An aircraft

Abstract

An aircraft 10 comprising a first swept wing 14A and a second swept wing 14B, each swept wing having an inner wing section 26A, 26B comprising an integrated payload fuselage 20 and an outer wing secti

Description

The present invention relates to an aircraft. BACKGROUND TO THE INVENTION V-shaped aircraft are known, and proof-of concept prototypes have been made. V-shaped aircraft can have fuselage sections integrated into the wings. Integrating the fuselage sections into the wings can reduce the bending moments in an aircraft by distributing the mass more evenly across its span. This means that the structure of the aircraft can be made sufficiently strong without adding excessive weight. Bending moments are experienced by aircraft because the shape of the wings is designed to generate an upwards lift force, while the central fuselage generally contains more mass and experiences a downwards gravitational force. Opposing forces generate a bending moment between the fuselage and the wings. Significant bending moments can be a problem as they exert stress on the materials and joints of an aircraft which can lead to cracks and eventually decommissioning or structural failure. They also mean the aircraft can be heavier as thicker materials are used in construction to withstand the forces exerted on the aircraft. V-shaped aircraft minimise this problem by integrating the fuselage section into the wings, thereby ensuring a lift is generated across the entire aircraft, and that the mass is distributed more evenly across the entire aircraft. A problem with all aircraft is the trade-off between space, weight and aerodynamics. All aircraft are designed to maximise the internal space of the aircraft, whilst keeping the weight as low as possible, whilst optimising the aerodynamics of the shape. The internal volume of the aircraft should be maximised so more passengers or cargo can be transported in order to make the flight cost-effective. However generally adding volume to the aircraft decreases the aerodynamic performance. Bigger aircraft are also heavier which means more fuel is required. The aerodynamic performance is dictated in part by the outer surface area of the aircraft, which should be kept as low as possible for low friction drag, and the wingspan, which should be kept as high as possible for low induced drag. In addition to that, if the aircraft should cruise at high subsonic speeds, it is desirable to keep the wing as thin as possible. Thin wings minimise aerodynamic shocks on the wing, which in turn minimise transonic drag. This is a problem with current V-shaped aircraft, which have fuselage in the wings and therefore thicker wings than conventional aircraft. Transonic drag generally increases with the thickness-to-chord ratio of the wing. The higher the thickness-to-chord ratio, the worse the transonic drag. Transonic drag further depends on the local sweep angle of the wing and the local lift coefficient Pressure drag exists behind the fuselage and is also an aerodynamic disadvantage. Finally, flow separation can exist and further hinder the aerodynamics of the aircraft. Conventionally, liquid fuel for an aircraft is stored in the wings. This is a good solution because it does not reduce the useful payload space (in a conventional aircraft, passengers or cargo cannot be put in the wings), and distributes the weight of the fuel over the span of the aircraft, reducing bending moments. Similarly, it has been proposed in “v-shaped” aircraft also to store liquid fuel in the wings, specifically in a rear section of the wing, between the integrated fuselage and the trailing edge of the wing. However, there is now increasing interest in the use of hydrogen as an aircraft fuel. This leads to certain challenges, firstly because the energy density of hydrogen is lower than conventional aviation fuel, and so a greater volume of hydrogen is needed to achieve the same range, and secondly because hydrogen needs to be stored in pressurised tanks. It cannot easily be put into the wing in the same way as conventional liquid fuels. The same issues arise with other alternative fuels, e.g. ammonia. The integration of fuel tanks can have a negative impact on a number of aspects of the aircraft. Firstly, the addition of tanks can impact the aerodynamics of an aircraft if they are fixed on the exterior of the aircraft. The addition of heavy tanks to the centre of the aircraft would add to the central weight and thereby increase the bending moments between the lift of the wings and the weight in the centre. This would necessitate increased weight to provide the structural integrity necessary to support these bending moments, and an aircraft with worse performance and shorter range. Retrofitting aircrafts with additional tanks is sometimes required to alter the range of the aircraft. Currently, this can be very time-consuming, expensive and can cause regulatory or safety issues. Fuel tanks that are too close to or not sufficiently isolated from the passenger carrier can be dangerous in the event of a crash. An additional consideration is that fuel tanks are ideally placed spaced from the engines in either a forwards or aft direction. This is because an engine burst event results in sharp debris being expelled radially from the engine which can cause fire and explosion if they hit the fuel tanks. Current fuel tanks can be inaccessible as they are embedded in the wings of the aircraft. This can hinder maintenance and add to the time and cost of routine checks and repairs. Aircraft with large wingspans require adequately wide landing spaces and space to manoeuvre on the ground. In many cases the main restriction is imposed by the gates I aprons at airports. An aircraft with a large