An eVTOL aircraft fuselage structure

Abstract

The present invention relates to the field of aircraft technology, and in particular to an eVTOL aircraft fuselage structure, comprising: a fuselage frame, a fuselage skin provided on the surface of the fuselage frame, a fuselage partition provided on the inner wall of the fuselage skin, and the fuselage frame, fuselage skin, and fuselage partition together forming a fuselage. The present invention has the advantages of inner and outer skin design, expanded visual range, integrated slide rails, increased support effect, and internal partition design, and provides an innovative inner and outer double-layer skin fuselage structure design, aiming to improve the overall strength, rigidity, and lightweight effect of the fuselage. The outer skin is made of carbon fiber prepreg and the inner skin is made of 7075 high-strength aluminum alloy to ensure the high strength of the outer layer and the lightweight of the inner layer. Adhesive bonding and bolting are used to ensure the close connection between the two skin layers and the efficiency of load transfer. Then, under the action of the driver's field of view window and the passenger field of view window, the fuselage structure design of the visual range is optimized.

Description

Technical Field

[0001] The present invention relates to the field of aircraft technology, and in particular to an eVTOL aircraft fuselage structure. Background Art

[0002] The fuselage of an electric aircraft is the main body of the aircraft that carries passengers and equipment. On the one hand, it protects the safety of passengers, and on the other hand, it transfers the aircraft load, maintains the aircraft's streamlined shape, and reduces automatic resistance. At the same time, the design of the fuselage also needs to meet many functional requirements, such as carrying equipment, opening and closing cabin doors, maintenance passages, etc. There are also many current designs of electric aircraft, all of which are designed to meet the functional requirements of different electric aircraft themselves.

[0003] The common fuselage structure on the market has only one layer of skin. The single-layer skin may not be as good as the double-layer skin structure in terms of compression, torsion and shear resistance, resulting in insufficient stability and strength of the fuselage when bearing complex loads; then the common fuselage will use traditional metal alloy materials, which will lead to excessive overall weight, which not only increases fuel consumption during flight, but also limits the aircraft's flight performance and load-bearing capacity; further, the fuselage may be optimized in load distribution. For example, high-strength materials are not used in areas with larger loads, while unnecessary thick materials are used in areas with smaller loads, which not only wastes materials, but also reduces the overall strength and stability of the fuselage; finally, the fuselage structure is relatively complex, containing multiple components and connection points, which not only increases the difficulty of manufacturing and assembly, but also makes maintenance and repair more difficult. In addition, the complex structure also increases the probability of failure.

[0004] Therefore, there is an urgent need for an eVTOL aircraft fuselage structure to solve the above problems. Summary of the Invention

[0005] The object of the present invention is to provide an eVTOL aircraft fuselage structure with the advantages of inner and outer skin design, expanded visual range, integrated slide rails, increased support effect and internal partition design, which solves the problems raised by the above-mentioned background technology.

[0006] To achieve the above-mentioned objectives, the present invention provides the following technical solutions: an eVTOL aircraft fuselage structure, comprising: a fuselage frame, a fuselage skin is provided on the surface of the fuselage frame, a fuselage partition is provided on the inner wall of the fuselage skin, and the fuselage frame, fuselage skin and fuselage partition together constitute the fuselage.

[0007] The fuselage frame includes a longitudinal frame, a circumferential frame, a transverse frame and a combined frame. The longitudinal frame is located at the bottom of the fuselage, the upper part of the fuselage, and the middle and rear part of the fuselage. The longitudinal frame at the bottom of the fuselage and the longitudinal frame at the upper part of the fuselage are both machined from 7075 high-strength aluminum alloy. The longitudinal frames at other positions are formed by paving carbon fiber composite materials and then hot-autoclaving. The circumferential frame is located at the tail of the fuselage. The transverse frame is located at the bottom of the fuselage and the joint between the wing and the fuselage. The transverse frame is 7075 and is mostly machined from high-strength aluminum alloy, and some are formed by paving carbon fiber composite materials and then hot-autoclaving. The combined frame is located at the front and center of the fuselage.

[0008] The longitudinal skeleton includes a fuselage longitudinal reinforcement beam, a fuselage longitudinal beam is arranged below the fuselage longitudinal reinforcement beam, a cockpit reinforcement beam is arranged on the front side of the fuselage longitudinal beam, a central wing longitudinal reinforcement beam is arranged above the fuselage longitudinal reinforcement beam, the central wing longitudinal reinforcement beam is located at the junction of the wing and the fuselage, and the central wing longitudinal reinforcement beams are fixedly connected by bolts.

