WO2008125868A2

WO2008125868A2 - Aircraft

Info

Publication number: WO2008125868A2
Application number: PCT/GB2008/001344
Authority: WO
Inventors: Peter Jeremy Dodd; Ian Prutton; Norman Baillie; Nigel Brackley; Jolyon Bambridge; Justin Bambridge; Edward Core; Gary Carter; Greg Morgan; Grant Lloyd; Terry Bridle
Original assignee: Peter Jeremy Dodd; Ian Prutton; Norman Baillie; Nigel Brackley; Jolyon Bambridge; Justin Bambridge; Edward Core; Gary Carter; Greg Morgan; Grant Lloyd; Terry Bridle
Priority date: 2007-04-16
Filing date: 2008-04-16
Publication date: 2008-10-23
Also published as: GB0707320D0; WO2008125868A3

Abstract

A wing structure for an aircraft comprises an inboard spar (30) pivotally mounted at an inboard end with respect to an aircraft fuselage, an outboard spar (32) pivotally coupled at one end to an outboard end of the inboard spar (30) a plurality of elongate, overlapping wing elements (26) each of which is attached at one end to a respective spar (30, 32) and extends rearwardly from the spar (30, 32), means (72) for rotating the inboard spar (30) relative to said fuselage, means (76) operable in conjunction with said actuating means (72) for rotating said outboard spar relative to said inboard spar upon operation of said actuating means (72), and means for maintaining said wing elements (26) generally aligned throughout a range of rotational positions of said inboard and outboard spars (30, 32).

Description

AIRCRAFT

The present invention relates to aircraft and in particular to a novel aircraft construction and configuration.

Traditionally aircraft comprise a fuselage, a main plane attached to the fuselage for providing lift and a tail plane and rudder provided at the rear of the aircraft primarily for providing control. The aircraft is controlled in pitch, roll and yaw by controlling movement of various control surfaces provided on the main plane, tail plane and rudder. In particular, the main plane may be provided with ailerons towards its tips which assist primarily in controlling the roll of the aircraft. The tail plane is provided with elevators which are used primarily to control the pitch of the aircraft. The rudder is used primarily to control the yaw of the aircraft. The various control surfaces may be used in conjunction in order to provide desired control. For example, in a turn, both the ailerons and the rudder are typically used. Such a construction has the disadvantage that many control surfaces have to be provided and suitably coordinated to provide a full range of control forthe aircraft.

Since the inception of manned flight, there have been attempts to emulate the flight of birds which are able to control all aspects of flight through their wings and tail. One example of such an attempt is disclosed in FR- A-2697442. This^'document illustrates certain types of manoeuvre which can be achieved in a bird-like construction. However, the disclosure of this document is entirely hypothetical and has no detail at all of how in practice such control may be achieved.

The present invention seeks to provide a practical aircraft whose flight principals are similar kv those of birds. The necessary lift and control can be provided by changing the area, configuration and position of the wing and tail. For example, the lift provided by the aircraft may be controlled by varying the area of the wing and tail, an option which is not available in a traditional aircraft. When an aircraft is flying at low speed, near its stall velocity, then the lifting surfaces of the aircraft should have as large an area as possible. In high speed flight, however, the area of the lifting surface need not be so great due to the increased velocity of air over the wings. Control moments required for manoeuvring can be generated by appropriate movement of the wing and tail.

The present invention in its various aspects provides an aircraft which is capable of adopting a wide range of configurations so as to provide the necessary lift and control essentially through just a wing and a tailplane.

The main lifting force of the aircraft is provided by the wings and from a first aspect, the present invention provides a wing structure which allows the area and configuration of the wing to be changed as required.

From a first aspect, therefore, the invention provides a wing structure for an aircraft, said structure comprising: an inboard spar pivotally mounted at an inboard end with respect to an aircraft fuselage; an outboard spar pivotally coupled at one end to an outboard end of the inboard spar; a plurality of elongate, overlapping wing elements each of which is attached at one end to a respective spar and extends rearwardly from said spar; means for rotating said inboard spar relative to said fuselage; means operable in conjunction with said actuating means for rotating said outboard spar relative to said inboard spar; and means for maintaining said wing elements generally aligned throughout a range of rotational positions of said inboard and outboard spars.

Thus in accordance with this aspect of the invention, a wing has a number of wing elements attached to inboard and outboard spars which are pivotal relative to each other. As the inboard spar is rotated, the outboard spar rotates relative thereto such that the wing can move between retracted and extended positions relative to the fuselage. During this movement, however, the wing elements are maintained generally aligned by suitable means. As the wing extends and retracts, the degree of overlap of the wing elements changes in a controlled manner, thereby varying the wing area.

In the preferred embodiment, the wing elements are maintained aligned by links extending between the wing elements. The links are preferably at least semirigid such that they can maintain the wing elements generally aligned, but permit some relative movement between them. The links could be resilient.

Preferably the links extend generally parallel to the respective spars so as to form a parallelogram type linkage therewith.^' Preferably as the spars rotate, the wing elements will maintain a generally parallel orientation. This is assisted in use by the airflow over the wing.

The wing elements are preferably attached to the respective spars through clevis-type joints. Most preferably, the clevis extends over a section of the wing element so as to prevent movement of the wing element in a vertical direction, thereby preventing or at least limiting flapping of the wing elements in use.

The inboard spar member may be mounted to the fuselage in any manner which allows it to rotate relative thereto about a generally vertical axis.

Any suitable actuator for example a hydraulic, pneumatic or ball screw type actuator, may rotate the inboard spar.

Separate actuators may be provided for the inboard spar of each wing. However, a common actuator may be provided, connected to the respective inboard spars through suitable connections. For example a rotary actuator may be connected to the inboard spars though suitable linkages. The rotation of the outboard spar relative to the inboard spar is achieved by a second actuator. In a first embodiment, the second actuator may be provided between the fuselage and the outboard spar. In such an arrangement, the second actuator may comprise a generally fixed length link extending between the fuselage and the outboard spar. Such an arrangement may, therefore, form a parallelogram- type linkage between the fuselage, inboard spar and outboard spar such that as the first actuator rotates the inboard spar, the link pulls or pushes on the, outboard link in order to cause its movement relative to the inboard spar. Thus the second actuator need not itself be a powered actuator. A single such actuator may be provided which can both push and pull on the outboard spar. However, two second actuators may be provided, both of which only pull on the outboard spar. This has the advantage that each second actuator may be constructed as a tension member, e.g. a cable, rather than a more bulky actuator which would also have to be able to withstand compression forces. Moreover, a tension cable may accommodate any changes in its length which may occur during movement of the spars by simply becoming slack.

In one embodiment, a first second actuator may therefore be attached between the fuselage and a portion of the outboard spar on one side of the pivotal attachment of the inboard and outboard spars while a second second actuator may be attached between the fuselage and a portion of the outboard spar on the other side of the pivotal attachment of the inboard and outboard spars, e.g. to a lug projecting inboard from an end of the outboard spar. The inboard end of the second actuator may be mounted to the fuselage in a translationally fixed position, for example at a ball or other form of pivotal joint. In one embodiment, however, the inboard end of the second actuator may be moved or translated so as to move the outboard spar. For example, the inboard end of the second actuator may be mounted to a bell crank mechanism or other mechanism which is operable to move the position of the inboard end of the second actuator. Separately operable second actuators may be provided for each wing, although in a preferred embodiment, commonly operable second actuators are provided. In a preferred arrangement, respective linkage mechanisms are driven by a common drive. For example a pair of bell cranks may be driven by a rack and pinion or other mechanism such that both outboard spars are moved at the same time. In one arrangement, the second actuators may be operable such that as one actuator operates to extend the outboard spar of one wing, the other operates to retract the outboard spar of the other wing. This can therefore be used as a roll control mechanism. However, when the inboard ends of the second actuators are held in a given position by the drive mechanism, rotation of the inboard spars will cause the outboard spars to extend and retract together.

In a yet further arrangement an extensible and retractable second actuator may be mounted between the inboard and outboard spars. This actuator (which may, for example, be a hydraulic or pneumatic actuator) will be operated in conjunction with the first actuator so that the outboard spar will move in the required manner relative to the inboard spar during operation of the first actuator. This may be more versatile that the first arrangement described above, potentially allowing a wider range of relative positions of the inboard and outboard spars. However, such an arrangement is likely to be more bulky and heavy than the arrangement described above. It is envisaged in such an arrangement that the second actuator will be attached to an end of the outboard spar adjacent its connection to the inboard spar. In the various above arrangements, by actuation of the appropriate actuators, the aircraft wings can be extended or retracted in either a symmetrical or asymmetrical manner. An asymmetrical operation will allow different shapes and areas for each wing, leading to asymmetrical forces on the wings which can be used for control purposes.

In further broad terms, therefore, the invention also provides an aircraft comprising a pair of wings, means for extending or retracting said wings or one or more parts thereof symmetrically, and means for extending or retracting said wings or one or more parts thereof asymmetrically. i The wing is preferably also provided with a tip spar, pivotally connected, directly or indirectly to outboard end of the outboard spar. Preferably a plurality of wing tip elements are attached pivotally to and extend rearwardly from said wing tip spar and are connected by links which position the wing tip elements in a desired relative orientation. In an extended condition of the wing, the wing tip elements may fan out from one another, whereas in a retracted position, they may lie more parallel to one another and to the wing elements attached to the inboard and outboard spars.

A third actuator is preferably provided to move the tip spar relative to the outboard spar. In the preferred embodiment, the third actuator comprises an element which is attached at an inboard end to the inboard spar and which cooperates with the tip spar at its outboard end. Preferably the element is guided by guide means provided on the outboard spar.

