface is controllable to affect substantially both of the pitch and yaw of the vehicle, or substantially only the yaw, or in the first position the planar control surface is controllable to affect substantially only the yaw of the vehicle and in the second position the planar control surface is controllable to affect substantially both of the pitch and yaw of the vehicle, or substantially only the pitch of the vehicle. The present invention also provides a method of controlling a vehicle.

RELATED APPLICATIONS

This application is a national phase application filed under 35 USC § 371 of PCT Application No. PCT/GB2018/052283 with an International filing date of Aug. 10, 2018, which claims priority of GB Patent Application 1713062.6 filed on Aug. 15, 2017 and EP Patent Application 17187056.1 filed on Aug. 21, 2017. Each of these applications is herein incorporated by reference in its entirety for all purposes.

FIELD

This specification relates generally to a vehicle and a method of controlling a vehicle. More specifically, the present invention relates to a vehicle having an adaptable control surface.

BACKGROUND

In a typical aircraft configuration, the empennage (i.e. rear section of the fuselage) includes a horizontal tail plane for controlling not only vertical stability but the pitch of the aircraft. The empennage also includes one or two vertical tail fins for controlling lateral stability and yaw. Either portions (i.e. elevators and rudders) of these control surfaces or these control surfaces in their entirety are rotatable about the lateral axis of the respective control surface.

While known control surfaces allow directional control of the aircraft, they do not provide a means to control the radar return signature and performance envelope of the aircraft. The same problem can be found on other vehicles, such as submarines, missiles, and torpedoes. The present invention seeks to address at least this problem across a range of platforms.

SUMMARY

According to a first aspect of the present invention, there is provided a vehicle comprising:

a rotor and a stator;

at least one planar control surface coupled to the rotor,

wherein the rotor is configured to rotate relative to the stator such that, in use, the at least one planar control surface moves from a first position to a second position, and wherein:

in the first position the planar control surface is controllable to affect substantially only the pitch of the vehicle and in the second position the planar control surface is controllable to affect substantially both of the pitch and yaw of the vehicle, or substantially only the yaw, or

in the first position the planar control surface is controllable to affect substantially only the yaw of the vehicle and in the second position the planar control surface is controllable to affect substantially both of the pitch and yaw of the vehicle, or substantially only the pitch of the vehicle.

As such, the vehicle is adaptable to change the control dynamics and/or profile of the vehicle during operation of the vehicle. The actuating system for adapting the vehicle can be lightweight and uncomplicated relative to the number of actuators required to adapt the control surface by the same degrees of freedom by conventional means. Due to the vehicle not requiring a pivot point in the root of the planar control surface, the radar cross section of the vehicle tends to be reduced.

The longitudinal axis of the vehicle and central axis of the stator may be arranged in parallel.

In the first position, the planar control surface may be substantially co-planar with the local horizontal plane of the vehicle and in the second position the planar control surface may be arranged in a plane at 45 degrees to the local horizontal plane.

The vehicle may comprise an actuating system, comprising:

at least one actuator coupled to the rotor and to the stator,wherein the control surface is coupled to the rotor such that the control surface extends perpendicularly from the periphery of the rotor wherein, in use, when a first actuator is in a first configuration the rotor is rotated such that the planar control surface is arranged in the first position, and when the first actuator is in a second configuration the rotor is rotated such that the planar control surface rotates about the central axis of the rotor to be arranged in the second position.

The at least one actuator may comprise a piston and the at least one actuator may be pivotably connected to the stator and to the rotor.

The at least one actuator may comprise a plurality of actuators arranged in pairs. The plurality of actuators may comprise only three pairs of actuators.

The plurality of actuators may define an elongate space having the stator and rotor disposed at the longitudinal ends of the elongate space, and wherein the stator and rotor are annular. An engine may be disposed in the elongate space. The at least one actuator may be configured to incline the central axis of the rotor with respect to the central axis of the stator in order to vector the thrust of the engine.

In use, when one actuator expands another actuator may retract to cause rotation of the rotor. Particularly, within the pairs of actuator, as one actuator extends, the other can retract to cause rotation of the rotor.

