An aircraft

Abstract

An aircraft 10 comprising: a starboard swept wing 14 and a port swept wing, each swept wing having an inner wing section 32 comprising an integrated payload fuselage (20 not shown) and an outer wing section 30 for extending the wingspan, in which each outer wing section has a thickness less than a thickness of the corresponding inner wing section; a front spar 36 close to a leading edge 18 of each wing, in at least the outer wing section, and a rear spar 38 close to a trailing edge 22 of each wing, for carrying loads along each wing; a central region 12 where the starboard swept wing meets the port swept wing; a starboard transition region where the starboard inner wing section meets the starboard outer wing section; a port transition region where the port inner wing section meets the port outer wing section; and a load-bearing structural element 40 disposed across each wing from the front spar to the rear spar in the transition region for carrying loads between the front spar and the rear spar of each wing.

Description

The present invention relates to an aircraft. BACKGROUND TO THE INVENTION A V-shaped aircraft has been previously devised by the inventor, which has been disclosed in his patent applications (including DE102014201040A1, N2034242 and GB2319730.4) and in a number of published papers on the concept. Proof-of concept prototypes have been made. V-shaped aircraft can have fuselage sections integrated into the wings. In all aircraft, the structure of the fuselage and the structure of the wing are subject to different design constraints. The fuselage is generally pressurised, and as such needs to be shaped to maximise strength and avoid adding undue thickness and weight, for instance with ovular, cylindrical, or hemispherical shapes. Wings need to be shaped like an aerofoil in order to maximise lift and minimise drag and are generally thin compared to the fuselage. In conventional aircraft, there is no tension between these two constraints, as the fuselage and wings are different components of the aircraft. V-shaped aircraft can have inner wing sections, and outer wing sections. These can be highly-swept, and the inner wing can have a larger sweep angle than the outer wing. The inner wing section can house the fuselage. The fuselage in the inner wing must be thick enough to hold sufficient cargo and passengers, whilst being thin enough to be aerodynamically optimised. The outer wing section is shaped like a conventional wing and is relatively thin for maximum lift and optimal aerodynamics. A transition region exists between the inner wing and outer wing. Since each section is subject to different design constraints, this intermediary region presents a structural optimisation challenge. Suitable optimisation of this region is of paramount importance, as it will dictate to a large extent the overall fuel burn per passenger and cargo of the aircraft. Furthermore, this transition region is a region of load-bearing importance. The outer wing section generates significant upwards lift, whilst the inner section is loaded with cargo and passengers and experiences a net downwards gravitational force, albeit that this is less than the weight of the passengers and cargo due to the lift generated by the inner wing. A significant amount of the loads on the outer wing are carried by an outer wing front spar, while heavy components such as the landing gear and the engines attach to an inner wing rear spar. Efficient force transfer between these spars is needed to avoid undue reinforcement and the subsequent thickness and additional weight that would create. Previously, various design iterations have been performed on V-shaped aircraft, which are described along with their shortcomings below. Optimisations have been considered individually on the inner wing and the outer wing, but no coherent design for the transition wing region of such an aircraft exists. In one proposed concept, the pressurised payload fuselage terminates long before it reaches the transition region, which results in limited volume for passengers and cargo (Figure 7). In an alternative concept, the pressurised region extends all the way to the transition region, which results in a fuselage that is too thick in the transition region and inhibits aerodynamic performance (Figure 8). The load paths of the concept shown in Figure 8 are shown as thickened spars in Figure 9. A partly optimised exterior design has been developed (Figure 10) but this does not present any solutions as to how to create an optimised interior structure. In this design, the exterior design has been prioritised. Therefore, even though the pressurised payload fuselage extends to the transition region, its height is compromised by the need for optimal aerodynamics and therefore inadequate internal volume is provided. It is an object of the present invention to reduce or substantially obviate the aforementioned problems. STATEMENT OF INVENTION According