The element may be fixedly attached to the tip spar, but in one embodiment the element cooperates with an actuating member provided on the tip spar such that as the inboard spar is rotated, the tip actuator pushes or pulls against the actuating member on the tip spar, causing it to pivot relative to the outboard spar. In one embodiment, respective first and second actuating members are provided on the tip spar, one of which is engaged by the actuating element in a pushing mode and the other of which is engaged by the element in a pulling mode. Preferably the respective tip spar actuating members are spaced apart such that there is a degree of lost motion between the element and the tip spar in certain operating conditions. For example, the outboard end of the third actuator may comprise a pin arranged to slide within a slot provided on the tip spar.

Preferably, the third actuator is arranged so as to fail when a predetermined load on the third actuator is exceeded. For example, in the pin and slot arrangement described above, the pin may be arranged so as to detach from the third actuator in the event of the wingtip contacting the ground upon landing of the aircraft. The advantage of such an arrangement is that upon exceeding a predetermined load on the tip spar, the movement of the tip spar is uncoupled from the inboard and outboard spars. Transmission of the excess load to the inboard and outboard spars is therefore reduced, thus minimising any potential damage to the wing.

A further wing element support may be pivotally attached to the tip spar so as to be selectively pivoted forwardly relative to the wing about a generally vertical axis. Such an arrangement is desirable as it provides an additional degree of control to the aircraft by allowing selective deployment or retraction of a wing element which will increase or decrease wings area and therefore lift. Differential operation of the elements may therefore provide a rolling moment.

This is itself thought to be a novel and inventive arrangement, so from a further aspect the invention provides an aircraft wing having a control element at the tip of the wing mounted for pivotal movement about a generally vertical axis. The control element support may be actuated by a fourth actuator. In one embodiment, the fourth actuator comprises a tension member which acts to pivot the control element about the axis. The control element may be mounted to the tip spar in such a manner that after the control element reaches a limit position with respect to the tip spar, continued pulling by the actuator produces rotation of the wing spar and the control element together. The lost motion mechanism referred to above may accommodate such movement.

Alternatively, the fourth actuator may comprise an actuator mounted between the tip spar and the control element for movement of the control element relative to the tip spar. For example, the fourth actuator may comprise a hydraulic or a pneumatic actuator. Alternatively, the fourth actuator may comprise a worm gear mounted on the tip spar for engagement with a second gear mounted on the control element about the pivotal axis of the control element. Rotation of the worm gear rotates the second gear (and therefore the control element) about the pivotal axis.

However, the fourth actuator preferably comprises a motor driven leadscrew mounted to the tip spar for engagement with a captive leadscrew nut, which is mounted to a portion of the control element located on an inboard side of the pivotal axis. Driving the motor driven leadscrew moves the captive leadscrew nut along the leadscrew, thus changing the length of the leadscrew portion extending between the motor and the control element. Increasing and decreasing the length of the leadscrew portion therefore retracts and deploys the control element respectively. The wing of the aircraft may thus be extended and retracted by means of the arrangement disclosed above. In the extended position, the wing elements overlap to a relatively small degree. The effect of this is that the wing has a relatively large surface area, affording maximum lift. In the retracted position of the wing, the wing elements overlap to a maximum extent, thereby minimising the area of the wing. Each wing element preferably comprises a shaft for attachment to the respective spar and a lifting body mounted on the shaft.

In the preferred embodiment, the lifting body is asymmetric about the shaft axis, having a relatively narrow leading edge portion and a relatively wider trailing edge portion. The wing elements are arranged such that the leading edge portion of one wing element overlaps the trailing edge portion of the adjacent wing element. Preferably, a trailing edge region of the trailing edge lifting body is relatively flexible compared to the leading edge lifting body. It will be understood that in use air pressure will act on the under surface of the wing element such that the trailing edge portion of one wing element will be forced into firm contact with the leading edge portion of the adjacent, overlying wing element. The relative flexibility of the trailing edge portion will allow a good, seal to be formed between the wing elements preventing air escaping from between them. This is itself a novel arrangement, so from a further aspect, therefore, the invention provides an aircraft wing element comprising an elongate shaft and a lifting body mounted to said shaft; said lifting body having a leading edge portion and a trailing edge portion, wherein said leading edge portion is narrower than said trailing edge portion and said trailing edge portion has a trailing edge region which is relatively flexible compared to the leading edge portion.

The leading and trailing edge portions of the lifting body may be formed as^{^} separate bodies and suitably attached to the shaft. However, the respective portions may be formed together, integrally with the shaft. For example, they may be moulded about the shaft in a suitable manner. Typically the lifting body will be of a lightweight, relatively rigid material such as a foam. A suitable protective skin may be provided on the body.

The shaft or its attachment to the spar may be flexible such that it allows the lifting body to twist about its axis to a limited degree. For example, the attachment may incorporate a resilient material such as rubber to accommodate such twisting. Preferably, one or more intermediate wing elements are provided between adjacent wing elements so as to substantially cover any gaps between the adjacent wing elements. The intermediate wing elements are preferably shorter in length than the main wing elements and extend within the region proximate the wing spars. The advantage of this arrangement is thaf it reduces airflow through the wing surface and increase inflow over and underneath the wing, thus increasing the lift provided by the wing. The intermediate wing elements may be provided above or below the main wing elements, or both. However, the intermediate wing elements, if present, are preferably provided above the main wing elements.

Suitable shrouding may be provided over the wings so as to cover the spars, ^■ wing element attachments, actuators and so on. Preferably, the shrouds are attached to the respective spars and are configured such that they may overlap one another as appropriate. The internal space of the shroud may be filled with a filler material, such as a foam material, e.g. a latex, or synthetic foam in order to provide some rigidity to the shrouding.

Preferably the filler material is resiliency deformable such that as the spars move, the chordwise profile of the wing also changes. In particular, as the wing is pulled in, the filler material is compressed such that the wing preferably becomes fatter in the vertical direction. This may be desirable both aerodynamically and for strength reasons, especially in front of the wing spars. From a yet further aspect, the invention provides a wing having relatively rotatable inboard and outboard spars, one or more shrouds covering said spars and a resiliently deformable filler material arranged within said shroud(s).

Of course, the shrouding material itself may provide sufficient rigidity for the purpose.

The leading edge of the wing may also be configured such that as the wing extends, the leading edge of the wing drops to increase the camber on the inboard part of the wing. This is advantageous in that it provides increased lift which is advantageous when the aircraft is in low speed e.g. a landing configuration. From a further aspect therefore, the invention provides an aircraft having a wing which is pivotable about a generally vertical axis, the wing being configured such that as an inboard portion of the wing pivots rearwardly, the leading edge of the wing moves so as to increase the camber of an inboard section of the wing.

The desired effect can be achieved by providing a control member extending along the leading edge region from the fuselage to an outboard portion of the wing, the attachment of the control member to the fuselage being positioned such that during rotation of the wing, the control member becomes taut and pulls the leading edge down, thereby increasing its camber.

As will be described further below, the wings of the aircraft can be operated in unison or independently in order to provide desired flight characteristics.

Preferably the wing is configured such that in the event of a power failure, it will move to a maximum surface area configuration, thereby facilitating landing. The respective wing spars may be resiliently biased towards such a position.

The above discussion focuses on the extension and retraction of a wing mechanism in order to change its configuration and surface area. In addition to the above, however, it is also desirable to allow the wing to assume a dihedral which is appropriate for its particular mode of operation. For example, in low speed flight, for example when the aircraft is coming into land, it may be desirable to have a relatively high angle of dihedral in order to improve stability. Conversely, when operating at high speeds, for example when the wings will be retracted, a smaller dihedral can be tolerated by the aircraft. Accordingly, from a further aspect, the invention provides an aircraft comprising a wing mounted to a fuselage, said mounting being such as to accommodate variations in an angle of dihedral of the wing occurring due to variations in an operative configuration of the wing.

In the preferred embodiment, therefore, the wing mounting should be able to accommodate not only rotation of the wing about a generally vertical axis, but also about a generally longitudinal axis of the aircraft.

From a further broad aspect, therefore, the invention provides an aircraft comprising a wing mounted to a fuselage, said mounting being such as to permit rotation of the wing about a generally vertical axis, and also about an axis extending in a longitudinal direction of the aircraft. In one embodiment, the inboard spar may be rotatably mounted on a support which is rotatably mounted on the fuselage.

In the preferred embodiment, the support is mounted for rotation about a generally vertical axis, allowing the wing to extend and retract relative to the aircraft fuselage as described above. The wing is mounted on the support such that it may rotate about an axis extending in a direction along the length of the fuselage, i.e. such that it may rotate up and down with respect to the fuselage.

In the above described embodiment, the inboard spar may, therefore, be mounted on a suitable pivot on the support.

Preferably the change in dihedral is accommodated by spring means provided in the mounting. The springs provide an appropriate reaction force counteracting the lift provided by the wing at any given angle of dihedral.

In a preferred embodiment, spring means may be provided which counter displacement of the wing in both upward and downward directions. In such arrangement, the spring resisting upward motion of the wing will tend to be stronger than the other, as the lifting loads tend to rotate the wing upwardly

Any suitable form of spring may be used, for example simple compression or tension springs although hydraulic springs could equally be used.

It will also be appreciated that it would be advantageous for the mounting arrangement to provide a degree of shock absorbing within the system such that sudden changes in lift, as might occur during turbulent flight will, to some extent be absorbed. To this end, one or more dampers may be provided in the mounting arrangement. From a further aspect, therefore, the invention provides an aircraft having a wing mounted to an aircraft fuselage for vertical movement relative thereto, said mounting comprising damping means for damping such vertical movements.

Thus in a preferred embodiment, the mounting may comprise one or more spring dampers.

Preferably means are provided to limit the upward pivoting of the wings relative to the fuselage. In the preferred embodiment, a tether member is provided extending between the wing and the fuselage.

From a further aspect, therefore, the invention provides an aircraft having a pair of wings pivotally mounted to a fuselage, and comprising tether means attached between the wings and the fuselage to limit rotation of the wings relative to the fuselage.