wingspan will not be able to use some airport gates. It is an object of the present invention to reduce or substantially obviate the aforementioned problems. STATEMENT OF INVENTION According to a first aspect of the present invention there is provided an aircraft comprising: a first swept wing and a second 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 central region where the first swept wing meets the second swept wing; a first transition region where the first inner wing section meets the first outer wing section; a second transition region where the second inner wing section meets the second outer wing section; in which at least one extension having an internal volume is provided behind the trailing edge of the wings, the at least one extension being disposed in the central region. The arrangement of the central extension advantageously increases the internal volume of the aircraft while only a small impact on the aerodynamic performance of the aircraft can be expected. Increasing the internal volume provides additional space, particularly for fuel. The central extension may be a fuel tank, for example a pressurised tank for hydrogen, or may at least provide a space in which such a tank may be disposed or partially disposed. The outer wing section has a thickness (meaning absolute thickness) less than a thickness of the corresponding inner wing section. Thickness-to-chord ratio is also discussed in parts of this disclosure, which will always be referred to as thickness-to-chord ratio. “Thickness” on its own means absolute thickness. It may be in the central region where the two wings meet that the inner wing sections are at their thickest. The result is a large thickness to chord ratio which produces a relatively high pressure drag in this region. The introduction of the central extension increases the chord length in the central region, decreasing the thickness to chord ratio and the pressure drag. Providing at least one extension on the trailing edge facilitates retrofitting aircraft by introducing the extension to the central region. Positioning the at least one extension on the trailing edge also improves accessibility for maintenance of the extension and any tank or other equipment within the extension. The size of the extension may be scalable dependent on the additional internal volume required. For example, the size of the at least one extension is not dependent on the size of the fuselage sections. The size of the integrated fuselage section may be increased or decreasing by adding or removing parts of the fuselage sections without necessitating movement of the extension. Advantageously, the extension is positioned in proximity to a structural element which connects the wings of the aircraft. The aircraft may have a substantially V-shaped planform. The sweep angle of inner wing sections may be greater than 50°, preferably greater than 55° or even more preferably greater than 60°. The sweep angle of the inner wing sections may be greater than the sweep angle of the outer wing sections. The higher sweep angle of the inner wing section allows the inner wing section to be relatively thick while minimising transonic drag. The at least one extension may be cylindrical. The at least one extension may be cone-shaped. A cylindrical or cone shape may improve or at least not be detrimental to aerodynamic performance. The at least one extension may be orientated with its elongate extent in a streamwise direction. The cylinder or cone diameter may be substantially similar to a maximum thickness of the swept wings at the central region. This may improve or at least not be detrimental to aerodynamic performance, compared to the aircraft without the extension. The at least one extension may be disposed forwards of an engine. The at least one extension or tank may be positioned forwards of the engine so that in the event of an engine burst, the extension is outside the engine burst region. The at least one extension may be a pressurised fuel tank for receiving pressurised fuel. For example, the pressurised fuel may be liquid hydrogen. The at least one extension may be separated from the integrated fuselage. In particular, where the extension is pressurised, separating the pressurised extension from the passengers may be beneficial for safety in the event of an accident. This arrangement may be beneficial for crashworthiness. The outer wing section may comprise winglets. At least one inner fuel tank may be provided in at least one of the first and second inner wing sections. This further provides additional internal volume for fuel. The at least one inner tank may be orientated in a direction corresponding to the direction of the corresponding inner wing section. The inner fuel tank may protrude from the respective inner wing to the corresponding transition region. According to a second aspect of the present invention, there is provided an aircraft comprising: a first swept wing and a second swept wing, each swept wing having an inner wing section having 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 central region where the first swept wing meets the second swept wing; a first transition region where the first inner wing section meets the first outer wing section; a second transition region where the second inner wing section meets the second outer wing section; in which at least one extension having an internal volume is provided behind the trailing edge of the wings, the at least one extension being disposed in at least one of the first and second transition regions. Many of the advantages have been discussed with respect to the first aspect of the present invention. The outer wing sections may have a lower sweep angle and a smaller chord length compared to the inner wing