to the present invention there is provided an aircraft comprising: a starboard swept wing and a port swept wing, each swept wing having an inner wing section comprising an integrated payload fuselage and an outer wing section for extending the wingspan, in which each outer wing section has a thickness less than a thickness of the corresponding inner wing section; a front spar close to a leading edge of each wing, in at least the outer wing section, and a rear spar close to the trailing edge of each wing, for carrying loads along each wing; a central region where the starboard swept wing meets the port swept wing; a starboard transition region where the starboard inner wing section meets the starboard outer wing section; a port transition region where the port inner wing section meets the port outer wing section; and a load-bearing structural element disposed across each wing from the front spar to the rear spar in the transition region for carrying loads between the front spar and the rear spar of each wing. Each wing comprises a front spar and a rear spar for carrying loads along each wing. This allows transmission of loads in substantially longitudinal directions along each wing. For example, along an inner wing portion and then along an outer wing portion. The load-bearing structural element provides an efficient load path that can transmit loads, for example those generated by lift forces or weight forces, across the aircraft. That is, to transmit forces from the front spar, close to the leading edge of the wing, to the rear spar, close to the trailing edge of the wing. More efficient load transfer can be expected to lead to a lower weight of the aircraft. Advantageously, the load-bearing structural element may help to reduce the loads transferred into the front section of the inner wing. In the inner wing, there may not be a definite, solid front spar, but instead reinforcements to a ceiling and floor of the aircraft, and / or tension struts between the ceiling and the floor. Moving further towards a tip of the wing, as the absolute wing thickness decreases, these reinforcements may become more apparent, transitioning to a more conventional spar-like structure, in particular in the transition wing region, where no windows are envisioned anymore along the leading edge, and where the region begins to store cargo. Further aft in the transition wing region close to the outer wing region (where thickness is even further decreased), it is likely that the structure looks like a conventional front spar. The integrated payload fuselage may have a bulkhead in each transition region. The bulkhead may not be a traditional curved bulkhead. The bulkhead may instead be flat. Each bulkhead may have a first portion and a second portion. The first portion may be streamwise and will henceforth be known as the streamwise portion. The second portion may be parallel to a direction between a vertex in the front spar and a vertex in the rear spar. As such, it may be at an angle to the first portion of the bulkhead. It may in one embodiment be at a right angle to the first portion of the bulkhead. A portion of the load-bearing structural element may form part of or be attached to the second portion of the bulkhead of the payload fuselage. In alternative embodiments, the loadbearing structural element could be fixed to the payload fuselage rather than forming a part of the payload fuselage bulkhead. In the case of pressurised payload fuselages, the load-bearing structural element can therefore additionally support some of the pressure loads of the aircraft. The load-bearing structural element can be used to carry bending loads and pressurization loads at the same time, which reduces the number of parts required, and results in a lighter aircraft. The load-bearing structural element controls how the loads are introduced into the pressurised region, which may lead to a lower weight and may also improve aeroelasticity of the aircraft. Thickness referred to herein means an absolute thickness not a thickness-to-chord ratio. An oval retention parameter may be defined on each wing. As discussed previously, the outer wing shape may have a thin and aerodynamic outer wing profile whilst the inner wing may have a thicker ovular cross section for supporting the internal pressurisation loads. The oval retention parameter is defined as a streamwise line at the transition region between these two design considerations. Alternatively put, the oval retention parameter is at the position in the transition wing section the thick oval shape ends. The streamwise portion of the bulkhead may be disposed near the oval retention parameter. However, the streamwise portion may be further away from the centre of the aircraft than the oval retention parameter. In alternative embodiments, the streamwise portion of the bulkhead may be disposed in approximately the same position as the oval retention parameter. The load-bearing structural element may