AIRCRAFT WING STRUCTURE

Abstract

An aircraft has a wing providing the main lifting surface for the aircraft. The wing has a structure supporting an aero-dynamic surface, and the wing has a weight, the wing structure being unable to support its own weight when the aircraft is stationary and under a load of 1 g so as to cause structural failure of the wing.

Claims

1. An aircraft comprising a wing providing the main lifting surface for the aircraft, the wing having a structure supporting an aerodynamic surface, wherein the wing has a weight, the wing structure being unable to support its own weight when the aircraft is stationary and under a load of 1 g so as to cause structural failure of the wing.

2. An aircraft according to claim 1, wherein the wing is adapted to generate lift when the aircraft is moving relative to the surrounding air to achieve sustained flight, and the wing structure is able to support its own weight under a load of 1 g during flight.

3. An aircraft according to claim 1, wherein the wing structure is unable to support its own weight when the aircraft is stationary and under a load of 0.5 g so as to cause structural failure of the wing.

4. An aircraft according to claim 1, wherein the aircraft has powered propulsion.

5. An aircraft according to claim 4, wherein the aircraft comprises at least one propeller powered by a motor.

6. An aircraft according to claim 4, wherein the aircraft is solar powered, comprising solar energy collectors and batteries for storing electrical energy.

7. An aircraft according to claim 1, wherein the aircraft carries a payload and the total weight of the aircraft is comprised of greater than 30% payload, preferably greater than 40% payload and more preferably greater than 50% payload.

8. An aircraft according to claim 1, wherein the aircraft is an unmanned vehicle.

9. An aircraft according to claim 1, wherein the aircraft excluding any payload has a mass of between 30 kg to 150 kg.

10. An aircraft according to claim 1, wherein the aircraft wing has a span of from 20 to 60 metres.

11. An aircraft according to claim 1, wherein the aircraft has no landing gear.

12. An aircraft according to claim 1, wherein the structure of the wing comprises at least one space frame and at least one cover supported by the space frame, wherein the space frame has one or more structural members, the structural members including a structural foam material.

13. An aircraft according to claim 12, wherein the cover is pre-stressed and the pre-stressed cover has the aerodynamic surface of the wing.

14. An aircraft according to claim 12, wherein the space frame comprises one or more chordwise ribs and one or more longitudinal spars.

15. An aircraft according to claim 12, wherein the structural member(s) consist of a structural foam material.

16. An aircraft according to claim 12, wherein the structural foam material is a cellular core foam of for example polystyrene, polycarbonate, polyvinyl chloride, polypropylene, acrylonitrile-butadiene-styrene or a polymethacrylimide (PMI) foam such as Rohacell?.

17. An aircraft according to claim 1, wherein the structural failure of the wing includes plastic deformation of the wing.

18. An aircraft according to claim 6, wherein the aircraft is exclusively solar powered.

19. A method of launching an aircraft, the aircraft comprising a wing providing the main lifting surface for the aircraft, the wing having a structure supporting an aerodynamic surface, wherein the wing has a weight, the wing structure being unable to support its own weight when the aircraft is stationary and under a load of 1 g so as to cause structural failure of the wing the method comprising lifting the aircraft from a substantially ground level location to an elevated altitude by a lighter than air carrier.

20. A method of launching an aircraft according to claim 19, wherein the aircraft is attached to the lighter than air carrier via one or more tethers during ascent from the substantially ground level location to the elevated altitude, the tethers being severed at altitude to launch the aircraft.

21. A method of launching an aircraft according to claim 19, wherein the elevated altitude is from 18,000 to 30,000 metres.

22. An aircraft according to claim 1, wherein the wing structure is configured to be supported prior to flight and being unable to support its own weight when the aircraft is stationary and under a load of 1 g so as to cause structural failure of the wing if the wing is unsupported prior to flight.

23. A method of launching an aircraft according to claim 19, wherein the wing structure is configured to be supported prior to flight and being unable to support its own weight when the aircraft is stationary and under a load of 1 g so as to cause structural failure of the wing if the wing is unsupported prior to flight.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

[0024] Embodiments of the invention will now be described with reference to the accompanying drawings, in which:

[0025] FIG. 1 is a perspective view of an unmanned aerial vehicle in flight,

[0026] FIG. 2 is a perspective view of a stationary UAV on the ground, with unsupported wings, showing how one wing has failed,

[0027] FIG. 3a is a perspective partially exploded view of the wing structure of the UAV without the aerodynamic surface,

[0028] FIG. 3b is a perspective view of the wing structure of the UAV of FIG. 3a with the upper and lower covers and the aerodynamic surface assembled,

[0029] FIG. 4 is a perspective view of the UAV at altitude, having just been released from the lighter than air carrier, and

[0030] FIG. 5A-5G are perspective views of an exemplary launch and release method of a UAV.