section and so must be thinner than the inner wing section. However, there is a high thickness-to-chord ratio in the transition region. A high thickness-to-chord ratio in this region may lead to increased transonic drag. The extension increases the chord length and decreases the thickness-to-chord ratio in these regions and decreases pressure drag at the expense of increasing the outer surface area and increasing friction drag. The result is an increased internal volume with minimal effect on overall aerodynamic performance. Advantageously, the extension is positioned away from the centre of the aircraft resulting in lower bending moments. The at least one extension may comprise a vertical fin. The vertical fin may mitigate effects of vortices originating at the trailing edge of the transition region. The vertical fin may improve the directional stability of the aircraft. The at least one extension may be disposed longitudinally offset from at least one engine. The at least one tank may be positioned laterally adjacent to the engine so that in the event of an engine burst, the extension is outside the engine burst region. The trailing edge of the outer wing section may be disposed higher than the trailing edge of the inner wing section. This provides more clearance, allowing for a shorter landing gear and thereby reducing the weight of the aircraft. A maximum chord length of the outer wing section may be less than a minimum chord length of the inner wing section. The aircraft may be adapted to be powered by hydrogen fuel, or ammonia, or another fuel which is stored in pressurized tanks. In some embodiments, the aircraft may be a dual-fuel or multi-fuel aircraft, with conventional liquid fuel being stored in the wings (between the trailing edge of the wing and the integrated fuselage and / or in the outer wings). The aircraft may have any feature or combination of features of the first aspect of the present invention. Providing both extensions in the central region and transition regions may be beneficial for controlling the position of the centre of gravity of the aircraft by adjusting the size of the extensions. According to a third aspect of the present invention, there is provided an aircraft comprising: a first swept wing and a second swept wing, each swept wing having an inner wing section comprising an integrated 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 central region where the first swept wing meets the second swept wing; a first transition region where the first inner wing section meets the first outer wing section; a second transition region where the second inner wing section meets the second outer wing section; in which at least one inner fuel tank is provided in the integrated fuselage. The integrated fuselage in most embodiments will include space for cargo and / or seating. However, note that in embodiments the inner fuel tank(s) may be provided inside or outside any pressurised cabin section. A pressure bulkhead may be provided between the inner fuel tank(s) and any payload space for passengers I cargo. In other embodiments the integrated fuselage could be provided for example in the form of tubes. A wide tube, disposed at the front (i.e. close to the leading edge) of the wing could provide a space for passengers and cargo, with another tube being provided behind the wide tube (i.e. between the wide tube and the trailing edge of the wing) to accommodate inner tank(s). The inner fuel tank provides additional internal volume for fuel. The inner fuel tank may protrude from the respective inner wing to the corresponding transition region. The inner fuel tank may be cylindrical. The at least one inner tank may be orientated in a direction corresponding to the direction of the corresponding inner wing section. The inner fuel tank may be disposed parallel to the inner wing. The aircraft may have any feature or combination of features of the first and second aspect of the present invention. 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 a top view of a cross section of an aircraft according to the present invention; Figure 2 shows a top view of the aircraft of Figure 1; Figure 3 shows a rear view of the aircraft of Figure 1; Figure 4 shows a top view of a second embodiment of an aircraft; Figure 5 shows a schematic of the aircraft of Figure 4; Figure 6 shows a schematic of a third embodiment of an aircraft; Figure 7 shows a schematic of a fourth embodiment of an aircraft; Figure 8 shows a schematic of a fifth embodiment of an aircraft; Figure 9 shows how the thickness-to-chord ratio varies across the width of a v-shaped aircraft. DESCRIPTION OF PREFERRED EMBODIMENTS Referring firstly to Figure 1, an aircraft is indicated generally at 10. The aircraft comprises a central region 12, a first (starboard) wing indicated at 14A and a second wing (port) indicated at 14B. The wings extend laterally from the central region 12. The wings enclose an integrated payload fuselage indicated at 20. Each wing 14, 16 comprises leading edges 18A, 18B and trailing edges 22A, 22B. Starboard tank 38A and port tank 38B are disposed on the trailing edges 22A, 22B of each wing 14A, 14B. Central tank 44 is disposed aft of the central region 12, between the wings 14A, 14B. Each of the tanks 38A, 38B, 44 is of an elongate shape. Tanks may be roughly cylindrical but, as shown in Figure 1 the shape of the tank may be as a truncated cone, to maximise interior volume while not adversely affecting aerodynamic performance. The tanks are preferably disposed in conical fairings. The tanks may be pressurised to store, for example, hydrogen fuel. In other embodiments the tanks may store conventional liquid aviation fuel. For brevity, the starboard wing and not the port wing is described fully. The port wing is symmetrical to the starboard wing and will comprise identical features. Symmetrical I identical features labelled on the starboard wing as xA will be labelled on the port wing as xB. 