In a preferred embodiment, tethers extend between the wings and a keel provided at or adjacent a lower part of the fuselage. The advantage of attaching a tether member to the fuselage, and in particular to a keel region of the fuselage, is that at least some of the lift force applied to the fuselage during a limit excursion will be applied through the tether to the fuselage rather than that force being transferred to the aircraft through the pivotal mounting. The forces can be effectively be distributed throughout the cross section of the fuselage. Preferably, the tethers are relatively inextensible so as to advantageously provide automatic adjustment of the wing dihedral across ₍a range of wing configurations. Such automatic adjustment arises because the attachment point of the tethers to the wings naturally moves forward and rearward relative to the fuselage as the wings are extended and retracted respectively. Since the length of the tethers is substantially fixed, the dihedral angle of the wings is forced to increase and decrease as the wings are extended and retracted respectively.

From a further aspect, therefore, the present invention provides an aircraft having a wing mounted to a fuselage, said mounting being such as to permit rotation of the wing about a generally vertical axis, said aircraft further comprising means for automatically changing a dihedral angle of said wing in response to said rotation of said wing. .

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However, the tethers or their attachment to the wings and fuselage preferably have some degree of resilience such that during flight any sudden forces, e.g. due to turbulence, acting on the wings can lead to just a change in dihedral of the wings, rather than causing a change in the actual wing configuration which would occur if the tether were completely inextensible.

The tethers may be individual tether members, or they may be formed as a unitary body.

The main wings of the aircraft primarily provide lift for the aircraft, although depending on their mode of operation could also be used in the control of the aircraft. The tail plane of the aircraft can, however, contribute both to the lift and to the control of the aircraft. It will be understood that in order to change the lift provided by the tail plane, its area can be varied, as is the case of the main wing, as described above. From a further aspect, the invention provides a tail plane construction which facilitates such operation. From a further aspect, therefore, the invention provides an aircraft tail plane comprising a tail chassis; a plurality of tail elements pivotally attached to said chassis and at least partially overlapping one another; and actuating means for rotating said tail elements relative to each other so as to vary the degree of overlap between the tail elements such that the surface area of the tail can be changed. In a preferred embodiment, the tail chassis is generally V-shaped in section having diverging limbs. Respective tail elements are attached to and extend rearwardly from these limbs. The tail elements may have a similar construction to the wing elements discussed above, i.e. having a shaft which is attached to the tail chassis, e.g. through a clevis arrangement. Preferably, the tail elements are symmetrically arranged around a central tail element. Preferably the central tail element is relatively rigid and overlies the adjacent tail elements of either side. This provides a reaction surface for the adjacent tail members.

Any suitable actuation means may be provided for pivoting the tail elements. It would be possible, for example, to have individual actuators on individual tail elements in order to effect the movement. Preferably, however, a common actuator is provided for moving the tail elements simultaneously. In the preferred embodiment, the actuator comprises a linkage extending between the tail elements and connected to a common actuator.

In the preferred embodiment, the linkage comprises a cable or cables interconnecting the tail elements. The cable is moveable by a suitable linear or rotary actuator in order to rotate the tail elements. Most simply the actuator may be a rotary shaft around which the cable is wound. Alternatively, the actuator may comprise an extensible and retractable actuator suitably attached to one or more tail elements. The actuating movement may be transferred between adjacent tail elements by the links between them. Preferably the tail elements are resiliently biased towards a failsafe configuration in the event that the actuator should fail. The failsafe configuration is preferably the tail extended configuration.

Preferably the tail chassis is mounted for rotation about a pitch axis such that the tail can be pitched up and down to change its angle of attack. By pitching the tail up and down, the moment acting upon the fuselage changes thereby providing pitch control.

Preferably the tail is rotated in pitch about its leading edge. This is in itself a novel arrangement so from a further aspect, the invention provides an aircraft having a tail plane which is rotatable in pitch about its leading edge. The tail chassis is preferably mounted for such rotation on a support which is rotatable about a roll axis whereby the tail may be rotated about an axis generally longitudinal of the aircraft. For example, the support may have a journal supporting a rotatable shaft of the chassis about the pitch axis.

An actuator for effecting the pitch movement is preferably also mounted on the support, so as to move with the support. In a simple embodiment, the actuator may comprise, a simple actuated arm attached to the tail chassis.

The support member may be a shaft mounted for rotation about the roll axis. Any suitable means may be provided for rotating the support member about the roll axis. Any suitable actuator may be used, but in a simple embodiment, a rack and pinion type actuator may be provided. In a particularly preferred embodiment, a pair of actuators may be provided acting on opposed parts of the shaft so as to balance the rotational forces acting thereon. The tailpane described above, i.e. one which can rotate about both pitch and roll axes, may be suitably integrated with or fixedly mounted to the aircraft fuselage. Preferably the however, the tail plane is pivotable as a unit with respect to the fuselage. This will allow the relative orientations of the fuselage and the tail unit to change as may be advantageous in circumstances as described further below.

In conventional aircraft, the wings and tail plane are maintained in a relatively fixed relative configuration, e.g. generally co-planar or at a fixed vertical spacing from each other. However, it has been recognised, that in certain flight modes, for example particularly when landing, it will advantageous to change the relative positions of the wings and tail plane.

From a further aspect, the invention provides an aircraft comprising a main plane, a tail plane and fuselage, said main plane and said tail plane being translatable one relative to the other so as to change their relative positions.

Preferably the fuselage comprises a first section to which is mounted the main plane and a second section which is moveable with respect to said first fuselage section so as to change the relative positions of said main plane and said tail plane.

From a further aspect, the invention provides an aircraft comprising a main plane, a tail plane, a first fuselage section and a section fuselage section, said main plane being mounted to said first fuselage section and said tail plane being mounted to said second fuselage section, and means for moving said second fuselage section with respect to said first fuselage section so as to change the relative positions of said main plane and said tail plane.

Preferably the first and second fuselage sections are pivotally connected together, and a first actuator provided to effect pivotal movement between the two.

In a particularly preferred embodiment, the relative planar relationship between the main plane and the tail plane remains generally the same in the absence of any independent control movement of the tail plane as described above.. Thus in one embodiment, the main plane and the tail plane may remain generally parallel in both configurations.

In such arrangements, a further actuator should be provided so as to rotate the main plane or the tail plane relative to its respective fuselage section so as to maintain the desired positional relationship between the main plane and the tail plane. Most preferably an actuator is provided to rotate the tail plane relative to the second fuselage section.

This actuator may be provided on the second fuselage section and be operated in parallel with the first actuator in order to obtain the necessary relative planar configuration between the main plane and the tail plane. In a further, preferred embodiment, however, the second actuator may comprise a fixed length link extending between the first fuselage section and the tail section such that rotation of the second fuselage section relative to the first fuselage section automatically results in rotation of the tail section relative to the second fuselage section. In effect, the second fuselage section acts as a link of a parallelogram type linkage.

In normal flight, the main plane and tail plane would be spaced apart by a predetermined vertical distance. However, the vertical spacing between the main plane and the tail plane is preferably increased during landing. This in effect, brings the second fuselage section closer to the ground than the first fuselage section which facilitates landing. To this end, an undercarriage is preferably provided in the second fuselage section.

From a further aspect, the invention provides an aircraft having a first, forward fuselage section and a second, rearward fuselage section, said sections being rotatable relative to one another and further comprising an undercarriage mounted to said second fuselage section.

In a simple embodiment, the undercarriage comprises a collapsible strut which is retractable into the fuselage generally axially of the aircraft. The strut may comprise a plurality of pivotally interconnected and relatively collapsible links.

The undercarriage preferably comprises ground engaging members, e.g. wheels, spaced apart longitudinally on the end of the strut. This will avoid the need for a nose wheel or tail wheel. Due to the construction of the preferred embodiment, in flight, the centre of gravity of the aircraft will be forward of the undercarriage. However, the movement of the fuselage sections one relative to the other will bring the centre of gravity over the undercarriage, between the ground engaging members. This will allow for a stable landing on the undercarriage. Preferably the front fuselage section comprises a cockpit for one or more passengers.

The aircraft may be powered or unpowered. In a powered aircraft, the second fuselage portion preferably comprises an engine. By virtue of the rotation of the fuselage sections one relative to the other, the line of thrust of the engine may be changed. This is particularly useful in landing where the aircraft is travelling at low speeds near stall. Since the second fuselage section is pitched downwardly relative to the front fuselage section, the engine provides a downward thrust which will assist in maintaining the aircraft in the air even at low speeds. This means that the aircraft will be able to touch down at relatively low forward speeds, even speed below the stall speed of the wings.

This in itself is a novel arrangement, so from a further aspect the invention provides an aircraft having a front fuselage section and a rear fuselage section, said rear fuselage section being rotatable relative to said front fuselage section, said aircraft further comprising an engine provided in said rear section whereby the line of action of the thrust of the engine relative to the front fuselage section can be changed.