be a beam. The load-bearing structural element may be an I beam or an H beam. Alternatively, it may be a truss structure. The load-bearing structural element may be of varying height to match the height of the wing. The load-bearing structural element may connect a vertex in the front spar with a vertex in the rear spar. There may also be a streamwise portion of the bulkhead at an end of the payload fuselage. The streamwise portion of the bulkhead may join with the load-bearing structural element. That is to say, the streamwise portion of the bulkhead forms one edge of the payload fuselage bulkhead, and the load bearing spar forms another edge. The existence of two bulkhead edges at the rear of the payload fuselage disposed at an angle to each other, allows the payload fuselage to protrude into the centre of the transition region. This in turn allows edges of the inner wing to be thinner as the portions do not have to accommodate the full height of the payload fuselage. This enables a more aerodynamic profile in the transition region. The payload fuselage may be pressurised. The pressurised payload fuselage has a substantially curved cross section when viewed along a longitudinal axis of each inner wing. A surface of the streamwise portion of the bulkhead may be curved. This is to maximise the structural integrity of the pressurised fuselage of the aircraft. It helps the payload fuselage to withstand internal pressures without need for unduly thick and therefore heavy materials. The sweep angle of each inner wing section may be greater than 50°. The sweep angle of the inner wing sections may be greater than the sweep angle of the outer wing sections. The relatively large sweep angle of the inner wing section allows the inner wing section to be relatively thick while minimising transonic drag. At least one vertex may exist in the front spar and rear spar at each transition region. That is to say, there may be two substantially straight front spars that join together at a vertex, for matching the change in sweep angle of the wing. Likewise, there may be two substantially straight rear spars that join together at a vertex. This is so that the spars align with the outer profile of the aircraft. The load-bearing structural element preferably joins the vertex of the front spar and the vertex of the rear spar on each wing for transmitting loads across the wing. This is because additional reinforcement at these areas of stress concentration is advantageous. Additionally, aerodynamic simulations carried out have shown that the height of the front spar close to the leading-edge vertex or kink of the V-shaped aircraft must be relatively small. At the same time, the stresses in this region have been found to be relatively high in FEM simulations. In the current approach, a majority of the loads can advantageously be transferred though the load-bearing structural element which can have greater height in the centre, which means the structural elements can withstand greater bending loads at less weight. Each wing may further comprise a plurality of ribs for supporting a skin of the aircraft. The plurality of ribs may be disposed across sections of the wings. The ribs may be disposed periodically along each wing. The plurality of ribs may be disposed in the rear portion of the aircraft, relative to the direction of flight. The ribs may be supported by the front and rear spars. Ribs may advantageously transfer loads from the skin to the front and rear spar. In alternative embodiments, the ribs may be disposed in any suitable arrangement. For example, they may be disposed from a front spar or rear spar. Alternatively, they may be disposed from a leading edge or trailing edge of the wing. In one embodiment, the ribs may be disposed from a leading edge to a rear spar. In a preferred embodiment, the bulkhead does not have a traditional hemispherical form. In alternative embodiments to those described above, the bulkhead may however have a continuous curved surface. According to a second aspect of the invention, there is provided a method of constructing an aircraft comprising the steps of: (a) constructing an inner wing comprising an integrated payload fuselage and a first edge; (b) constructing an outer wing for extending the wingspan, the outer wing comprising a second edge, in which the outer wing section has a thickness less than a thickness of the inner wing section; and (c) joining the first edge of the inner wing and the second edge of the outer wing. Each of the first and second edge may be streamwise. Advantageously, this allows modular construction of the aircraft, thereby making construction and the associated logistical challenges, for example of transporting sections of the aircraft during construction, simpler. The aircraft may comprise a starboard swept wing and a port swept wing, each swept wing may have an inner wing section and an outer wing section. There may be a central region of the aircraft where the