DETAILED DESCRIPTION OF EMBODIMENT(S)

[0031] In an embodiment, the aircraft is a UAV 1 having two wings 2, a fuselage 4, and a tailplane 6, as shown in FIG. 1.

[0032] In this embodiment the fuselage 4 is a minimal structure, comprising simply a lightweight tube, with the wings 2 and tailplane 6 attached to the tube. The tube is of carbon fibre construction, having a diameter in the range of 60 to 120 mm and a wall section of 0.5 mm. In alternative embodiments, the fuselage may be constructed of any lightweight material, for example wood, plastic or fibre reinforced composite, and may be hollow or solid, and of any shape suitable for having wings and tailplane attached. The shape and dimensions of the fuselage may vary along the length of the fuselage, for example to provide weight balance, and may be elliptical or tapered. The nose 8 of the fuselage extends forwards of the wings and acts to counter balance the weight of the tailplane. The nose 8 also provides optional payload storage capacity.

[0033] The tailplane 6 has cruciform vertical and horizontal stabilising surfaces attached to the fuselage 4. The trailing portion of the stabiliser has an active movable rudder 10 located at the upper and lower portion of the vertical stabilising surface. An actuator controls the rudder 10, the actuator being located in the tailplane 6.

[0034] The wings 2 are elongate in a spanwise direction with a total wingspan of around 20 to 60 metres, extending either side of the fuselage 4. The wing may be straight or tapered in the outboard direction, and the wings may be horizontal or have a dihedral or an anhedral angle from the point the wing meets the fuselage, or from any point along the wing.

[0035] Each of the wings 2 carry a motor driven propeller 12 which may be powered by rechargeable batteries, or the batteries may be recharged during flight via solar energy collecting cells (not shown) located on the external surface of the aircraft. Each propeller is lightweight, in an embodiment the propellers each weigh less than one kilogram and are greater than 2 metres in length. FIG. 1 shows each wing carrying a single propeller, however in alternative embodiments multiple propellers may be provided on each wing. The propellers are shaped for high altitude, low speed flight. The payload of the vehicle is also carried mainly within the wing structure.

[0036] FIG. 1 shows the UAV 1 in sustained flight, with the wings 2 generating lift and the wing structure able to support its own weight. FIG. 2 shows the UAV 1 stationary and on the ground 14. In FIG. 2 the UAV 1 has not been supported on the ground and failure of the wings 2 has therefore occurred at least at point 16 along one of the wings 2. Failure may take the form of a fracture or complete break in the wing structure, as shown in FIG. 2, or may be excessive deformation such that the wing structure has plastically rather than elastically deformed and is therefore unable to support flight of the aircraft. Failure occurs as a result of the wing 2 being unable to support its own weight at the Earth's surface, i.e. under a loading of 1 g. In fact, the wings 2 are so fragile that each wing may be unable to support its own weight under a loading of less than 1 g, and will break when subjected to a load of, for example, 0.5 g.

[0037] In an embodiment, each wing 2 comprises an aerofoil 20 as shown in the exploded perspective view of FIG. 3a. The aerofoil 20 comprises a space frame 22 having a plurality of ribs 23 and spars 24, a cover 10 including an upper cover (skin) 26, a lower cover (skin) not visible in FIG. 3a, a leading edge assembly 28, a trailing edge assembly 30, and also includes an aerodynamic surface membrane 32 (shown in FIG. 3b). The aerofoil 20 has a cambered, low speed profile with a sharp leading edge radius.

[0038] The ribs 23 extend chordwise across the aerofoil 20, and are spaced equidistantly apart in a spanwise direction. Each rib 23 is of similar overall shape and dimension. Each spar 24 extends spanwise along the length of the aerofoil 20 and comprises an upper spar section and a lower spar section. The profile of the upper and lower spar sections varies chordwise across the aerofoil 22 according to the shape of the aerofoil 22, the upper and lower spar sections having a larger vertical cross-section at the quarter chord position. The space frame 22 is assembled by inserting slots in each lower spar into slots along the lower edge of each rib 23, and then repeating the process for all subsequent ribs 23. Similarly, slots in each upper spar insert into slots along the upper edge of each rib 23.