The engine may be any suitable type but is preferably a shrouded fan, turbo fan or turbo jet engine. A preferred embodiment of the invention will now be described by way of example only with reference to the accompanying drawings in which:

Figure 1 a shows, schematically, an aircraft in accordance with the invention in a first, landing configuration;

Figure Ia shows, schematically, an aircraft in accordance with the invention in a first, high speed configuration;

Figure Ib shows, schematically, the aircraft of Figure Ia in a second, lower speed configuration;

Figure Ic shows, schematically, the aircraft of Figure I a in a third, cruise configuration; Figure Id shows, schematically, the aircraft of Figure Ia in a fourth, landing configuration; Figure 2 shows the wing of the aircraft in the configuration of Figure Ic schematically in more detail;

Figure 3 shows the wing of the aircraft in the configuration of Figure Ib schematically in more detail; Figure 4 shows the wing of Figure 2 with its external shrouding removed;

Figure 5 shows the wing of Figure 3 with its external shrouding removed;

Figure 5A shows an alternative wing actuation mechanism;

Figure 5 B shows an alternative tip spar actuation mechanism;

Figure 5C shows an enlarged view of a portion of the actuation mechanism of Figure 5B;

Figure 5D shows an alternative actuation mechanism for the outermost wing element;

Figure 5E shows an embodiment of the wing comprising a plurality of intermediate wing elements; Figure 5F shows a further wing actuation mechanism;

Figure 6 shows a detail of the mounting of the wing elements of Figures 2 to 5;

Figure 6A shows an alternative mounting arrangement for the wing elements of Figures 2 to 5; Figure 7 shows a section through a plurality of wing elements;

Figure 8 shows a section through the wing of Figure 2 along the line XX- XX;

Figure 9 shows a schematic front elevation of the aircraft of Figure 1 ;

Figure 10 shows a detail of the arrangement of Figure 9 taken along the line XX of Figure 4;

Figure 1 1 shows the tail plane of the aircraft of Figure 1 in a first configuration;

Figure 12 shows the tail plane of Figure 1 1 in a second configuration;

Figure 13 shows the tail plane of Figure 1 1 in a third configuration; Figure 14 shows the tail plane of Figure 1 1 in a neutral attitude;

Figure 15 shows the tail plane of Figure 14 in schematic perspective; Figure 16 shows the tail plane of Figure 14 in upwardly and downwardly deflected positions;

Figure 17 shows the tail plane of Figure 16 in the upwardly deflected position, in schematic perspective; Figure 18 shows the tail plane of Figure 16 in downwardly deflected position in schematic perspective;

Figure 19 shows the tail plane of Figure 15 rotated in a first direction about a roll axis;

Figure 20 shows the tail plane of Figure 14 rotated in the opposite direction around a roll axis;

Figure 2OA shows an alternative tail plane arrangement;

Figure 2OB shows the elevator control mechanism of the tail plane of Figure 2OA;

Figure 2OC shows the actuation mechanism for the tail elements of the tail plane of Figure 2OA;

Figure 2OD shows the actuation mechanism for controlling lateral movement of the tail plane of Figure 2OA;

Figure 2OE shows an alternative view of the actuation mechanism of Figure 2OD; • Figure 21 shows the fuselage, wings and tail plane of the aircraft of Figure 1 in a first orientation;

Figure 22 shows the aircraft of Figure 21 in a second configuration;

Figure 23 shows under carriage of the aircraft of Figure 22 in a semi- deployed condition; and Figure 24 shows the under carriage of the aircraft of Figure 23 in a fully deployed position.

With reference to Figures I a to Id, an aircraft 2 in accordance with the invention comprises a first, forward fuselage section 4, a second, rear fuselage section 6, a main plane 8 comprising a port wing 10 and a starboard wing 12 and a tail plane 14. As is illustrated schematically in Figures Ia to Id, the main plane 8 is moveable between a number of configurations. The Figures show the wing configuration in both plan and front elevation views, with sections of the wing being shown at noted respective stations along the wing.

In the first configuration illustrated in Figure Ia,, the wings 10, 12 are retracted towards the fuselage 3. The tail plane 14 is also drawn in to a generally rectangular configuration. The wings 10, 12 have a slight anhedral. This configuration is suitable for high speed, e.g. diving flight. .

In the second, lower speed configuration illustrated in Figure Ib, the wings 10, 12 are extended slightly. This increases the area of the wings 10, 12. The wings 10, 12 still have a slight dihedral. The chordwise profile of the wings 10, 12 at an inboard station A is relatively thick as illustrated. The tail plane 14 is slightly extended to provide an increased area and has a slight anhedral.

In the third, cruise configuration illustrated in Figure Ic, the wings 10, 12 are extended further to provide a greatr surface area. The wings 10, 12 have a slight dihedral. The chordwise profile of the wings 10, 12 at the inboard station A is less thick than that of the wing in the configuration shown in Figure Ib. The tail plane 14 is also extended further to provide a greater area still and has a neutral dihedral.. In the fourth, landing configuration illustrated in Figure Id, the wings 10, 12 are extended fully to the extent that their tips extend forwardly. This is the maximum surface area configuration of the wing. The wings 10, 12 have a significant dihedral. The chordwise profile of the wings 10, 12 at the inboard station A is less thick than that of the wing in the configuration shown in Figure Ib, however, the camber of the wing does significantly increase. The tail plane 14 is also extended further to provide maximum tail plane area and has a significant dihedral. Figures 3 to 10, show a wing of the aircraft of Figures I a to Id in more detail. The configuration shown in Figure 2 and 4 corresponds generally to the configuration shown in Figure 1 c, while the configuration shown in Figures 3 and 5 corresponds generally to the configuration A in Figure Ib.

The wing 12 comprises three relatively moveable sections, namely an inboard section 14, an outboard section 16 and a tip section 18. The outer surface of each wing section is provided by a suitable shroud 20, 22, 24 which surrounds various wing supports and actuators as will be described further below. Extending from the trailing edge of the shrouds 20, 22, 24 are wing elements 26 which might be compared to feathers on a bird wing.

Figures 4 to 10 show further details of the wing mechanism. The wing 12 comprises an inboard spar 30, an outboard spar 32 and a tip spar 34. The inboard spar 30 has an inboard end 36 which is mounted by a clevis 38 to a turntable 40.

The turntable 40 is rotably mounted on a support 42 formed in the first fuselage section 4 by means of a spindle 44. A suitable bearing material 46 is formed between the turntable 40 and support 42. The inboard spar 30 further comprises an outboard end 48. The outboard spar 32 has an inboard end 50 which is pivotably mounted to the outboard end 48 of the inboard spar 30 about an axis 52.

The outboard spar 32 further has an outboard end 54 which is pivotably connected to an inboard end 56 of the tip spar 34 about an axis 57.

A plurality of wing elements 26 are mounted to the respective spars 30, 32,

34 around respective pivots 58. As shown in more detail in Figure 6, the inboard and outboard spars are provided with clevis plates 60 which support a shaft 62 of the wing element 26 by means of a clevis pin 64. The plates 60 prevent excessive flapping of the wing elements 26 in the vertical direction.

Figure 6A shows an alternative mounting arrangement for mounting the wing elements 26 to the inboard and outboard spars 30, 32. In the embodiment shown in Figure 6A, the clevis pin 64 extends beyond the clevis plates 60 and is connected at each end to the shaft 62 by upper and lower flexible support rods 300.

As shown in part in Figure 6A, each pair of upper and lower support rods 300 may be formed as a single support rod that extends past the respective spar 30, 32 and around the leading edge of the wing. Such an arrangement not only provides additional support to the shaft 62 to prevent excessive flapping of the wing elements

26, but also assists in providing the desired leading edge shape of the wing. For additional support, one or more support fins 302 may also be provided between the clevis pin 64 and the shaft 62, as shown in Fig. 6A.

The wing elements 26 may be mounted to the tip spar 34 using a similar arrangement to those shown in Figures 6 and 6 A, although as shown, the tip spar 34 is provided with a recess on its trailing edge which receives the wing elements 26, rather than having individual mounting clevises. The outermost wing element 66 is mounted to a separate support 68 which is pivotably mounted to the tip spar 34 about a pivot axis 70 as will be described further below.

Various actuators are provided in order to move the wing 12 between its various configurations. In a first embodiment shown in Figures 4 and 5, an actuator 72 which is fixed to the first fuselage section 4 is connected to a portion 74 of the inboard end 36 of the inboard spar 30 which projects beyond the clevis 38. Extension and retraction of the actuator 72 will cause the turntable 40 and thus the inboard spar 30 to rotate about the spindle 44. In an alternative arrangement shown in Figure 5 A, each turntable 40 is connected via a pushrod 304 to a common turntable actuator 306, the turntable actuator 306 being rotatably mounted to a servo 308. As shown, the pushrods 304 are pivotably connected to the turntable actuator 306 in an offset arrangement, such the rotation of the turntable actuator 306 via the servo 308 will cause the turntables 40 and thus the inboard spars 30 to rotate about the respective spindles 44. As in the earlier embodiment, the turntable 40 permits pivoting of the inboard spar 30 about a longitudinally extending axis of the aircraft. Thus the inboard end of the inboard ' spar may pivot about an axis extending front to back of the turntable 40. Moreover, the turntable may incorporate spring dampers for resisting and damping vertical movement of the inboard end of the inboard spar 30. For example, the internal volume of the turntable may be filled with a resilient damping material.

Additionally, respective second actuators 76a, 76b extend between the first fuselage 4 and the outboard spar 32. Both actuators are in the form of tension members e.g. cables. One second actuator 76a extends from a mounting location 77a on the first fuselage section 4 to a mounting lug 79a provided on the front of the outboard spar 32 outboard of the pivot axis 52. The other second actuator 76b extends from a mounting location 77b on the first fuselage section 4 to a mounting lug 79b provided on the outboard spar 32 on the other side of the pivot axis 52. As the first actuator 72 extends and retracts to rotate the inboard spar 30, one or other of the second actuators 76a, 76b will pull on a lug 79a, 79b and so cause the outboard spar to extend or retract. Alternatively, a powered actuator 76c (shown in dotted lines) may extend between the inboard spar 30 and the outboard spar 32. Extension and retraction of the second actuator will pivot the outboard spar 32 relative to the inboard spar 30 about the axis 52. In the embodiment shown in Figures 4 and 5, a further actuator in the form of a rod 78 which is, at least at its outboard end 80 relatively flexible, extends from a pivotable joint 82 at its inboard end 84 on the inboard spar 30 through a guide 86 provided on the distal end 54 of the outboard spar 32. The outboard end 88 of the rod 78 is formed with a pushing surface 90 for cooperation with a block 92 provided on a leading edge of the tip spar 34 and a pulling surface 91 for cooperation with a collar 93 provided on the leading edge of the tip spar 34 inboard of the block 92.