starboard swept wing meets the port swept wing. The aircraft may comprise a starboard transition region where the starboard inner wing section meets the starboard outer wing section and a port transition region where the port inner wing section meets the port outer wing section. The sweep angle of each inner wing section may be greater than 50 degrees. The sweep angle of the inner wing may be greater than the sweep angle of the outer wing section. According to a third aspect of the invention, there is provided a method of constructing an aircraft comprising the steps of: (a) constructing an inner wing comprising a payload fuselage, a rear spar close to a trailing edge, an inner portion of a load-bearing structural element configured to connect the rear spar of the inner wing with a front spar of an outer wing; (b) constructing an outer wing comprising a front spar close to a leading edge, a rear spar close to a trailing edge and an outer portion of the loadbearing structural element; and (c) joining the inner wing with the outer wing at the load-bearing structural element and at the rear spar. This enables the inner wing and outer wing can be independently constructed, and then joined together at strong, structural spars. This can save construction costs and facilitate simple construction. The aircraft may comprise a starboard swept wing and a port swept wing, each swept wing may have an inner wing section and an outer wing section. There may be a central region of the aircraft where the starboard swept wing meets the port swept wing. The aircraft may comprise a starboard transition region where the starboard inner wing section meets the starboard outer wing section and a port transition region where the port inner wing section meets the port outer wing section. The sweep angle of each inner wing section may be greater than 50 degrees. The sweep angle of the inner wing may be greater than the sweep angle of the outer wing section. Preferably, the inner wing is provided with a front spar, and the front spar of the outer wing is joined to the front spar of the inner wing in step c. The two methods for constructing claimed herein may be a method of constructing an aircraft having any of the features of the aircraft claimed herein. BRIEF DESCRIPTION OF THE DRAWINGS For a better understanding of the present invention, and to show more clearly how it may be carried into effect, reference will now be made by way of example only to the accompanying drawings, in which: Figure 1 shows the starboard half of an aircraft; Figure 2 shows a portion of the starboard half of the aircraft of Figure 1; Figure 3 shows the starboard half of an aircraft in an alternative embodiment; Figure 4 shows the starboard half of an aircraft in a different alternative embodiment; Figure 5 shows starboard half of the aircraft of Figure 4; Figure 6 shows a method of construction for the embodiment of Figure 1; Figure 7 shows a portion of a starboard half of the aircraft in one embodiment of a previous concept; Figure 8 shows a portion of a starboard half of the aircraft in another embodiment of a previous concept; Figure 9 shows a portion of the starboard half of the aircraft in the previous concept shown in Figure 8; and Figure 10 shows a portion of a starboard half of the aircraft in another embodiment of a previous concept. DESCRIPTION OF PREFERRED EMBODIMENTS Referring firstly to Figure 1, the starboard half of an aircraft is indicated generally at 10. The aircraft comprises a central region 12, a starboard wing indicated at 14, and a port wing (not shown). Each wing 14 extends laterally from the central region 12. Each wing 14 encloses an integrated payload fuselage indicated at 20. Each wing 14 further comprises a leading edge 18 and a trailing edge 22. The leading edge 18 and the trailing edge 22 of each wing are substantially parallel. An engine 16 is disposed on each wing 14. Each engine 16 is disposed on trailing edge 22 of its respective wing 14. For brevity, the starboard wing 14 and not the port wing is shown in the drawings and described fully throughout the description. The port wing is symmetrical to the starboard wing and will comprise identical features. It will be appreciated that in embodiments, the aircraft may not be completely symmetrical, for example different equipment may be provided on either side. However most embodiments are likely to be roughly symmetrical in their main features. The starboard wing 14 is disposed extending from the central region 12. The wing 14 extends substantially horizontally. The wing 14 comprises an inner wing section 32 that attaches to the central region 12. Engine 16 is disposed on the inner wing section 32. The wing 14 comprises an outer wing section 30 which extends from the inner wing section 32. The inner wing section 32 has a positive sweep angle. Preferably, the inner wing section 32 has a sweep angle of over 50 degrees. The outer wing section 30 also has a positive sweep