[0009] There are two longitudinal fuselage beams distributed at the bottom of the fuselage. The longitudinal fuselage beams consist of a, b, c, and d, and every two adjacent beams are detachably connected by films and bolts.

[0010] The cockpit reinforcement beam includes a transverse beam, which is fixedly connected to the fuselage longitudinal reinforcement beam by bolts. The inner cavity of the transverse beam is provided with a longitudinal beam, and a front landing gear mounting hole is opened on one side of the longitudinal beam.

[0011] The central wing longitudinal reinforcement beam includes a wing front beam, a wing rear beam is provided on one side of the wing front beam, and a plurality of longitudinal frames are fixedly connected to one side of the wing front beam and the wing rear beam. A transverse frame is provided on the surface of the longitudinal frame, and a wing mounting hole is provided on one side of the wing front beam and the wing rear beam.

[0012] The annular frame includes two annular frames 1, two annular frames 2 are arranged on the rear side of the annular frame 1, and two reinforcing ribs are fixedly connected between the two annular frames 2.

[0013] The combined skeleton includes an upper combined frame, a lower combined frame is arranged below the upper combined frame, and a main combined frame is arranged between the upper and lower combined frames. The combined skeleton is fixedly connected to the outer skin and the inner skin by adhesive bonding and bolts. The upper and lower combined frames are machined from 7075 high-strength aluminum alloy. The main combined frame is formed by paving carbon fiber composite materials and then curing them in an autoclave. The surface of the lower combined frame is provided with landing gear connection holes.

[0014] The fuselage skin includes an outer skin, an inner skin is provided in an inner cavity of the outer skin, the combined frame is fixedly connected to the outer skin and the inner skin by bolts, the outer skin covers the outer side of the fuselage frame, the inner skin covers the inner side of the fuselage frame, and the inner skin and the fuselage frame are connected by adhesive bonding bolts, and the surfaces of the outer skin and the inner skin are both provided with an opening component;

[0015] The outer skin includes a left outer skin, a right outer skin is provided on one side of the left outer skin, and a fuselage radome is provided on one side of the left outer skin and the right outer skin.

[0016] The inner skin includes a left inner skin, and a right inner skin is provided on one side of the left inner skin.

[0017] The opening assembly includes an aircraft door opening opened on the surface of the outer skin and the inner skin, an aircraft glass opening is commonly opened on the surface of the outer skin and the inner skin, a protruding frame is provided on the surface of the inner skin, the protruding frame is used in conjunction with the outer skin, and the protruding frame is connected to the outer skin by bolts and films, the protruding frame is used in conjunction with the aircraft door opening, a front landing gear opening is opened at the bottom of the outer skin, two rear landing gear openings are opened at the bottom of the outer skin, a cooling system exhaust port is opened at the bottom of the outer skin and located on the right side of the rear landing gear opening, and a tail disassembly and maintenance opening is opened on the right side of the bottom of the inner skin.

[0018] The fuselage partition includes a fuselage ground partition, a fuselage cabin partition, a central wing partition, an equipment cabin partition and a cockpit partition. The cockpit partition, the fuselage ground partition, the fuselage cabin partition and the equipment cabin partition are arranged at the bottom of the fuselage skin inner cavity from front to back, and the central wing partition is arranged above the fuselage skin. The fuselage ground partition, the fuselage cabin partition, the central wing partition and the equipment cabin partition are all made of carbon fiber prepreg and foam sandwich paving and curing. The fuselage ground partition is placed above the transverse frame and is connected to it to form an internal platform of the cabin.

[0019] Furthermore, as a preferred embodiment of the present invention, three pilot vision windows are provided on the surface of the fuselage skin, and five passenger vision windows are provided on the surface of the fuselage skin.

[0020] Furthermore, as a preferred embodiment of the present invention, a track groove is provided on the top of each of a and b, and a plurality of limit grooves are provided on both sides of the inner cavity of the track groove.

[0021] Furthermore, as a preferred embodiment of the present invention, the longitudinal reinforcement beams of the fuselage are distributed at the top and bottom of the fuselage, and the longitudinal reinforcement beams of the fuselage are fixedly connected to the longitudinal reinforcement beams of the central wing by bolts.