In an alternative embodiment shown in Figures 5B and 5C, the outboard end 80 of the rod 78 is formed as a cable 81 which is provided with a pin 310 which engages a slot 312 provided on a leading edge of the tip spar 34. The cable 81 is guided through a tube or channel (not shown) mounted to the outboard end of the outboard spar 32. The pin 310 cooperates with one end 314 of the slot 312 in order to extend the tip spar 34. The tip spar is biased towards its retracted position by means not shown. Preferably, the pin 310 is arranged to detach from the outboard end 88 of the rod 78 when a predetermined load on the tip spar 34 is exceeded, e.g. in the event of the tip spar 34 contacting the ground during the landing of the aircraft. The advantage of this arrangement is that the detachment of the pin 310 from the rod 78 effectively uncouples the movement of the tip spar 34 from the inboard and outboard spars 30, 32. Transmission of the excess load to the inboard and outboard spars 30, 32 is therefore reduced, thus minimising any potential damage to the wing.

In the embodiment shown in Figures 4 and 5, a third actuator, in the form of a Bowden cable or similar 94 runs generally along the leading edge of the wing 12. The cable 94 also passes through the guide 86 mounted on the outboard spar 32. At its inboard end 96, the outer sheath 97 of the cable 94 is connected to the first fuselage section 4 and at its outboard the outer sheath is connected to the outboard spar 32, for example to the guide 86. The inboard end of the inner cable 99 of the cable 94 is connected to an actuator 101. The outboard end of the inner cable 98 is attached to the support 68 for the outermost wing element 66.

The attachment of the Bowden cable 94 to the first fuselage section 4 is displaced downwardly from the attachment of the wing to the fuselage for reasons which will be explained further below.

In an alternative embodiment shown in Figure 5D, the third actuator comprises a servo 318 mounted to an outboard portion of the tip spar 34. The servo 318 drives a leadscrew 320 which engages a captive leadscrew nut 322 mounted on an inboard portion of the support 68 of the outermost wing element 66. By driving the leadscrew 320 with the servo 318, the captive leadscrew nut 322 moves along the leadscrew 320, thus rotating the support 68 and the outermost wing element 66 about the pivot 70.

The respective spars are resiliently biased, by means not shown, towards the fully extended configuration shown in Figure Id. In addition to the above, the respective shafts 62 of the wing elements 26 are connected together by links 100. The inboard most link 102 is rotably fixed to the front fuselage section 4. The links 100 are sufficiently rigid to maintain the relative positions between the wing elements 26 as the wing extends and retracts.

Extension and retraction of the wing 12 will be described with reference to Figures 4 and 5 (and with reference to the alternative embodiments of Figures 5A- 5D in parenthesis).

Starting from the configuration shown in Figure 4, should the pilot of the aircraft wish to retract the wing 12 to the position shown in Figure 5, the actuator 72 is extended (the turntable actuator 306 is rotated anticlockwise by the servo 308). Extension of the actuator 72 (rotation of the turntable actuator 306) causes the turntable 40 and thus the inboard spar 30 to rotate about spindle 44 thereby drawing the inboard spar 30 closer to the fuselage 3. This rotation also causes the actuation cable 76a to pull the outboard spar 32 towards the inboard spar 30 about the rotational axis 52, due to the relative positions of the axis 52 and the point of attachment oft the actuator cable 76a to the lug 79a. The second actuator cable 76b may slacken slightly to accommodate this movement. As this rotation occurs, the links 100 between the wing elements 26 attached to the inboard and outboard spars 30, 32 maintain the generally parallel relationship of the wing elements 26.

The rotation of the inboard and outboard spars 30, 32 leads to the outboard end 88 of the rod 78 pressing against the block 92 on the tip spar 34, causing that to rotate clockwise about the axis 57. (The tip spar 34 is biased to rotate clockwise when the pin 310 slides along the slot 312). This brings the wing tip elements 26 inwardly towards the more inboard elements 26.

In order then to extend the wing 12 from the position shown in Figure 5 to the position shown in Figure 4, the actuator 72 is retracted (the turntable actuator 306 is rotated clockwise by the servo 308). This causes the inboard spar 30 to rotate away from the fuselage 4. This in turn causes the second actuator cable 76b to pull on the second lug 79b so as to extend the outboard spar 32. With this movement, the pulling surface 91 of the third actuator 78 engages the collar 93 (the pin 310 engages the first end 314 of the slot 312) provided on the tip spar 34 causing it to rotate anticlockwise about the axis 57.

The support 68 may be moved relative to the tip spar 34 to its extended position by pulling on the inner cable 99 of the Bowden cable 94 through the actuator 101 (by driving the leadscrew 320 so as to move the captive leadscrew nut along the leadscrew 320 towards the servo 318). Such a movement will lead to an increase in lift as the wing area increases further, allowing a rolling moment to be generated by differential operation of the supports on either wing.

Referring to the embodiment of Figure 5F, this is somewhat of a hybrid between the embodiments of Figure 1 to 5 and Figures 5A to 5C.

In this embodiment, the inboard spar 30 of each wing is mounted to a turntable 40 as in the embodiment of Figure 5 A, and actuated by a common actuator not shown. However, each second actuator 76 is not fixedly mounted to the aircraft fuselage as it is in the earlier embodiments, but attached at its inboard end to one limb 500 of a bell crank 502 mounted to the fuselage about an axis 503. The other limb 504 of the bell crank 502 is attached to one end of a rack 506 which is driven by a pinion 508. A spring tensioner 510 is connected between the outboard spar 32 and the aircraft fuselage and maintains the second actuator 76 in tension. It will be understood that in the position shown in Figure 5F, when the common inboard spar actuator operates with the pinion 508 locked, the wings will extend or retract symmetrically, the second actuator 76 in effect being fixed at its inboard end. However, should the pinion 508 be operated in the configuration of Figure 5F, movement of the rack will, through the pivoting of the bell cranks 502 about their axes 503 simultaneously cause clockwise pivoting of one outboard spar 32 and anticlockwise pivoting of the other outboard spar 32. The spring tensioner 510 can change length to accommodate the change in geometry of the inboard and outboard spars 30, 32 and the second actuator 76, shortening as the outboard spar extends and lengthening as it retracts. Thus the outboard section of one wing will extend and the outboard section of the other wing retract, providing a difference in wing area and thus lift, producing, inter alia, a rolling moment and increased drag on one wing, allowing steering of the aircraft to be effected.

As discussed above, the wing 12 is mounted on a turntable 40 through a clevis arrangement 38. The clevis arrangement 38 allows the dihedral of the wing 12 to vary as shown schematically in Figure 9 by the arrows D. To accommodate this movement, the inboard extension of the inboard spar 30 extends into a compartment 1 10 formed on the turntable 40. The enclosure 1 10 has a lower surface 1 12 and an upper surface 1 14. A first spring damper 1 16 extends between an upper surface of the inboard spar extension 74 and the upper surface 1 14 of the enclosure 1 10, while a second spring damper 1 18 extends between the lower surface of the inboard spar 74 and the lower surface 112 of the enclosure. The lower spring damper 1 18 has a higher spring force than the upper spring damper 1 16 as it must counteract the lift loads transmitted through the inboard spar extension 74. However, the upper spring damper 116 acts as a shock absorber to counteract any oppositely directed forces created during flight.

In addition to the spring dampers 1 16, 1 18, tethers 120 extend from the respective inboard spars 30 to a keel 122 formed at the bottom of the fuselage 4. The tethers 120 are formed from a relatively inextensible material and as such act as limiters on the upward movement of the wings 10, 12. They also assist in transmitting lift forces into a lower part of the fuselage, thereby more evenly spreading the load around the fuselage. They will also provide automatic adjustment of the dihedral angle of the wings as the wings are extended and retracted. In particular, when the wings are retracted, the tethers (which are attached to the fuselage generally forward of the wings) will tend to pull the wings down, thereby decreasing their dihedral, in certain circumstances to the point of them having anhedral. When the wings are extended, the tethers allow the wings to pivot up under aerodynamic loading, thereby increasing their dihedral.

^■ The tethers should preferably not be completely inextensible so as to allow some small changes in dihedral in response to transient loads such as turbulence without a resultant change in wing configuration which may be undesirable.

A further tether 124 is provided around the top of the fuselage and connected to the wings 10, 12. This tether 124 limits the droop of the wings 10, 12 when the aircraft is at rest.

It will be seen from Figures 4 and 5 in particular, that the wing elements 26 overlap. This is to provide a substantially uninterrupted wing surface. As shown in greater detail in Figure 7, each wing element 26 comprises a generally U-shaped shaft portion 130 which, shown in Figure 8, is attached to a connector 132 for attachment to the wing spars'30, 32, 34. The wing elements 26 are generally asymmetrical about the axis 133 of the wing element 26. In particular, a typical wing element 26 has a relatively narrow leading edge portion 134 and a relatively wider trailing edge portion 136. The leading edge portion 134, and an inner portion 138 of the trailing edge portion 136 are relatively rigid to provide strength to the element 26 and to transmit aerodynamic forces into the shaft 130. However, a trailing portion 140 of the trailing edge portion 136 is relatively flexible. Such an arrangement means that the trailing edge 140 of the trailing edge portion 136 may be pushed upwardly by air pressure against the leading portion 134 of an adjacent wing element to seal the gap between adjacent wing elements 26.

In the embodiment shown in Figure 5E, the wing 10, 12 is provided with intermediate wing elements 324 mounted between the wing elements 26 in order to substantially cover any gaps between the adjacent wing elements 26. The intermediate wing elements 324 are preferably shorter in length than the wing element 26 and extend within the region proximate the wing spars 30, 32, 34. The intermediate wing elements 324 may be provided above or below the wing elements 26, or both. The construction and mounting of the intermediate wing elements 324 may be the same as described above with reference to the wing elements 26.