The planar control surface may comprise a portion for controlling the direction of the vehicle, the portion being moveable independently of the planar control surface.

The vehicle may comprise a second rotor and a second stator, the second rotor being coupled to a second planar control surface, wherein the second rotor may be configured to rotate relative to the second stator such that, in use, the second planar control surface moves from the first position to the second position, and wherein the rotor may be arranged adjacent and parallel to the second rotor such that the planar control surface and the second planar control surface are arranged on opposite sides of the vehicle with respect to each other. The vehicle may be an aircraft.

The planar control surface may be a tail plane. Alternatively, the planar control surface may be a canard.

The rotor may be a tail cone.

According to a second aspect of the present invention, there is provided a method of controlling a vehicle comprising rotating a rotor relative to a stator, a planar control surface being coupled to the rotor, such that the planar control surface moves from a first position to a second position, wherein:

in the first position the planar control surface is controllable to affect substantially only the pitch of the vehicle and in the second position the planar control surface is controllable to affect substantially both of the pitch and yaw of the vehicle, or substantially only the yaw, or

in the first position the planar control surface is controllable to affect substantially only the yaw of the vehicle and in the second position the planar control surface is controllable to affect both of the pitch and yaw of the vehicle, or substantially only the pitch.

The method may comprise rotating the rotor +/−22.5 degrees relative to the stator to effect a 45 degree rotation of the planar control surface relative to a fixed plane.

The method may comprise changing a flow path through the vehicle by controlling at least one actuator to incline the central axis of the rotor relative to the central axis of the stator.

The method may comprise controlling a first one of a plurality of actuators to expand while controlling a second one of the plurality of actuators to retract to cause rotation of the rotor.

DETAILED DESCRIPTION

Embodiments herein relate generally to an actuating system for adapting a control surface of a vehicle to suit the mode of the vehicle. For example, the present invention may relate to a tail structure for an aircraft having control surfaces able to reposition from the local horizontal plane toward the local vertical plane (or a midpoint between the horizontal and vertical planes) depending on whether the pilot requires high speed, stealth or high levels of lateral control. Each of these conditions are variables, among others, that define a mode, such as supersonic cruising mode, enemy airspace infiltration mode, or dogfighting mode.

An aircraft100on which the present invention is implemented will now be described with reference toFIG. 1. The aircraft100shown here is a fighter jet, but equally the invention can also be implemented on unmanned aerial vehicles (UAVs), transport, intelligence-gathering, airborne early warning and control, maritime patrol or civilian aircraft. In other embodiments, instead of an aircraft100, the invention is implemented on other vehicles such as a weapon (for example a cruise missile, ICBM, torpedo, air-to-air missile, air-to-ground missile or guided rocket), a waterborne or submersible vessel (for example, a hovercraft, submarine, or ship), or land vehicles such as tanks for controlling the direction of the main gun or turret, or direction of an antenna.

It should be noted that the local axis of the aircraft100moves with the airframe, and therefore the axes, directions and control inputs defined herein below are relative to the local axis system of the aircraft100rather than the global frame of reference.

The aircraft100shown inFIG. 1has a nose section20, wings10a,10b, fuselage15(i.e. main body), engine exhaust nozzles40a,40band tail surfaces30a,30b. Although two engine exhaust nozzles40a,40bare shown, in some embodiments the aircraft100has only one engine and associated nozzle40a,40b. Furthermore, in some embodiments, the aircraft100further includes canards disposed on the fuselage15between the wings10a,10band the nose20. The tail surfaces30a,30band canards are used to control the pitch of the aircraft100and stabilise the aircraft100in the local horizontal plane (otherwise known as the common fuselage horizontal datum) when the lateral axes2a,2bof the tail surfaces30a,30bor canards are in the local horizontal plane. To cause the aircraft100to perform a lateral roll, the tail surfaces30a,30bare independently rotatable about their respective lateral axes2a,2b. In other words, the tail surfaces30a,30band canards are control surfaces. In other embodiments, the tail surfaces30a,30bmay not be rotatable about their lateral axes2a,2band could comprise parts rotatable relative to the remaining parts of the tail surfaces30a,30bto control the pitch of the aircraft100. In other words, in some embodiments the tail surfaces30a,30bcomprise elevators.