angle. The sweep angle of the outer wing section 30 is less than the sweep angle of the inner wing section 32. Each wing section 30, 32 has a leading edge 18. In this embodiment, the leading edge 18 extends along the inner wing section 32 and the outer wing section 30. The leading edge 18 runs in a straight line along the front of the inner wing section 32. There is then a vertex at the transition region where the leading edge 18 transitions to a significantly smaller sweep angle in the outer wing section 30. The trailing edge 22 runs in a straight line along the rear of the inner wing section 32. There is then a vertex at the transition region where the trailing edge 22 transitions to a significantly smaller sweep angle in the outer wing section 30. The outer wing section 30provides an increased wingspan of the aircraft. In this embodiment, it is the inner wing section 32 which integrates the payload fuselage 20. The outer wing section 30 is much thinner, more akin to the wing of a conventional aircraft. An additional vertex is present along the trailing edge 22 of the inner wing section 30. This additional vertex shows a smaller change in angle than at the transition region. The vertex is around halfway along the trailing edge 22 of the inner wing section 30. The outer wing section 30 provides an increased wingspan of the aircraft 10. In this embodiment, it is the inner wing section 32 which integrates the payload fuselage 20. The outer wing section 30 is much thinner, more akin to the wing of a conventional aircraft. A front spar 36 is disposed proximate to the leading edge 18 of the aircraft 10. The front spar 36 is substantially parallel to the leading edge 18. A rear spar 38 is disposed proximate to the trailing edge 22. The rear spar 38 is substantially parallel to the trailing edge 22 of the aircraft. In alternative embodiments, the rear spar may be further away from the trailing edge of the aircraft to enable storage of additional kerosine in the trailing edge, integration of a pressurized hydrogen tank in the trailing edge region or positioning the rear spar at a position of larger building heights to save weight. As such, there is a front vertex 50 in the front spar 36 at the bend in the leading edge 18. The front vertex 50 is close to the transition region. The front vertex 50 is at the join of the inner wing front spar, which has a high sweep angle, to the outer wing front spar, which has a lower sweep angle. There is a rear vertex 52 in the rear spar 38 near the bend in the trailing edge 22. The rear vertex 52 is close to the transition region. The rear vertex 52 is at the join of the inner wing front spar, which has a high sweep angle, to the outer wing front spar, which has a lower sweep angle. The front spar 36 and rear spar 38 are each disposed at edges of the payload fuselage 20. The front spar 36 and rear spar 38 further form a boundary with the payload fuselage 20. A winglet 34 is provided on the outer wing portion 30 at an outer end of the wing. The winglet 34 is substantially perpendicular to the outer portion of the wings 30 for reducing drag. The winglet 34 is oriented substantially vertically. The payload fuselage 20 is integrated into the inner wing section 32. In other words, the payload fuselage 20 is integrated near to the leading edge 18. Part of the wing 14 extends behind the payload fuselage 20 to the trailing edge 22. The payload fuselage 20 runs broadly parallel to the leading edge 18 of the wing 14. The payload fuselage 20 is integrated into the inner wing section 32 for more even weight distribution and therefore reduced bending moments. The payload fuselage 20 may be pressurised for accommodation of passengers. The payload fuselage 20 terminates at the transition region. The payload fuselage 20 terminates in a bulkhead. In this embodiment, the payload fuselage 20 extends from a leading edge of the aircraft to a rear spar of the aircraft. In alternative embodiments, it may extend from either one of the leading edge or the front spar of the aircraft, to either one of the trailing edge or the rear spar of the aircraft. A load-bearing structural element 40 is disposed from the front vertex 50 in the front spar 36 to the rear vertex 52 in the rear spar 38 to carry loads from the leading edge 18 of the outer wing 30 to the trailing edge 22 of the inner wing 32. In this embodiment, the load-bearing structural element 40 is a continuation of the rear spar 38 and front spar 36.. The load-bearing structural element 40 runs from the front vertex 50 in the front spar 36 to the rear vertex 52 in the rear spar 38. The payload fuselage 20 has a rear bulkhead. The rear bulkhead has a streamwise edge 42 and a second edge in the transition region. The second edge may be parallel to a line between the front vertex and rear vertex. In this embodiment, the streamwise edge 42 and second edge are perpendicular to each other. A portion of the second edge of the payload fuselage 20 is formed