Referring to Figure 8, further detail of the wing construction can be seen. The leading edge portion of each wing 10, 12 is covered by a sheath 22. The space within the sheath is filled by a resiliently deformable foam material 150. The foam 150 envelops the spars 30, 32, 34 and as the wing retracts will be compressed. This will cause the thickness of the wing, particularly in the inboard wing section to increase, giving greater structural rigidity and also changing the camber of the wing. Similarly, as the wing extends, the foam will relax returning to a thinner configuration..

It will also be understood that as the wing extends, the Bowden cable 94 encased within the leading edge portion of the wing will, due to its low attachment point to the fuselage 3, tend to pull the leading edge of the wing down in its inboard portion at least. This acts to increase the camber on the wing in this region thereby significantly increasing its lift, which is important at low speeds e.g. landing. This is shown schematically in Figure Id.

Turning now to Figures 1 1 to 20, details of the tail plane 14 are shown. The tail plane 14 comprises a generally V-shaped chassis 150 comprising upper and lower V-shaped plates 152, 154. A plurality of tail elements 156 are pivotably mounted between the upper and lower plates 152 by means of pivot pins 158. The construction of the tail elements 156 may be generally similar to that of the wing elements 26. Thus, generally, the tail elements 156 may be asymmetrical in construction. However, a central tail element 160 is formed generally symmetrically and is relatively rigid so as to provide a reaction surface for the tail elements 156 which underlie it on either side. A shroud, not shown, is provided over the chassis 150 and the adjacent ends of the tail elements 156.

The tail plane chassis 150 is mounted on a shaft 162 which is journaled in an opening 164 provided in a second shaft 166. As can be seen in Figure 1 1 , pulleys 168 are mounted at either end of the shaft 162. These receive a control cable 170 which is coupled to the shafts of the tail elements 156. The cables are wound around a drum 172 accommodated within the central opening 174 of the chassis 150. The arrangement is such that as the drum 172 is rotated the cable 170 winds onto the drum thereby moving the tailelements from a closed configuration as shown in Figure 12, through an intermediate configuration shown in Figure 13 to the fully open configuration shown in Figure 11. Of course other arrangements are possible. For example, the drum may be replaced by respective extensible and retractable actuators acting, through a suitable linkage on one or more tail elements.

The tail elements may be resiliently biased to an open configuration and the actuator or actuators act against that biasing to close the tail, such that in the event of actuator failure, the tail opens. This is useful in landing.

It will be appreciated that the tail area is significant and may provide significant lift to the aircraft. This is in contradistinction to traditional aircraft where the tail plane is used primarily as a control surface rather than a lift generating surface. The tail chassis 150 may be rotated around the axis 180 of the shaft 162

(which is effectively at the leading edge of the tail plane) by means of an actuator 182 mounted- on the second shaft 166. In particular, a yoke 184 is coupled to the shaft 162 such that when the actuator 182 retracts, the chassis 150 pivots upwardly as shown in Figure 17. Similarly, if the actuator 182 is extended, then the chassis 150 pivots downwardly about the axis 180. This allows the angle of incidence of the' tail plane to be varied, thereby providing pitch control to the aircraft.

As will also be seen from Figure 11 , the second shaft 166 is journaled in a housing 190 so as to be rotatable about a longitudinal axis 192. This rotation is effected by a rack and pinion mechanism 194. The rack and pinion mechanism 194 comprises upper and lower racks 196, 198 mounted on extensible and retractable upper and lower actuators 200, 202. These racks engage with a splined section 204 of the second shaft 166. The racks 196, 198 are operated in unison such that when they are extended, the shaft 166 and tail chassis 150 connected to the shaft 166 will rotate in a clockwise direction as shown in Figure 20. Similarly, when the racks 196, 198 are retracted, the shaft 166 and chassis 150 will rotate in an anticlockwise direction as shown in Figure 19. Rotation of the tail plane around the axis 192 provides some pitch and yaw control for the aircraft. It will be appreciated that a wide range of tail plane orientations can be achieved by operating the actuator 182 and the actuators 196, 198 in unison.

An alternative tail plane 400 is shown in Figures 20A-20E. The tail plane 400 comprises a generally V-shaped chassis 402 which is pivotably mounted to a 5 tail crossbar 404. The tail crossbar 404 is pivotably mounted in a clevis joint 406 by a clevis pin 408. The clevis joint 406 is formed integrally with a tail plane mounting block 410 which is attached to a tail plane mounting shaft 412. The mounting shaft 412 is journal ed in an opening provided in a rear bulkhead 414 of the rear fuselage section 6 in order to mount the tail plane 400 to the aircraft.

10 As shown in Figure 2OA, a roll actuator 416 is provided to rotate the- mounting shaft 412, and therefore the tail plane 400, about a central axis 418. The roll actuator comprises a first idler shaft 420 which is drivingly connected to a servo 422 for rotation about a pivot axis 424. The first idler shaft 420 is connected by a pair of linkages 426 to a second idler shaft 428, which is mounted for rotation about

15 a pivot axis 430. Finally, the second idler shaft is connected by a second pair of linkages 432 to a third idler shaft 434 which is mounted to the tail plane mounting shaft 412. Rotational movement of the first idler shaft 420 by the servo 422 will therefore be transmitted to the mounting shaft 412 (and therefore to the tail plane 400) via the second and third idler shafts 428, 434 and the linkages 426, 432. '20 Rotation of the tail plane 400 about the axis 418 provides some pitch and yaw control for the aircraft.

As shown in Figures 2OA, 2OB and 2OD, the tail plane 400 may also be pivoted about a horizontal axis 436 in order to provide pitch control for the aircraft. In order to provide the pitch control, an elevator control mechanism 438 is provided.

25 The control mechanism 438 comprises a leadscrew 440 attached to a primary gear 442. A pair of servos 444 drive a pair of secondary gears 446 which are meshed with the primary gear 442 so as to rotate the leadscrew 440. The primary and secondary gears 442, 446 are mounted in a carrier housing 448 which is pivotably mounted on two support members 450, the support members 450 being mounted to

30 the tail crossbar 404. The leadscrew 440 is received in a captive leadscrew nut 452 which is mounted to a rod 454. The rod 454 is rotatably mounted to the tail plane chassis 402 via a pair of bearing assemblies 456. In order to adjust the pitch of the tail plane 400, the leadscrew 440 is rotated by the servos 444 via the primary and secondary gears 442, 446. The rotation of the leadscrew 440 causes the captive leadscrew nut 452 to move along the leadscrew 440. The movement of the captive leadscrew nut 452>alters the distance between the leadscrew nut 452 and the carrier housing 448, thus causing the tail plane chassis to rotate about the tail crossbar 404. The changing pitch angle of the tail plane chassis 402 is accommodated by the pivotal mounting of the carrier housing 448 to the support members 450 and the pivotal mounting of the rod 454 in the bearing assemblies 456. The tail plane 400 comprises a plurality of tail elements 460 pivotably mounted to the chassis 402. As shown in Figure 2OC, the tail plane 400 comprises an actuation assembly 462 for moving the tail elements 460 between open and closed configurations, as discussed previously with respect to Figures 11-13. The actuation assembly comprises a servo 464 which drives a first gear 466, the first gear 466 driving a second gear 468. Received in the second gear 468 is a leadscrew nut which engages a leadscrew 470 which is fixedly mounted to the tail plane chassis 402 along the axis 418. The servo 464 and the first and second gears 466, 468 are mounted to a carrier 472. By driving the leadscrew nut using the servo 464, the carrier 472 can be made to move along the leadscrew 470. A pair of pushrods 474 connect the carrier 472 to a pair of master tail elements 476 so as to translate the movement of the carrier 472 along the leadscrew 470 into rotational movement of the master tail elements 476 about the mounting axes 478. The tail elements 460 are interconnected by linkages (not shown) such that movement of the master tail elements 460 will result in movement of the remaining tail elements 460. The tail plane 400 may also be rotated laterally about the clevis pin 408 by an actuation mechanism 480 shown in Figures 2OD and 2OE. The actuation mechanism comprises a servo 482 (Figure 20A) which is mounted to the mounting block 410 via a servo plate 484. The servo 482 drives a shaft 486 which in turn drives a first idler shaft 488. The first idler shaft drives a second idler shaft 490 via a pair of linkages 492, the first and second idler shafts 488, 490 also being connected by a support bar 494 in order to prevent distortion of the actuation mechanism 480 when under load. The second idler shaft is connected to the tail crossbar 404 by a pushrod 496 in order to rotate the tail plane 400 about the clevis pin 408. By using the actuation mechanism 480 to move the tail plane 400 laterally, it is possible to shift the drag caused by the tail plane 400 when making a turn with the aircraft, i.e. the tail plane 400 can provide a rudder control. From the above discussion it will be seen that the main plane 8 and tail plane

14, 400 may be moved through a large number of configurations. However, the aircraft of the present invention also permits the relative positions of the main plane 8 and tail plane 14, 400 to be changed.

As shown in Figure 21, in a normal flight configuration, the main plane 8 which is attached to the front fuselage section 4 and the tail plane 14, 400 which is attached to the rear fuselage portion 6 lie generally parallel to one another with a vertical offset D. The housing 190 of the tail plane 14, 400 is pivotally attached to the rear fuselage section 6 about a first pivot axis 210 while the rear fuselage section 6 is pivotally attached to the front fuselage section 4 about a pivot axis 212. A rigid link 214 extends between a front mounting 216 provided on the front fuselage portion 4 and a rear mounting 218 provided on the tail plane housing 190. First and second actuators 220, 222 mounted to the forward fuselage section 4 are attached to the forward end of the rear fuselage section 6 through joints 224. This arrangement means that when, as shown in Figure 22, the upper actuator 222 is extended and the lower actuator 220 retracted, the rear fuselage section 6 rotates clockwise about the axis 212 relative to the forward fuselage section 4. However, due to the linkage 214, the tail plane rotates anticlockwise about the axis 210 such that the tail plane 14, 400 maintains its generally parallel relationship with the main plane 8 but at a much greater vertical offset D'. In effect, the rear fuselage portion 6 has been swung forwardly which is particular advantageous when the aircraft is landing. It will be appreciated that in either configuration, it is possible to rotate the tail plane 14, 400 about the axes 180, 192 and to open or close the tail elements.