[0039] The ribs 23 and upper and lower spar sections are formed by cutting the rib or spar profile from a sheet of structural foam. The material thickness is of the order of 4 mm, but may be thinner for example 2 mm or 3 mm, or could also be thicker than 4mm. In this embodiment, the ribs and spar sections are cut from structural foam material of the same thickness, alternative embodiments may have specific ribs and/or spars of a differing material thickness. The structural foam used is Rohacell? 31 IG-F although there are a number of structural foam types suitable for use, as long as the key features of lightweight combined with rigidity are present. Rohacell? 31 IG-F is the lightest grade of structural foam currently available, with the finest cell structure. Other products with an equivalent density have a coarser cell structure and surfaces are therefore not as suitable for bonding. Heavier grades of structural foam would provide an increased stiffness, but at an increased weight.

[0040] FIG. 3b shows the assembled aerofoil including the aerodynamic surface membrane 90. The membrane is pre-stressed and then stretched over the aerofoil and taped on to the aerofoil. This enables the aerofoil to behave as a monocoque structure, i.e. the membrane becomes a structural component also and increases the loads that can be supported by the aerofoil. The membrane may be of any lightweight material that is dimensionally stable both with temperature and under UV light conditions, whilst providing good tensile strength to weight ratio. In this embodiment, a polyimide film such as Kapton? is used, with a thickness of 12.5 micron. Alternatives include Mylar film.

[0041] With the upper 28 and lower 26 covers assembled on to the space frame 22, the aerofoil structure comprises multiple hollow cells 31 formed by adjacent ribs, spars and upper 28 and lower 26 covers. The payload is carried inside these hollow sections 31, typically along successive hollow sections formed at the quarter chord position, where the hollow sections have the largest dimensions. Optional reinforcement the area of the aerofoil carrying the payload may be provided, for example one or more additional reinforcing spars can be located either side of the hollow cells 31 carrying the payload. Additionally, the lower face of the hollow cells 31 carrying payload can also be reinforced and shaped so as to facilitate location of items or cables within particular hollow cells 31.

[0042] Payload is distributed within the wing so as to balance the centre of mass of the wing and payload with the centre of lift of the wing within a predetermined range for the defined flying characteristics.

[0043] The UAV 1 is lifted to altitude by a balloon 40 in a wingtip up configuration and then reoriented at altitude in readiness for release. The UAV 1 is attached to the balloon during the lift phase by one or more tethers 42. FIG. 4 shows the UAV 1 having reached its launch altitude and the tethers 42 attaching the UAV 1 to the balloon 40 having been released. The UAV 1 is thereby released into its flight mode.

[0044] FIG. 5 shows the method of launching an aircraft of the invention. The design detail of the aircraft in FIG. 5 differs in a number of ways when compared to the UAV shown in FIGS. 1, 2 and 4. For example, the aircraft in FIG. 5 has a larger fuselage and a T-tail design of tailplane, with the horizontal stabiliser located at the end of the vertical stabiliser. The propeller design and location on the wing also differs. However, the method of launching is equally applicable to both aircraft designs, and is not intended to be limited to an aircraft of particular proportion or detail design.

[0045] FIG. 5 shows an unmanned aerial vehicle 1 attached to a lighter-than-air balloon 40 acting as the carrier at a ground level location at stage A. The UAV 1 is attached to the balloon by three tethers 42, one attached to a wing-tip and two attached to the fuselage. The method is similar to that described with reference to WO 2014/013268, the contents of which are incorporated herein by reference.

[0046] During launch, the balloon 40 lifts the UAV 1 which enters a wing-tip-up orientation, as shown in stage B, achieved by pulling on the tether attached to the wing-tip.

[0047] The balloon 40 then begins the ascent, carrying the UAV 1 suspended below in a wing-tip-up orientation, as shown in stage C.

[0048] Once at the launch altitude the UAV 1 is moved back into a horizontal attitude by letting out the tether 42 attached to the wing-tip, as shown in stage D.

[0049] The tether 42 attached to the wing-tip is then severed, together with one of the tethers attached to the fuselage, causing the UAV 1 to enter a nose-down configuration, as shown in stage E and F.

[0050] Finally the last tether 42 is severed, releasing the UAV to begin its descent, as shown in stage G. The motors are powered and the UAV increases flight velocity until the descent is controlled and the UAV 1 is capable of independent flight.

[0051] The UAV 1 may have a flight duration of days, weeks, months or possibly even years. At the end of the mission, the UAV 1 is required to land. Since the UAV 1 does not possess any landing gear structure, landing on a runway may not be possible. The UAV 1 may be designed to crash land, be caught in a net or land on a moving platform in the vicinity of the landing location for example.

[0052] Although the invention has been described above with reference to one or more preferred embodiments, it will be appreciated that various changes or modifications may be made without departing from the scope of the invention as defined in the appended claims.