by the load-bearing structural element 40. The payload fuselage 20 has a substantially curved cross section, as viewed along an axis of the wing. The load-bearing structural element 40 is a spar. The load-bearing element may alternatively be a truss. The load-bearing element has a height. The height of the load-bearing element may match a height of the wing. The load-bearing structural element 40 may be substantially sheet-shaped. The load-bearing structural element 40 is of sufficient thickness to carry loads between the front and rear spars, in addition to the loads from the pressurisation of the payload fuselage 20. The streamwise edge 42 may be curved to maximise load-bearing capability. Alternatively, the streamwise edge 42 may be another vertical sheet. The position of the oval retention parameter 70, defined herein, is indicated by a dashed line. The oval retention parameter is shown as a line in a streamwise direction at the transition region. In this embodiment, the oval retention parameter is closer to a centre of the aircraft than the streamwise edge of the bulkhead 42. Figure 2 shows a section of the starboard wing 14, with additional thickness on the spars used to indicate the primary load bearing path. Lift loads are carried on the front spar 36 of the outer wing 30. At the front vertex 50 of the front spar 36, the lift load is transferred to the continuation of the front spar 36 on the inner wing 32 and the load-bearing structural element 40. The load-bearing structural element 40 transfers the lift load to the rear spar 38 of the inner wing 32. Conversely, the weight load, for example from the engine 16, is carried by the rear spar 38 of the inner wing 32. The weight load is transferred to the load-bearing structural element 40. The load-bearing structural element 40 transfers the weight load to the front spar 36 of the outer wing 30. The load-bearing structural element 40 transfers the weight load to a lesser extent to the front spar 36 of the inner wing 32. Figure 3 shows an alternative embodiment of the present invention. The aircraft 110 comprises a central region 112, a starboard and a port wing (not shown), an integrated payload fuselage 120, a leading edge 118, trailing edge 122 and engine 116, disposed as in previous embodiments. The integrated payload fuselage terminates at a bulkhead 144 in each transition region. In contrast to the previous embodiment, there is no streamwise and second portion of the bulkhead. Instead, the bulkhead is a substantially continuous surface. At least a portion of bulkhead 144 is disposed substantially perpendicular to the leading edge 118. The bulkhead 144 is disposed from a rear vertex of the rear spar 138 to a leading edge of the wing 118. The bulkhead 144 is a curved surface. Alternatively, the bulkhead is a flat surface. The load-bearing structural element 140 is disposed as in the previous embodiments, from a front vertex 150 to the rear vertex 152. Unlike in previous embodiments, the load-bearing structural element 140 does not form a part of the bulkhead of the payload fuselage 120. Figure 4 shows a third embodiment, identical to the first embodiment, but in which an indicative position of structural ribs is shown. Ribs, one of which is indicated at 260, are disposed periodically along the wing. In this embodiment, ribs 260 are disposed from a leading edge 218 to a rear spar 238. Each rib 260 is substantially perpendicular to a leading edge 218 of the wing 214. Each rib 260 is substantially perpendicular to the trailing edge 222. As such, the angle of each rib 260 changes progressively through the transition region. The ribs 260 are disposed periodically from close to the tip of the winglet 234 to around halfway along the inner wing 232. Figure 5 is an annotated copy of Figure 4, showing the same embodiment as in Figure 4. The oval retention parameter, as defined herein, is indicated at 70. Figure 6 shows how the wing of the embodiment shown in Figure 1 can be assembled with simplicity. The outer wing 32 can be constructed separately to the inner wing 30 and payload fuselage 20. Subsequently, the outer wing 32 can be attached to the inner wing 30 via at least three attachment points, shown as dots. Corresponding features of the wing are shown as Xa on the inner wing 32 and Xb on the outer wing 30. The first attachment point is at the front spar 36a of the inner wing portion. The second attachment point is at the load-bearing structural element 40. The second attachment point is substantially halfway along the load-bearing structural element 40. The third attachment point is at the rear spar 38b of the outer wing portion 30. Figures 7 to 10 show previous concepts, as discussed in the background. 5 The embodiments described above are provided by way of example only, and various changes and modifications will be apparent to persons skilled in the art without departing from the scope of the present invention as defined by the appended claims.