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 26A that attaches to the central region 12. The wing 14 comprises an outer wing section 30A which extends from the inner wing section 26A. The inner wing section 26A has a positive sweep angle. Preferably, the inner wing section 26A has a sweep angle of over 60 degrees. The outer wing section 30A also has a positive sweep angle. The sweep angle of the outer wing section 30A is less than the sweep angle of the inner wing section 26A. Each wing section 26A, 30A has a leading edge. In this embodiment a leading edge 18A extends along the inner wing section 26A and the outer wing section. The leading edge 18A runs in a straight line along the front of the inner wing section 26A. There is then a discontinuity in the transition region where the leading edge 18A transitions to a significantly smaller sweep angle in the outer wing section 30A. The outer wing section 30A provides an increased wingspan of the aircraft. In this embodiment, it is the inner wing section which integrates the payload fuselage. The outer wing section 30A is much thinner, more akin to the wing of a conventional aircraft. The payload fuselage 20 is integrated into a front part of the inner wing section 26A. In other words, the payload fuselage 20 is integrated near to the leading edge 18A, and part of the wing extends behind the payload fuselage 20 to the trailing edge. The payload fuselage 20 runs broadly parallel to the leading edge of the wing. The payload fuselage 20 is integrated into the inner wing section 26A for more even weight distribution and therefore reduced bending moments. The payload fuselage 20 may be pressurised for accommodation of passengers. The inner wing section 26A comprises payload fuselage 20. Payload fuselage 20 contains a pressurised cabin for passengers 24A. The pressurised cabin 24A is placed in front of the tanks 38A, 38B relative to the direction of travel for increased safety to passengers in case of fuel fire. Although tank 44 is close to the front of the aircraft, there is no part of the fuselage or cabin immediately behind it. The engines 46A, 46B are also positioned aft of the pressurised cabin 24A for protecting the passengers from debris in case of engine burst. The payload fuselage 20 comprises a cargo region indicated generally at 28A, immediately aft of the passenger cabin 24A. The payload fuselage 20 comprises an additional fuel tank 29A. The additional fuel tank 29A is partitioned from the rest of the payload fuselage 20 for protection of the passengers in case of fuel fires. The additional fuel tank 29A is positioned aft of the pressurised cabin 24A, and in this embodiment aft of the cargo region 28A as well. The wing 14A is in profile substantially aerofoil-shaped for producing upwards lift to the aircraft. This generates lift across a majority of the width of the aircraft 10. Starboard engine 46A is provided on the trailing edge of the starboard wing 14A, approximately halfway along the inner wing section 26A. The engine 46A is a conventional jet engine and is substantially cylindrically shaped. The long axis of engine 46A is oriented in a streamwise direction, although in some embodiments the engine may be tilted slightly either upwards, downwards, or sideways from the streamwise direction. The wing 14A comprises a transition region 23A where the inner wing 26A is joined to the outer wing 30A. Starboard outer tank 38A is disposed proximate to transition region 23A. The starboard tank 38A (or, a fairing surrounding the tank) has an outer surface that is substantially parabolic-cone shaped for reducing pressure drag. The pressure drag can be high behind the inner wing 26A due to the high sweep angle and the thickness of the inner wing 26A. Starboard outer tank 38A has a substantially circular cross section. The tip of the parabolic-cone shaped tank 38A is aft of the base. The tank 38A is oriented such that the long axis of the cone is substantially aligned with a streamwise direction. Outer tank 38A has a root diameter similar to the maximum thickness of the wing. Outer tank 38A is distanced laterally and longitudinally from the engine 46A such that it is unlikely to be hit by debris in the event of an engine burst. Outer tank 38A is enclosed by a fairing 39A for encasing the tank and ensuring the exterior of the aircraft is suitably aerodynamic. Similarly, port outer tank 38B is disposed on the transition region 23B of the port-side wing. The central region 12 comprises a joining region 42 connecting the two trailing edges of the starboard and port wings. A central tank 44 is mounted onto this joining region 42, and is disposed aft of the joining region and between the trailing edges of the two wings. The central tank 44 (or an associated fairing) has an outer surface that is substantially parabolic-cone shaped for reducing pressure drag. The central tank 44 is oriented such that the long axis of the cone is substantially aligned with a streamwise direction. The tip of the parabolic-cone shaped tank 44 is aft of the base. The central tank 44 has a root diameter similar to the maximum thickness of the wing profile. The engine 46A is positioned aft of the central tank 44, for protecting the central tank 44 from potential engine burst. The tanks 38A, 38B, 44 are similar in function. The tanks 38A, 38B, 44 are used to accommodate fuel. The tanks 38A, 38B, 44 can accommodate a pressurized volume. The pressurised fuel tanks can be used to store liquid hydrogen. More generally, in other embodiments, instead of fuel tanks 