[0022] Furthermore, as a preferred embodiment of the present invention, the annular skeleton 1 is formed by paving carbon fiber composite materials and then curing them in an autoclave, and the annular skeleton 2 is machined from 7075 high-strength aluminum alloy.

[0023] Furthermore, as a preferred embodiment of the present invention, the fuselage skin is formed by laying carbon fiber prepreg and foam core on a mold and curing.

[0024] Furthermore, as a preferred embodiment of the present invention, a plurality of tail wing docking holes are provided on one side of the annular skeleton 2, and a plurality of skin connection holes are provided on the surface of the annular skeleton 2.

[0025] Furthermore, as a preferred embodiment of the present invention, the surfaces of the outer skin and the inner skin are both provided with aircraft charging openings.

[0026] Furthermore, as a preferred embodiment of the present invention, the annular skeleton is fixedly connected to the outer skin by bolts.

[0027] Beneficial effects: The technical solution of the present application has the following technical effects: the present invention has the advantages of inner and outer skin design, expanded visual range, integrated slide rail, increased support effect and internal partition design, and provides an innovative inner and outer double-layer skin fuselage structure design, aiming to improve the overall strength, rigidity and lightweight effect of the fuselage. The outer skin is made of carbon fiber prepreg and the inner skin is made of 7075 high-strength aluminum alloy to ensure the high strength of the outer layer and the lightweight of the inner layer. Adhesive bonding and bolt connection are used to ensure the close connection between the two layers of skin and the efficiency of load transfer; then, under the action of the driver's field of view window and the passenger field of view window, the fuselage structure design of the visual range is optimized, and the observation range and safety of the driver and passengers are improved; further, the integrated slide rail is large The beam design tightly combines the beam and the slide rail to form an integral structure, which not only improves the stability and load-bearing capacity of the beam, simplifies the assembly process inside the fuselage, and reduces manufacturing costs. At the same time, the integrated slide rail beam is also adjustable and can be flexibly adjusted according to different flight conditions and passenger needs; further, the optimized internal fuselage frame beam structure combines the advantages of transverse, longitudinal, circumferential and combined frame beams to form a complete and efficient support system to improve the durability and reliability of the overall structure; finally, in order to meet the passengers' needs for comfort and privacy, the internal fuselage partition is designed. The partition is made of lightweight, high-strength material and can be flexibly adjusted according to the needs of passengers to improve the overall quality of the aircraft. BRIEF DESCRIPTION OF THE DRAWINGS

[0028] The accompanying drawings are used to provide a further understanding of the present invention and constitute a part of the specification. Together with the embodiments of the present invention, they are used to explain the present invention and do not constitute a limitation of the present invention. In the accompanying drawings:

[0029] Figure 1 It is a schematic diagram of the structure of the present invention;

[0030] Figure 2 This is a three-dimensional schematic diagram of the partial structure of the fuselage skin of the present invention in a disassembled state;

[0031] Figure 3 This is a bottom view of the local structure of the fuselage skin of the present invention;

[0032] Figure 4 This is a schematic diagram of the three-dimensional structure of the fuselage skeleton of the present invention;

[0033] Figure 5 This is a schematic diagram of the three-dimensional structure of the longitudinal skeleton of the present invention;

[0034] Figure 6 It is a three-dimensional schematic diagram of the partial structure of the longitudinal beam of the fuselage of the present invention;

[0035] Figure 7 For the present invention Figure 6 A partial enlarged view of middle A;

[0036] Figure 8 This is a three-dimensional schematic diagram of the local structure of the cockpit reinforcement beam of the present invention;

[0037] Figure 9 It is a schematic diagram of the partial three-dimensional structure of the fuselage partition, annular frame and central wing longitudinal reinforcement beam of the present invention;

[0038] Figure 10 It is a three-dimensional schematic diagram of the local structure of the annular skeleton of the present invention;

[0039] Figure 11 It is a schematic diagram of the three-dimensional structure of the combined skeleton of the present invention.