Three longitudinal axes1a,1b,1cof the aircraft100are shown drawn through the aircraft100. The fuselage15centre line1a, or primary longitudinal axis, passes through the aircraft100from nose20to tail, and equally bisects the aircraft100. The port engine centre line1bpasses along the length of the port engine. The starboard engine centre line1cpasses along the length of the starboard engine. The lateral axis2a,2bof each tail surface30a,30bis shown drawn through the tail surfaces30a,30bfrom their outboard edge to the inboard edge. The lateral axes2a,2bmay be perpendicular to the primary longitudinal axis1aof the aircraft100.

The rear section of the aircraft100having the tail surfaces30a,30bis known as an empennage (or tail assembly).

An actuating device200for use in controlling the tail surfaces30a,30bwill now be described with reference toFIG. 2. The same actuating device200is adaptable for use in controlling other control surfaces, such as canards and the main wings10a,10b. The actuating device200is capable of achieving six degrees of freedom, namely X, Y, Z translations and X, Y, Z rotations. While the actuating device200is physically capable of achieving these diverse degrees of freedom, in certain embodiments these degrees of freedom are limited by software such that the central point of a rotatable member220adoes not move out of alignment with the central point of a fixed member220b. In other words, the rotatable member220ais only rotated relative to the fixed member220b, and not translated.

The actuating device200includes a plurality of linear actuators210. As will be explained later, for optimum translations and rotations required for moving control surfaces in the context of the present invention, the actuating device is provided with six linear actuators210.

In plan view, the six actuator210centrelines form a hexagonal figure. Where the actuators210are all at equal length (i.e. mid stroke) and if the effective rotatable and fixed member diameters were the same, this is a regular hexagon. As alternate actuators210expand and contract to maintain the centre locus while rotating the rotatable member220ato a limit value (in this case +/−22.5 degrees), the mechanical advantage to power the rotation changes, and the mid-point of the longer actuators moves towards the engine casing disposed such that it passes through the figure defined by the actuators210(shown inFIGS. 4aand 4b). These variables, along with the available stroke and the distance between the rotatable and fixed members give a practical limit to the motion. Thus each pair of actuators210occupies a 120 degree segment of the circle to achieve a +/−22.5 degrees rotary motion (where the zero degree point is where all actuators210are equally extended). If a greater number of actuators was used (for example, eight) each pair would occupy a 90 degree segment of the circle and could no longer extend to achieve this rotation.

In other words, the optimal number of actuators210is six. If more were used, the necessary 45 degree translation of the tail surface30a,30bcould not be achieved. Meanwhile, if fewer actuators210were used, the rotatable member220awould lose a structural constraint (i.e. become unbounded and lose rigidity in at least one axis).

The actuators210include a coupling device to pivotably connect the actuators210to the rotatable member220aand fixed member220b. In certain embodiments, the actuators210have a coupling device at each end. The coupling devices have spherical bearings. In the embodiment shown here, the coupling devices are in the form of universal joints which allow adequate freedom of rotation to achieve the required mechanism motions. In other embodiments, the actuators210include ball joints at both ends for coupling to sockets. In other embodiments, the actuators210include dual-axis hinges. In further embodiments, any of the previously described coupling devices may be combined in the same actuating device200. The actuators210include pistons between the universal joints for driving the universal joints apart or bringing them together. The pistons are hydraulically actuated. In other embodiments, the pistons are pneumatically actuated. In other embodiments, the pistons are driven by an electric motor.

The universal joints at one end of the actuators210are connected to the rotatable member220a. The universal joints at the opposite end of the actuators210are connected to the fixed member220b. While the fixed member220bis fixed relative to the body of the aircraft100, the rotatable member220ais able to tilt and rotate according to which actuators210are extended and which are retracted. This will be explained in more detail later with reference toFIGS. 4aand 4b. In the embodiments shown, the rotatable and fixed members220a,220bare ring-shaped, or annuli.