38A, 38B, 44, additional volumes may be provided in corresponding positions for other purposes. For example, volumes in place of one or more of the tanks 38A, 38B, 44 could be used as additional space for passengers or cargo. Figure 2 shows another embodiment. In the embodiment of Figure 2, the outer tank 38A extends through the full width of the starboard wing 14A, and aft of the trailing edge 22A into an extension behind the wing. Figure 3 shows a rear view of the aircraft 10 of Figure 1. A fin 48A is provided on the top of the outer tank 38A for minimising vortices in air flow and maintaining directional stability. The fin 48A is substantially vertical. The fin 48A is substantially perpendicular to the tangent of the surface of the outer tank 38A. The wing 14A has an upper surface indicated generally at 16A and a lower surface indicated generally at 17A. The upper surface of the wing 16A is substantially continuous across the width of the wing, including the inner and outer wing sections. The lower surface of the wing 17A however is at a significantly increased height on the outer wing section as compared to the height of the lower surface of the inner wing section. In other embodiments, the outer tank 38A can be positioned to bridge a more significant height difference between the inner wing portion 26A and the outer wing portion 30A. In other words, both the upper and lower surfaces of the outer wing section could be at a height above the upper surface of the inner wing section. This is advantageous as it ensures ground clearance during take-off and landing, while minimising the extent and weight associated with landing gear (50A, 52A, 54A). A winglet 34A is provided on the outer wing portion 30A at the distal end of each wing. The winglet 34A is substantially perpendicular to the outer portion of the wings 30A for reducing drag. The winglet 34A is oriented substantially vertically. A further alternative embodiment is shown in Figure 4. As before, an outer wing portion 130A is provided attached to inner wing portion 126A. The outer wing portion 130A can be offset from the inner wing portion 126A such that it is substantially set back from the inner wing portion, relative to the direction of travel. In alternative embodiments, the outer wing portion 130A can be offset from the inner wing portion 126A such that it is substantially set forwards from the inner wing portion, relative to the direction of travel. The position of the tank 138A allows any horizontal or vertical discontinuity between the inner and outer wing portions to be effectively bridged. Still referring to Figure 4, the outer tank 138A is oriented and shaped as described in previous embodiments. The outer tank 138A is disposed proximate to transition region 123A. The outer tank 138A may extend to the leading edge 118A of the inner wing 126A. In alternative embodiments, the outer tank 138A may protrude from the leading edge 118Aofthe inner wing 126A. Figure 5 shows another alternative embodiment. The inner wing portion 226A comprises a configuration of additional fuel tanks, indicated generally at 229A. The additional fuel tanks 229A are substantially cylindrical in shape. The additional fuel tanks are elongated. The additional fuel tanks are stored parallel to each other. The additional fuel tanks are further stored parallel to the outer edge of the wing 218A. Three additional tanks indicated generally at 229A are provided. Figures 6, 7 and 8 show further alternative embodiments showing different arrangement options for the additional fuel tanks. The aircraft 10 comprises the outer fuel tanks as described in previous embodiments. The additional space provided by the outer fuel tanks can in some embodiments be used to enlarge and provide more space for the additional tanks. The addition of an outer tank such as tank 38, or indeed any other volume in a corresponding position to hold for example cargo or passengers, can minimise the thickness-to-chord ratio in the relevant cross section of the aircraft 12. This subsequently minimises the transonic drag experienced by the wing when operating at speeds approaching the speed of sound. This is otherwise high in V-shaped aircraft due to the increased thickness of the wing. Figure 9 illustrates how the additional volume in the form of the tank 38 or fairing affects the thickness-to-chord ratio. Addition of tanks, or more generally, volumes which may be used for fuel, cargo, passengers, or aircraft equipment, in the positions described, provides for an aircraft with additional volume. This is required for a practical hydrogen-powered aircraft since, compared with conventional liquid aviation fuel, a greater volume of hydrogen needs to be used to achieve the same range. Also, hydrogen needs to be stored in pressurised tanks. In any case, the additional volume is provided without significantly compromising on aerodynamic performance, and advantageously the outer tanks in the transition regions can be used to bridge discontinuities between the inner and outer wing sections. This in particular means that the outer wing section can be set at a higher level, reducing the required extent and weight of landing gear. 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 first swept wing and a second 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 central region where the first swept wing meets the second swept wing;a first transition region where the first inner wing section meets the first outer wing section;a second transition region where the second inner wing section meets the second outer wing section;in which at least one extension having an internal volume is provided behind the trailing edge of the wings, the at least one extension being disposed in the central region.