Claims

1. An aircraft comprising:a starboard swept wing and a port swept wing, each swept wing having an inner wing section comprising an integrated payload fuselage and an outer wing section for extending the wingspan, in which each outer wing section has a thickness less than a thickness of the corresponding inner wing section;a front spar close to a leading edge of each wing, in at least the outer wing section, and a rear spar close to a trailing edge of each wing, for carrying loads along each wing;a central region where the starboard swept wing meets the port swept wing;a starboard transition region where the starboard inner wing section meets the starboard outer wing section;a port transition region where the port inner wing section meets the port outer wing section; anda load-bearing structural element disposed across each wing from the front spar to the rear spar in the transition region for carrying loads between the front spar and the rear spar of each wing.

2. An aircraft as claimed in claim 1, in which the integrated payload fuselage comprises a bulkhead in each transition region.

3. An aircraft as claimed in claim 2, in which a portion of the load-bearing structural element forms part of or is attached to a second portion of the bulkhead of the payload fuselage.

4. An aircraft as claimed in any preceding claim, in which the load-bearing structural element is a beam.

5. An aircraft as claimed in any preceding claim, in which the load-bearing structural element is of varying height for matching a height of the wing.

6. An aircraft as claimed in any preceding claim, in which there is a streamwise portion bulkhead of payload fuselage.

7. An aircraft as claimed in claim 6, in which the streamwise portion of the bulkhead joins with the load-bearing structural element.

8. An aircraft as claimed in any previous claim, in which the payload fuselage is pressurised.

9. An aircraft as claimed in claim 8, in which the pressurised payload fuselage has a substantially curved cross section.

10. An aircraft as claimed in claim 8 or 9, in which a surface of the streamwise portion of the bulkhead may be curved.

11. An aircraft as claimed in any preceding claim, in which the sweep angle of each inner wing section is greater than 50°.

12. An aircraft as claimed in any preceding claim, in which the sweep angle of the inner wing section is greater than the sweep angle of the outer wing section.

13. An aircraft as claimed in claim 12, in which at least one vertex exists in the front spar and rear spar at each transition region, for matching the change in the sweep angle of the wing.

14. An aircraft as claimed in claim 13, in which the load-bearing structural element joins the vertex of the front spar and the vertex of the rear spar at each transition region for transmitting loads across the wing.

15. An aircraft as claimed in any previous claim, in which each wing further comprises a plurality of ribs for supporting a skin of the aircraft.

16. An aircraft as claimed in claim 15, in which each rib is disposed across sections of each wing.

17. An aircraft as claimed in claim 15 or 16, in which each rib is disposed periodically along each wing.

18. An aircraft as claimed in any of claims 15to 17, in which each rib is disposed along a rear portion of the aircraft.

19. An aircraft as claimed in any of claims 15 to 18, when dependent on claim 4, in which each rib attaches to the front and rear spars of the aircraft.

20. A method of constructing an aircraft, comprising the steps of:(a) constructing an inner wing comprising an integrated payload fuselage and a first edge;(b) constructing an outer wing for extending the wingspan, the outer wing comprising a second edge, in which the outer wing section has a thickness less than a thickness of the inner wing section; and(c) joining the first edge of the inner wing and the second edge of the outer wing.

21. A method of constructing an aircraft as claimed in claim 20, in which the first and second edge are each streamwise.

22. A method of constructing an aircraft, comprising the steps of(a) constructing an inner wing comprising a payload fuselage, a rear spar close to a trailing edge, an inner portion of a load-bearing structural element configured to connect the rear spar of the inner wing with a front spar of an outer wing;(b) constructing an outer wing comprising a front spar close to a leading edge, a rear spar close to a trailing edge and an outer portion of the load-bearing structural element; and(c) joining the inner wing with the outer wing at the load-bearing structural element and at the rear spar.

23. A method of constructing an aircraft as claimed in claim 22, in which the inner wing is provided with a front spar, and the front spar of the outer wing is joined to the front spar of the inner wing.