The actuators210are arranged in pairs, such that for a first pair of actuators210the universal joints at one end are disposed proximate to each other when connected to the rotatable member220a. Meanwhile, the universal joints at the opposite end of the same pair of actuators210are spaced apart from each other and disposed proximate to the universal joints of another pair of actuators210.

While linear actuators210in the forms of pistons have been shown and described herein, it would be appreciated that the rotatable member220aand fixed member220bmay be coupled by other driving means, such as a rack and pinion, worm drive or direct motor. However, it would require a complex and heavy arrangement of such driving means to provide the same degrees of freedom provided by the actuating device200shown inFIG. 2. Moreover, the embodiment having a plurality of pistons provides a level of redundancy not found when the other driving means are used.

While the actuating device200according to embodiments described above, comprising pistons210and ring-shaped rotatable and fixed members220a,220b, is limited by geometry to cause a maximum 60 degree deflection of the control surface relative to the local horizontal plane, it would be appreciated that other actuating devices could be used to provide an adaptable control surface capable of 90 degrees of dihedral and/or anhedral deflection. For instance the stator (fixed member220b) and rotor (rotatable member220a) of the present invention could be integrated with the stator and rotor of an electric motor.

FIG. 2shows the central axis3of the rotatable member220aand the central axis4of the fixed member220b. These central axes3,4are parallel and collinear when the rotatable member220ais not tilted.

An empennage featuring the present invention will now be described with reference toFIGS. 3aand 3b.FIG. 3ashows tail surfaces30a,30barranged in horizontal configuration (or 0 degrees dihedral). Here, the term horizontal is used in relation to the local reference frame of the aircraft100, and is therefore not necessarily parallel to the horizon. Horizontal means the tail surfaces30a,30bare substantially co-planar with each other and the aircraft's100horizontal plane (i.e. substantially co-planar with the wings10a,10b). In other words, the lateral axes2a,2bof each tail surface30a,30bare not inclined relative to each other. When the tail surfaces30a,30bare horizontal, the distance between the outboard edges of the tail surfaces is at its greatest. Rotating the tail surfaces30a,30babout their lateral axes2a,2bwill cause the aircraft's100nose20to pitch up or down (in the aircraft's local frame of reference). This configuration is particularly useful when the aircraft100is travelling at supersonic speeds and the pilot has no need to change the heading of the aircraft100. Moreover, this configuration is useful when the aircraft100is required to present a low radar cross section.

Meanwhile,FIG. 3bshows the tail surfaces30a,30bin a configuration in which they are arranged at 45 degrees to the local horizontal plane as defined above. The lateral axes2a,2bof the tail surfaces30a,30bform an angle of 90 degrees between each other. In other words, here the tail surfaces30a,30bmake an angle of 45 degrees with a plane generally parallel with the wings10a,10b. In this configuration, the tail surfaces30a,30bcan be used to control yaw of the aircraft100and/or pitch of the aircraft100.

In further embodiments, the actuating devices200are configured to allow the tail surfaces30a,30bto rotate about the central axes3of the rotatable members220ato be perpendicular to the local horizontal plane as explained above. In other words, in this configuration, the lateral axes2a,2bof the tail surfaces30a,30bdo not intersect each other.

FIGS. 3aand 3bshow a cutaway part on one side of the empennage. The internal structure of the corresponding part of the side not shown in cutaway is identical to the structure described below for the cutaway part.

The empennage according to this embodiment includes two engines50laterally displaced either side of the central axis1a. In other embodiments, there may only be a single engine, or an engine may be disposed elsewhere on the aircraft100and not form part of the present invention. While jet engines are shown here, the engines50may also be ram jet engines, rocket engines or hybrid engines.

An actuating device200as described with reference toFIG. 2is disposed around each engine50. Here, a jet engine is disposed such that it passes through the annuli or the rotatable member220aand fixed member220b. The casing of the engine50and its exhaust have a circular cross section.