As can be seen in Figures 21- and 22, an engine e.g. a ducted fan, turbofan or similar engine 230 may be accommodated within the second fuselage portion 6. During normal flight as illustrated in Figure 21 , the thrust generated by the engine 230 will be substantially rearward. However, in the landing configuration shown in Figure 22, the thrust will be directed rearwardly and downwardly, thereby providing a lift component to the aircraft. This lift component means that the aircraft may even be able to land at a speed below its normal stall speed.

As shown in Figures 23 and 24, the rear fuselage section 6 may accommodate an undercarriage 250. The undercarriage 250 comprises a pair of collapsible strut structures 252, one arranged on either side of the aircraft. Wheels 254 or other ground engaging members are provided on a cradle 256 attached to the lower most link of the mechanism 252. During normal flight, the centre of gravity and the centre of lift of the aircraft will be relatively far forward on the aircraft, ahead of the undercarriage. However, due to the pivoting of the rear fuselage section during landing, the centre of gravity of the aircraft will move back over the cradle to a position between the wheels 254. This will allow the aircraft to land without the use of a tail wheel or nose wheel which is normally required to stabilise the aircraft during landing and while on the ground.

The undercarriage 250 can be extended and retracted by virtue of the actuators 258 which also incorporate spring dampers to absorb landing loads. Links 260 arranged between the strut members 262 transfer the retracting or extending movement between the strut members 262.

It will be understood that a wide range of control movements may be made by suitable operation of the main plane 8 and tail plane 14, 400. The lift generated by the aircraft can be varied by varying the area of the main plane 8 and tail plane 14, 400 by extending and retracting the main plane 8 and tail plane 14, 400. For example, the aircraft can be trimmed in pitch by varying the area of the tail plane, for example. As the aircraft travels faster, the centre of lift on the wing will travel rearwardly, closer to the centre of gravity. A smaller corrective moment need therefore be applied by the tail plane which can be achieved by reducing its surface area.

Pitch and' yaw movements may be induced by movement of the tail plane 14, 400 about the orthogonal rotation axes 180, 192. Also, rolling and yawing control may be effected by differential extension and retraction of the wings 10, 12 of the aircraft. Yaw will be induced due to the difference resulting from differential drag on the wings, while rolling will be induced by the difference in lift generated by the wings. Operation of the outermost wing tip element will also have an effect in this regard, potentially changing the lift and drag on the relevant wing.

It will be understood that the above is a description of just a preferred embodiment of the invention that many variations are possible within the scope of the invention. For example, while the aircraft has been described as having an engine 230, it may of course operate as glider.

Claims

1. A wing structure for an aircraft, said structure comprising: an inboard spar pivotally mounted at an inboard end with respect to an aircraft fuselage; an outboard spar pivotally coupled at one end to an outboard end of the inboard spar; a plurality of elongate, overlapping wing elements each of which is attached at one end to a respective spar and extends rearwardly from said spar; means for rotating said inboard spar relative to said fuselage; means for rotating said outboard spar relative to said inboard spar; and means for maintaining said wing elements generally aligned throughout a range of rotational positions of said inboard and outboard spars.

2. A wing structure as claimed in claim 1 wherein the wing elements are maintained aligned by links extending between the wing elements.

3. A wing structure as claimed in claim 2 wherein the links are at least semirigid.

4. A wing structure as claimed in claim 1 , 2 or 3 wherein the links extend generally parallel to the respective spars so as to form a parallelogram type linkage therewith.

5. A wing structure as claimed in any preceding claim wherein as the spars rotate, the wing elements will maintain a generally parallel orientation.

6. A wing structure as claimed in any preceding claim wherein the wing elements are attached to the respective spars through clevis-type joints.

7. A wing structure as claimed in claim 6 wherein the clevis prevents or at least limits flapping of the wing elements in use.

8. A wing structure as claimed in any preceding claim wherein the inboard spar member is mounted to the fuselage in a manner which allows it to rotate relative thereto about a generally vertical axis.

9. A wing structure as claimed in any preceding claim comprising an actuator for rotating the inboard spar.

10. A wing structure as claimed in claim 9 comprising a pair of wing structures and separate actuators for the inboard spar of each wing structure.

1 1. A wing structure as claimed in claim 9 comprising a pair of wing structures and a common actuator connected to the respective inboard spars through respective linkages.

12. A wing structure as claimed in any preceding claim wherein the rotation of the outboard spar relative to the inboard spar is effected by a second actuator.

13. A wing structure as claimed in any preceding claim wherein the second actuator is provided between the fuselage and the outboard spar.

14. A wing structure as claimed in claim 13 wherein the second actuator comprises a generally fixed length link extending between the fuselage and the outboard spar.

15. A wing structure as claimed in claim 14 wherein the inboard end of the second actuator is mounted to the fuselage in a translationally fixed position.

16. A wing structure as claimed in any of claims 1 to 14 wherein the inboard end of the second actuator is movable or translatable to move the outboard spar.

17. A wing structure as claimed in claim 16 wherein the inboard end of the second actuator is mounted to a bell crank or other mechanism which is operable to move the position of the inboard end of the second actuator.

18. A wing structure as claimed in claim 17 comprising two wing structures and wherein separately operable second actuators are provided for the outboard spar of each wing.

19. A wing structure as claimed in claim 17, comprising two wing structures and wherein commonly operable second actuators are provided for the outboard spar of each wing..

20. A wing structure as claimed in claim 19 wherein respective linkage mechanisms are driven by a common drive., for example a rack and pinion drive.

21. A wing structure as claimed in claim 19 or 20 wherein the respective second actuators are operable such that as one actuator operates to extend the outboard spar of one wing, the other operates to retract the outboard spar of the other wing.

22. A wing structure as claimed in claim 12 comprising an extensible and retractable second actuator mounted between the inboard and outboard spars.

23. A wing structure as claimed in any of claims 12 to 21 wherein the or each second actuator is operable independently of said first actuator. ^l

24. A wing structure as claimed in any preceding claim comprising a tip spar, pivotally connected, directly or indirectly to outboard end of the outboard spar.

25. A wing structure as claimed in claim 24 comprising a plurality of wing tip elements attached pivotally to and extending rearwardly from said-wing tip spar, said wing tip elements being connected by links.

26. A wing structure as claimed in claim 25 comprising a third actuator to move the tip spar relative to the outboard spar.

27. A wing structure as claimed in claim 26 wherein the third actuator comprises an element which is attached at an inboard end to the inboard spar and which cooperates with the tip spar at its outboard end.

28. A wing structure as claimed in claim 27 wherein the actuator element cooperates with an actuating member provided on the tip spar such that as the inboard spar is rotated, the tip actuator pushes and/or pulls against the actuating member on the tip spar, causing it to pivot relative to the outboard spar.

29. A wing structure as claimed in claim 29 wherein there is a lost motion mechanism provided between the element and the tip spar.

30. A wing structure as claimed in claim 29 wherein the outboard end of the third actuator comprises an element for example a pin arranged to slide within a slot provided on the tip spar.

31. A wing structure as claimed in any of claims 26 to 30 wherein the third actuator is arranged so as to fail when a predetermined load on the third actuator is exceeded.

32. A wing structure as claimed in any of claims 26 to 31 comprising a further wing element support pivotally attached to the tip spar so as to be selectively pivoted forwardly relative to the wing about a generally vertical axis, said element support being actuated by a fourth actuator.

33. A wing structure as claimed in claim 32 wherein the fourth actuator comprises a tension member which acts to pivot the control element about the axis.

34. A wing structure as claimed in claim 32 or 33 wherein the fourth actuator comprises an actuator mounted between the tip spar and the control element for movement of the control element relative to the tip spar.

35 A wing structure as claimed in any preceding claim wherein the wing elements comprise a shaft for attachment to a spar and a lifting body mounted to the shaft, the lifting body being asymmetrical about the shaft axis, having a relatively narrow leading edge portion and a relatively wider trailing edge portion.

36. A wing structure as claimed in claim 35 wherein the leading edge portion of one wing element overlaps the trailing edge portion of the adjacent wing element.

37. A wing structure as claimed in claim 35 or 36 wherein a trailing edge region of the lifting body is relatively flexible compared to a leading edge region of the lifting body.

38. An aircraft wing element comprising an elongate shaft and a lifting body mounted to said shaft; said lifting body having a leading edge portion and a trailing edge portion, wherein said leading edge portion is narrower than said trailing edge portion and said trailing edge portion has a trailing edge region which is relatively flexible compared to the leading edge portion.

39. A wing structure or wing element as claimed in claim 37 or 38 wherein the leading and trailing edge portions of the lifting body are formed as separate bodies and suitably attached to the shaft or are formed together, integrally with the shaft.

40. A wing structure or wing element as claimed in any of claims 35 to 39 wherein the lifting body is of a lightweight, relatively rigid material such as a foam.

41. A wing structure or wing element as claimed in any of claims 35 to 40 wherein the shaft or its attachment to the spar is flexible such that it allows the lifting body to twist about its axis to a limited degree.

42. A wing structure or wing element as claimed in any of claims 35 to 41 wherein one or more intermediate wing elements are provided between adjacent wing elements so as to substantially cover any gaps between the adjacent wing elements, the intermediate wing elements preferably being shorter in length than the main wing elements and extending within the region proximate the wing spars.

43. A wing structure or wing element as claimed in any of claims 35 to 42 comprising one or more shrouds over the wings.

44. A wing structure or wing element as claimed in claim 43 wherein the shrouds are attached to the respective spars and are configured such that they may overlap one another.