2. An aircraft as claimed in claim 1, in which the sweep angle of each inner wing section is greater than 50°.

3. An aircraft as claimed in claim 1 or claim 2, in which the sweep angle of the inner wing section is greater than the sweep angle of the outer wing section.

4. An aircraft as claimed in any preceding claim, in which the at least one extension is cylindrical or cone shaped.

5. An aircraft as claimed in claim 4, in which the at least one extension is orientated in a streamwise direction.

6. An aircraft as claimed in claim 5, in which the cylinder or cone diameter is substantially similar to a maximum thickness of the swept wings at the central region.

7. An aircraft as claimed in any preceding claim, in which the at least one extension is separated from the integrated fuselage.

8. An aircraft as claimed in any preceding claim, in which the at least one extension is disposed fore of an engine.

9. An aircraft as claimed in any preceding claim, in which the at least one extension is a pressurised fuel tank for receiving pressurised fuel.

10. An aircraft as claimed in any preceding claim, in which the outer wing section comprises winglets.

11. An aircraft as claimed in any preceding claim, in which at least one inner fuel tank is provided in at least one of the first and second inner wing sections.

12. An aircraft as claimed in claim 11, in which the at least one inner fuel tank is orientated in a direction corresponding to the direction of the corresponding inner wing section.

13. An aircraft comprising:a first swept wing and a second swept wing, each swept wing having an inner wing section having 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 central region where the first swept wing meets the second swept wing;a first transition region where the first inner wing section meets the first outer wing section;a second transition region where the second inner wing section meets the second outer wing section;in which at least one extension having an internal volume is provided behind the trailing edge of the wings, the at least one extension being disposed in at least one if the first and second transition region.

14. An aircraft as claimed in claim 13, in which the at least one extension comprises a vertical fin.

15. An aircraft as claimed in claim 13 or claim 14, in which the at least one extension is disposed laterally adjacent to at least one engine.

16. An aircraft as claimed in any of claims 13 to 15, in which the trailing edge of the outer wing section is disposed higher than the trailing edge of the inner wing section.

17. An aircraft as claimed in any of claims 13 to 16, in which a maximum chord length of the outer wing section is less than a minimum chord length of the inner wing section.

18. An aircraft comprising:a first swept wing and a second 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 central region where the first swept wing meets the second swept wing;a first transition region where the first inner wing section meets the first outer wing section;a second transition region where the second inner wing section meets the second outer wing section;in which at least one inner fuel tank is provided in at least one integrated payload fuselage.

19. An aircraft as claimed in claim 18, in which the at least one inner fuel tank protrudes from the inner wing to the transition region.

20. An aircraft as claimed in claim 18 or 19, in which the inner fuel tank is cylindrical.

21. An aircraft as claimed in any of claim 20, in which the inner fuel tank is disposed parallel to the inner wing.

22. An aircraft as claimed in any preceding claim, in which the aircraft comprises a liquid hydrogen fuel.17

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