The tail surfaces30a,30bare coupled radially to the outside edge of respective rotatable members220aand extend outward therefrom. Therefore, when the actuators210extend or retract in a predetermined way, the tail surfaces30a,30bare caused to rotate about the longitudinal axis3of the rotatable member220ato be horizontal (i.e. transverse to the fuselage15) as inFIG. 3aor make an angle with the local horizontal plane as shown inFIG. 3b.

The outside edge of the fixed member220bis fixed to the inside wall of the fuselage15such that its longitudinal axis4is parallel to the longitudinal axis1aof the fuselage15. When all actuators210are at their mid-stroke neutral position, the central axis3of the rotatable member220ais arranged collinearly with the central axis4of the fixed member220b. The outside periphery of the rotatable member220aforms part of the outside surface of the aircraft100. The rotatable member220acomprises a baffle for coupling it to the fuselage15to prevent gaps forming in the outside surface of the aircraft100when the rotatable ring220apivots up, down, left or right with respect to the fuselage15.

While dihedral deflection of the tail surfaces30a,30bis shown inFIG. 3b, the present invention can also be used to deflect the tail surfaces30a,30banhedrally (i.e. downwards relative to the local horizontal plane).

The engine exhaust nozzles40a,40bare coupled axially to the periphery of respective rotatable members220a. Therefore, when the central axis3of the rotatable member220ais caused to be inclined relative to the central axis4of the fixed member220b, such that the central axis3of the rotatable member220amakes an angle with that of the fixed member220b, the engine exhaust nozzle40ais pivoted such that its longitudinal axis1bmakes the same angle with the central axis4of the fixed member220b. For example, the engine exhaust nozzle40acan be controlled to point up or down relative to the fuselage15. Therefore, a flow path of the engine exhaust through the empennage is adjusted. This is known as thrust vectoring.

In these embodiments, the present invention alleviates the need for an aircraft to have both a horizontal tail plane and a tail fin, as it allows the empennage to be adaptable to perform the function of both features. Therefore, radar cross section of the aircraft100and drag tend to be reduced when necessary, and the performance envelope of the aircraft100tends to be dynamically controllable. Moreover, the aircraft100is generally lighter than an aircraft having both horizontal and vertical tail surfaces.

In other embodiments, instead of or in addition to the tail surfaces30a,30b, control surfaces such as the main wings10a,10b, weapons pylons or front canards of the aircraft100can be controlled in a similar manner to that described above.

The present invention will now be described in more detail with respect toFIGS. 4aand 4b. These Figures show cross sections through the longitudinal axis of the port side of the empennage shown inFIGS. 3aand 3b. The view is taken from the nose20of the aircraft100through to the tail.

As shown, the engine casing surrounding the engine50is cylindrical. When actuators210surrounding the engine casing are at maximum extension, the midpoints of each fully extended actuator210are accommodated by a recess in the casing. A clearance should be maintained between moving and static parts, with enough allowance for emergency extension limits of the actuators210without contact with equipment disposed between the plurality of actuators210.

The actuators210are arranged in pairs. First actuator210aand second actuator210bform a first pair of actuators. Third actuator210cand fourth actuator210dform a second pair of actuators. Fifth actuator210eand sixth actuator210fform a third pair of actuators. As shown inFIGS. 4aand 4b, the area between the engine casing and the skin of the aircraft100is limited. To include more actuators210would reduce the angular range in which each actuator operates and therefore prevent the tail surface30brotating to 45 degrees above the local horizontal plane. Meanwhile, fewer actuators210would reduce the degrees of freedom of the rotatable member220a.

In certain embodiments, the number of actuators210is an even number. The number of actuators210could therefore be reduced to four. In either a symmetrical or asymmetrical array, assuming pin-joint behaviour, these would then act like two joined 4-bar links, and could not be constrained in all degrees of freedom without additional constraint at the rotational centre, which cannot be achieved when an engine50is in the way.

The rotatable member220ais of a smaller diameter than the fixed member220band is axially yet collinearly spaced apart from the fixed member220b. Here, the rotatable member220ais shown being further aft in the empennage than the fixed member220b. In some embodiments, the rotatable member220ais disposed forward of the fixed member220bwith respect to the nose20of the aircraft100.