[0040] In the figure, the meanings of the reference numerals are as follows: 1. fuselage frame; 11. longitudinal frame; 111. fuselage longitudinal beam; 112. fuselage longitudinal reinforcement beam; 113. cockpit reinforcement beam; 1131. transverse beam; 1132. longitudinal beam; 1133. front landing gear mounting hole; 114. central wing longitudinal reinforcement beam; 1141. wing front beam; 1142. wing rear beam; 1143. longitudinal frame; 1144. transverse frame; 1145. wing mounting hole; 12. circumferential frame; 121. circumferential frame 1; 122. circumferential frame 2; 123. reinforcement rib; 124. tail docking hole; 125. skin connection hole; 13. transverse frame; 14. combined frame; 141. upper combined frame; 142. lower combined frame; 143. total combined frame; 14 4. Landing gear connection hole; 2. Fuselage skin; 21. Outer skin; 211. Left outer skin; 212. Right outer skin; 213. Fuselage radome; 22. Inner skin; 221. Left inner skin; 222. Right inner skin; 23. Opening assembly; 231. Aircraft door opening; 232. Aircraft window opening; 233. Protruding frame; 234. Front landing gear opening; 235. Aircraft charging opening; 236. Rear landing gear opening; 237. Cooling system exhaust vent; 238. Tail disassembly and maintenance opening; 3. Fuselage partition; 31. Fuselage ground partition; 32. Fuselage compartment partition; 33. Center wing partition; 34. Equipment compartment partition; 35. Cockpit partition; 4. Pilot's field of view window; 5. Passenger field of view window; 6. Track groove; 7. Limit groove. DETAILED DESCRIPTION

[0041] The technical solutions in the embodiments of the present invention will be clearly and completely described below in conjunction with the drawings in the embodiments of the present invention. In order to better understand the technical content of the present invention, specific embodiments are cited and explained in conjunction with the drawings as follows. Various aspects of the present invention are described in this disclosure with reference to the drawings, which show many illustrative embodiments. It should be understood that the various concepts and embodiments introduced above, as well as those described in more detail below, can be implemented in any of many ways. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without making creative work are within the scope of protection of the present invention.

[0042] As attached Figure 1 To the attached Figure 11 As shown: This embodiment provides an eVTOL aircraft fuselage structure, including: a fuselage frame 1, a fuselage skin 2 is provided on the surface of the fuselage frame 1, a fuselage partition 3 is provided on the inner wall of the fuselage skin 2, and the fuselage frame 1, the fuselage skin 2 and the fuselage partition 3 together constitute the fuselage.

[0043] The fuselage frame 1 includes a longitudinal frame 11, a circumferential frame 12, a transverse frame 13 and a combined frame 14. The longitudinal frame 11 is located at the bottom of the fuselage, the upper part of the fuselage, and the middle and rear part of the fuselage. The longitudinal frame 11 at the bottom of the fuselage and the longitudinal frame 11 at the upper part of the fuselage are both machined from 7075 high-strength aluminum alloy. The longitudinal frames 11 at other positions are formed by laying carbon fiber composite materials and then hot-autoclaving. The circumferential frame 12 is located at the tail of the fuselage. The transverse frame 13 is located at the bottom of the fuselage and the joint between the wing and the fuselage. The transverse frame 13 is mostly machined from 7075 high-strength aluminum alloy, and some are formed by laying carbon fiber composite materials and then hot-autoclaving. The combined frame 14 is located at the front and center of the fuselage.

[0044] The longitudinal skeleton 11 includes a fuselage longitudinal reinforcement beam 112, a fuselage longitudinal beam 111 is arranged below the fuselage longitudinal reinforcement beam 112, a cockpit reinforcement beam 113 is arranged on the front side of the fuselage longitudinal beam 111, and a central wing longitudinal reinforcement beam 114 is arranged above the fuselage longitudinal reinforcement beam 112. The central wing longitudinal reinforcement beam 114 is located at the junction of the wing and the fuselage, and the central wing longitudinal reinforcement beams 114 are fixedly connected by bolts.

[0045] There are two longitudinal fuselage beams 111 distributed at the bottom of the fuselage. The longitudinal fuselage beams 111 are composed of a, b, c, and d, and every two adjacent beams are detachably connected by films and bolts.

[0046] The cockpit reinforcement beam 113 includes a transverse beam 1131 , which is fixedly connected to the fuselage longitudinal reinforcement beam 112 by bolts. A longitudinal beam 1132 is provided in the inner cavity of the transverse beam 1131 , and a front landing gear mounting hole 1133 is provided on one side of the longitudinal beam 1132 .