45. A wing structure or wing element as claimed in claim 43 or 44 wherein the internal space of the shroud is filled with a filler material.

46. A wing structure or wing element as claimed in claims 45 wherein the filler material is resiliency deformable.

47. A wing having relatively rotatable inboard and outboard spars, one or more shrouds covering said spars and a resiliently deformable filler material arranged within said shroud(s).

48. Apparatus as claimed in any preceding claim wherein the leading edge of the wing is configured such that as the wing extends, the leading edge of the wing drops to increase the camber on the inboard part of the wing.

49. An aircraft having a wing which is pivotable about a generally vertical axis, the wing being configured such that as an inboard portion of the wing pivots rearwardly, the leading edge of the wing moves so as to increase the camber of an inboard section of the wing.

50. Apparatus as claimed in claim 48 or 49 comprising a control member extending along the leading edge region of the wing from the fuselage to an outboard portion of the wing, the attachment of the control member to the fuselage being positioned such that during rotation of the wing, the control member becomes taut and pulls the leading edge down, thereby increasing its camber.

51. Apparatus as claimed in any preceding claim wherein the wing is configured such that in the event of a power failure, it will move to a maximum surface area configuration, thereby facilitating landing.

52. Apparatus as claimed in any preceding claim wherein the wing is mounted to a fuselage in such a manner as to accommodate variations in an angle of dihedral of the wing occurring due to variations in an operative configuration of the wing.

53. Apparatus as claimed in claim 52 wherein the wing mounting accommodates rotation of the wing about a generally vertical axis and about a generally longitudinal axis of the aircraft.

54. An aircraft comprising a wing mounted to a fuselage, said mounting being such as to permit rotation of the wing about a generally vertical axis, and also about an axis extending in a longitudinal direction of the aircraft.

55. Apparatus as claimed in claim 53 or 54 wherein the or an inboard spar is mounted on a support which is rotatably mounted on the fuselage.

56. Apparatus as claimed in claim 55 wherein the support is mounted for rotation about a generally vertical axis and the wing is mounted on the support such that it - may rotate about an axis extending in a direction along the length of the fuselage.

57. Apparatus as claimed in any of claims 52 to 56 wherein the change in dihedral is accommodated by spring means.

58. Apparatus as claimed in claim 57 comprising spring means which counter displacement of the wing in both upward and downward directions.

59. Apparatus as claimed in claim 58 wherein the spring force resisting upward motion of the wing is greater than that resisting downward motion of the wing.

60. Apparatus as claimed in claim 57, 58 or 59 wherein said spring means comprises compression or tension springs, a compressible elastic material or hydraulic springs.

61. Apparatus as claimed in any preceding claim further comprising a wing mounting comprising one or more dampers for damping vertical movements of the wing.

62. An aircraft having a wing mounted to an aircraft fuselage for vertical movement relative thereto, said mounting comprising damping means fordamping such vertical movements.

63. Apparatus as claimed in claim 61 or 62 wherein the mounting comprises one or more spring dampers.

64. Apparatus as claimed in any of claims 52 to 63 comprising means for limiting the upward pivoting of the wings relative to the fuselage.

65. Apparatus as claimed in claim 64 wherein said means comprises a tether member extending between the wing and the fuselage.

66. An aircraft having a pair of wings pivotally mounted to a fuselage, and comprising tether means attached between the wings and the fuselage to limit rotation of the wings relative to the fuselage.

67. Apparatus as claimed in claim 65 or 66 wherein tethers extend between the wings and a keel provided at or adjacent a lower part of the fuselage.

68. Apparatus as claimed in claim 66 or 67 wherein the tethers are relatively inextensible

69. An aircraft having a wing mounted to a fuselage, said mounting being such as to permit rotation of the wing about a generally vertical axis, said aircraft further comprising means for automatically changing a dihedral angle of said wing in response to said rotation of said wing.

70. An aircraft tail plane comprising a tail chassis; a plurality of tail elements pivotally attached to said chassis and at least partially overlapping one another; and actuating means for rotating said tail elements relative to each other so as to vary the degree of overlap between the tail elements such that the surface area of the tail can be changed.

71. A tailplane as claimed in claim 70 wherein the tail chassis is generally V- shaped in section having diverging limbs, tail elements being attached to and extending rearwardly from these limbs.

72. A tailplane as claimed in claim 71 wherein the tail elements are symmetrically arranged around a fixed central tail element.

73. A tailplane as claimed in claim 72 wherein the central tail element is relatively rigid and overlies the adjacent tail elements of either side.

74. A tailplane as claimed in any of claims 70 to 73 comprising a common actuator for moving a plurality of pivotally mounted tail elements simultaneously and a linkage extending between one or more of the tail elements and the common actuator.

75. A tailplane as claimed in claim 76 wherein the actuator comprises a leadscrew, and a runner movable along the leadscrew and coupled to a tail element.

79. A tailplane as claimed in any of claims 70 to 78 wherein the tail elements are resiliency biased towards a failsafe configuration.

80. A tailplane as claimed in any of claims 70 to 79 wherein the chassis is mounted for rotation about a pitch axis such that the tail can be pitched up and down to change its angle of attack.

81. A tailplane as claimed in claim 80 wherein the tail is rotated in pitch about its leading edge.

82. An aircraft having a tail plane which is rotatable in pitch about its leading edge.

83. A tailplane as claimed in claim 80 or 81 wherein the chassis is mounted for rotation on a support which is rotatable about a roll axis.

84. A tailplane as claimed in claim 83 comprising a pitch actuator mounted to the support.

85. A tailplane as claimed' in claim 84 wherein the actuator comprises a leadscrew mounted to the support and a runner mounted on the leadscrew and attached to the chassis.

86. A tailplane as claimed in any of claims 83 to 85 wherein the support member comprises a shaft mounted for rotation about the roll axis.

87. A tailplane as claimed in claim 86 comprising an actuator for rotating the support member about the roll axis.

88. A tailplane as claimed in claim 87 comprising a pair of actuators provided acting on opposed parts of the shaft so as to balance the rotational forces acting thereon.

89. A tailplane as claimed in any of claims 70 to 88 wherein the chassis is rotatable about a yaw axis.

90. A tailplane as claimed in any of claims 70 to 89 wherein the tail plane is pivotable as a unit with respect to an aircraft fuselage.

91. An aircraft comprising a main plane, a tail plane and fuselage, said main plane and said tail plane being translatable one relative to the other so as to change their relative positions.

92. An aircraft as claimed in claim 91 wherein the fuselage comprises a first section to which is mounted the main plane and a second section which is moveable with respect to said first fuselage section so as to change the relative positions of said main plane and said tail plane.

93. An aircraft comprising a main plane, a tail plane, a first fuselage-section and a section fuselage section, said main plane being mounted to said first fuselage section and said tail plane being mounted to said second fuselage section, and means for moving said second fuselage section with respect to said first fuselage section so as to change the relative positions of said main plane and said tail plane.

94. An aircraft as claimed in claim 91 or 92 wherein the first and second fuselage sections are pivotally connected together, and comprising a first actuator to effect pivotal movement between the two.

95. An aircraft as claimed in any of claims 91 to 94 wherein the relative planar relationship between the main plane and the tail plane remains generally the same in the absence of any independent control movement of the tail plane.

96. An aircraft as claimed in claim 95 wherein the main plane and the tail plane remain generally parallel in both configurations.

97. An aircraft as claimed in claim 96 comprising a further actuator for rotating the main plane or the tail plane relative to its respective fuselage section so as to maintain the desired positional relationship between the main plane and the tail plane.

98. An aircraft as claimed in claim 97 wherein an actuator is provided to rotate the tail plane relative to the second fuselage section.

99. An aircraft as claimed in claim 98 wherein the actuator is provided on the second fuselage section and is operable in parallel with the first actuator in order to obtain the necessary relative planar configuration between the main plane and the tail plane.

100. An aircraft as claimed in claim 98 wherein the second actuator comprises a fixed length link extending between the first fuselage section and the tail section such that rotation of the second fuselage section relative to the first fuselage section automatically results in rotation of the tail section relative to the second fuselage section.

101. An aircraft having a first, forward fuselage section and a second, rearward fuselage section, said sections being rotatable relative to one another and further comprising an undercarriage mounted to said second fuselage section.

102. An aircraft as claimed in claim 100 wherein the undercarriage comprises a collapsible strut which is retractable into the fuselage generally axially of the aircraft.

103. An aircraft as claimed in claim 102 wherein the strut comprises a plurality of pivotally interconnected and relatively collapsible links.

104. An aircraft as claimed in any of claims 100 to 104 wherein the undercarriage comprises ground engaging members spaced apart longitudinally on the end of the strut.

105. An aircraft as claimed in any of claims 91 to 104 wherein the front fuselage section comprises a cockpit for one or more passengers.

106. An aircraft as claimed in any of claims 91 to 105 comprising a an engine in the second fuselage portion.

107. An aircraft having a front fuselage section and a rear fuselage section, said rear fuselage section being rotatable relative to said front fuselage section, said aircraft further comprising an engine provided in said rear section whereby the line of action of the thrust of the engine relative to the front fuselage section can be changed.

108. An aircraft wing having a control element at the tip of the wing mounted for pivotal movement about a generally vertical axis.

109. An aircraft comprising a pair of wings; means for extending or retracting ^' said wings or one or more parts thereof symmetrically; and means for extending or retracting said wings or one or more parts thereof asymmetrically.

1 10. An aircraft as claimed in claim 109 wherein each wing comprising an inboard spar pivotally mounted at an inboard end with respect to an aircraft fuselage and an outboard spar pivotally coupled at one end to an outboard end of the inboard spar; and wherein the outboard spar of each wing can be moved independently of the movement of said inboard spar to effect the asymmetrical extension or retraction of the wings.