The tail surface30bis coupled to the rotatable member220aby any suitable means. For example, the tail surface30bmay be riveted, bolted or welded to the rotatable member220a. In other embodiments, the rotatable member220aand tail surface30bare formed as a single integrated unit. In further embodiments, a tail cone (or fairing) is coupled axially to the rotatable member220a, and the tail surface30bis instead coupled to the tail cone in a similar manner to as previously described. Therefore, when the rotatable member220arotates, the tail surface30bpivots about the longitudinal axis of the rotatable member220a.

In order to fully rotate the tail surface30bto 45 degrees with respect to the local horizontal plane, which is the plane parallel to the horizon when the aircraft100is flying straight and level (i.e. in the same plane as the primary longitudinal axis1aof the aircraft100), the second actuator210b, fourth actuator210dand sixth actuator210fextend in unison, while the first actuator210a, third actuator210cand fifth actuator210eretract in unison. The resulting configuration is shown inFIG. 4b. In order to bring the tail surface30bback to the local horizontal plane, the second actuator210b, fourth actuator210dand sixth actuator210fretract in unison, while the first actuator210a, third actuator210cand fifth actuator210eextend in unison. The resulting configuration is shown inFIG. 4a. As shown inFIGS. 6 and 7, intermediate dihedral inclinations of the tail surface30bcan be achieved through predetermined combinations of actuator extensions. The desired pitch of the tail surface30bcan be maintained while the tail surface30bmoves dihedrally.

If one of the actuators210malfunctions (e.g. jams), it is possible for some degree of control of the tail surface30bto be maintained through actuation of the remaining actuators210. The end of the failed actuator coupled to the rotatable member220ais swung further aft of the vehicle100by its paired actuator extending. The remaining two pairs of actuators translate in unison to allow this motion, which moves the rotatable member220afurther aft and out of standard alignment with the fixed member220b. This new aft position can be used as a basis for reversionary motions of the remaining five actuators. Although full freedom of movement will not be achieved, enough control can be maintained until the aircraft100is brought to a stop (or, where an aircraft, lands).

By extending each of the first actuator210aand second actuator210bby a greater amount than that of the third through sixth actuators210c-f, the rotatable member220acan be pivoted such that its central longitudinal axis3of the rotatable member220amakes an angle with the longitudinal axis4of the fixed member220b. In the embodiment shown here, where an engine50extends through the fixed member220band an engine exhaust nozzle40bis coupled to the rotatable member220a, pivoting the central longitudinal axis3of the rotatable member220aaway from the central longitudinal axis4of the fixed member220bcauses the thrust generated by the engine50to be vectored thereby providing the aircraft100with increased manoeuvrability. This is useful, for example, in aircraft and water-jet powered vehicles100such as jet skis and torpedoes.

In further embodiments, sensors such as imaging systems and radar warning receivers are disposed on the distal tips of the tail surfaces30a,30b. Therefore, by controlling the tail surfaces30a,30bto move out of the aircraft's100primary horizontal plane (i.e. anhedrally or dihedrally), the obscuration of the field of regard of the sensors by the aircraft's100fuselage15or wings10a,10bcan be reduced.

In the embodiment shown here, the same features shown inFIG. 4aare mirrored on the starboard side of the aircraft100in order to provide the aircraft100with two adaptable tail surfaces30a,30b.

FIGS. 5aand 5bshow an alternative arrangement for causing movement of a tail surface30b. Here, the actuators210are coupled directly to the inside surface of a tail fairing (or tail cone)60. The opposite end of each actuator210is coupled to the inside surface of the fuselage15, for example a fuselage, of the aircraft100. A hinge65couples the tail fairing60to the fuselage15. The hinge65is able to compress and expand to prevent gaps forming between the tail fairing60and the fuselage15when the tail fairing60is moved. The hinge65acts as a baffle. In effect, the tail fairing60is an elongated rotatable member220aas previously described with reference toFIGS. 3ato 4b. Meanwhile, in effect the fuselage15of the aircraft100acts as a fixed member200bas previously described with reference toFIGS. 3ato4b.

A planar tail surface30bis coupled to the tail fairing60. The planar tail surface30bmay be coupled to the tail fairing60by any suitable means, such as bolting, riveting, welding, or the tail surface30band tail fairing60may be integrally formed.

In further embodiments, an auxiliary power unit (APU) is disposed inside the tail fairing. Here, the distal end of the tail fairing60comprises an engine exhaust nozzle. In some embodiments, sensors, such as magnetic anomaly detectors or synthetic aperture radars, are disposed on the tail fairing60of the aircraft100. Therefore, the direction in which these sensors point can be adjusted by pivoting the tail fairing60using the appropriate actuators210.

Similarly to as shown inFIGS. 2aand 2b, two actuating devices200may be disposed side-by-side (or in parallel) to provide two adaptable tail surfaces30a,30b. These two tail surfaces30a,30bcan then rotate upwards, i.e. dihedrally, to provide an aircraft100with a twin tail fin. Such rotation would tend to happen in unison so that each tail fin was inclined by an equivalent angle to the horizontal plane.

In other embodiments again, two actuating devices200are coupled axially in the same aircraft100. In other words, the rotatable member220aof one actuating device200is disposed proximate to the distal end of another actuating device200, or concentrically with the other actuating device. A control surface is coupled to the rotatable plate220aof each actuating device200on opposite sides to each other. This embodiment is particularly useful where the control surfaces are canards towards the front of the aircraft100or where there the aircraft100has only a single tail cone or engine fairing about which a tail surface30a,30bcan rotate.

FIG. 6shows experimental data plotted for extensions of actuators210necessary to achieve different degrees of dihedral displacement of tail surfaces30a,30bat a pitch angle of 6 degrees (i.e. the tail surfaces30a,30bare pivoted about their lateral axes2a,2bby 6 degrees relative to the local horizontal plane). Here, the tail cone60diameter is 1200 mm. The graph shows a slightly non-linear relationship between extension of the actuators210and resultant rotation of the rotatable member220awhen pitch of the tail surface30a,30bis also effected. When the rotatable member220ahas a rotation of zero degrees, all six actuators210are partly extended.

FIG. 7shows experimental data plotted for extensions of actuators210necessary to achieve different degrees of dihedral displacement of tail surfaces30a,30bat a pitch angle of 0 degrees. Here, the tail cone60diameter is 1200 mm. The graph shows a linear relationship between rotation of the rotatable member220aand extension of the actuators210when there is no pitch or yaw effected in the tail surface30a,30b.

The aforementioned embodiments provide a control surface able to move from a first position to a second position by rotating about a point inside the main body of the aircraft100. While a tail surface30a,30bhas been described as moving from the local horizontal plane to a plane at 45 degrees to the horizontal, it would be readily appreciated that the first position need not be the local horizontal plane. For example, the actuating device200could rotate the tail surface30a,30bfrom the plane at 45 degrees to the local horizontal to the local vertical plane of the aircraft100.

By using the actuating device200described above to cause the tail surface30a,30b(or any other control surface) to rotate about a longitudinal axis4of the rotatable member220a, instead of rotating about the root of the tail surface30a,30b, the need for external joints is alleviated. Therefore, radar cross section tends to be reduced.

While plus/minus 22.5 degree motion limits of the rotatable member220ahave been used for the embodiments described above, other limits of angularity may be attained, dependant on actuator210and fixed/rotatable member220a,220bgeometry. For instance, if the stator220band rotor220awere provided by the stator and rotor of an electric motor, the range of inclinations relative to the horizontal plane would be increased.

It will be appreciated that the above described embodiments are purely illustrative and are not limiting on the scope of the invention. Other variations and modifications will be apparent to persons skilled in the art upon reading the present application.

Moreover, the disclosure of the present application should be understood to include any novel features or any novel combination of features either explicitly or implicitly disclosed herein or any generalization thereof and during the prosecution of the present application or of any application derived therefrom, new claims may be formulated to cover any such features and/or combination of such features.