[0047] The central wing longitudinal reinforcement beam 114 includes a wing front beam 1141, and a wing rear beam 1142 is provided on one side of the wing front beam 1141. One side of the wing front beam 1141 and the wing rear beam 1142 are fixedly connected to a plurality of longitudinal frames 1143, and the surface of the longitudinal frame 1143 is provided with a transverse frame 1144. One side of the wing front beam 1141 and the wing rear beam 1142 are provided with wing mounting holes 1145.

[0048] The annular frame 12 includes two annular frames 121 . Two annular frames 122 are disposed on the rear side of the annular frame 121 . Two reinforcing ribs 123 are fixedly connected between the two annular frames 122 .

[0049] The combined skeleton 14 includes an upper combined frame 141, a lower combined frame 142 is arranged below the upper combined frame 141, and a main combined frame 143 is arranged between the upper combined frame 141 and the lower combined frame 142. The combined skeleton 14 is fixedly connected to the outer skin 21 and the inner skin 22 by adhesive bonding and bolts. The upper combined frame 141 and the lower combined frame 142 are machined from 7075 high-strength aluminum alloy. The main combined frame 143 is formed by paving carbon fiber composite materials and then curing them in an autoclave. The surface of the lower combined frame 142 is provided with landing gear connection holes 144.

[0050] The fuselage skin 2 includes an outer skin 21. The circumferential frame 12 is fixedly connected to the outer skin 21 by bolts. The inner skin 22 is provided in the inner cavity of the outer skin 21. The combined frame 14 is fixedly connected to the outer skin 21 and the inner skin 22 by bolts. The outer skin 21 covers the outer side of the fuselage frame 1, and the inner skin 22 covers the inner side of the fuselage frame 1. The inner skin 22 and the fuselage frame 1 are connected by glue and bolts. The surfaces of the outer skin 21 and the inner skin 22 are both provided with an opening component 23.

[0051] The outer skin 21 includes a left outer skin 211 , a right outer skin 212 is provided on one side of the left outer skin 211 , and a fuselage radome 213 is provided on one side of both the left outer skin 211 and the right outer skin 212 .

[0052] The inner skin 22 includes a left inner skin 221 , and a right inner skin 222 is provided on one side of the left inner skin 221 .

[0053] The opening assembly 23 includes an aircraft cabin door opening 231 opened on the surface of the outer skin 21 and the inner skin 22. The surfaces of the outer skin 21 and the inner skin 22 are jointly opened with an aircraft glass opening 232. The surface of the inner skin 22 is provided with a protruding frame 233. The protruding frame 233 is used in conjunction with the outer skin 21, and the protruding frame 233 is connected to the outer skin 21 by bolts and films. The protruding frame 233 is used in conjunction with the aircraft cabin door opening 231. A front landing gear opening 234 is opened at the bottom of the outer skin 21. Two rear landing gear openings 236 are opened at the bottom of the outer skin 21. A cooling system exhaust vent 237 is opened at the bottom of the outer skin 21 and on the right side of the rear landing gear opening 236. A tail disassembly and maintenance opening 238 is opened on the right side of the bottom of the inner skin 22.

[0054] The fuselage partition 3 includes a fuselage ground partition 31, a fuselage cabin partition 32, a central wing partition 33, an equipment cabin partition 34 and a cockpit partition 35. The cockpit partition 35, the fuselage ground partition 31, the fuselage cabin partition 32 and the equipment cabin partition 34 are arranged at the bottom of the inner cavity of the fuselage skin 2 from front to back, and the central wing partition 33 is arranged above the fuselage skin 2. The fuselage ground partition 31, the fuselage cabin partition 32, the central wing partition 33 and the equipment cabin partition 34 are all made of carbon fiber prepreg and foam sandwich paving and curing. The fuselage ground partition 31 is placed above the transverse frame 13 and is connected to it to form an internal platform of the cabin.

[0055] Specifically, three pilot vision windows 4 are provided on the surface of the fuselage skin 2 , and five passenger vision windows 5 are provided on the surface of the fuselage skin 2 .

[0056] In this embodiment, by cooperating with the driver's field of view window 4 and the passenger's field of view window 5, the driver's field of view window 4 provides the driver with three window visual ranges, and the passenger's field of view window 5 provides the passenger with five window visual ranges.

[0057] Specifically, a track groove 6 is provided on the top of each of a and b, and a plurality of limiting grooves 7 are provided on both sides of the inner cavity of the track groove 6 .

[0058] In this embodiment, the coordinated use of the track groove 6 and the limiting groove 7 allows the seat to be connected thereto and to slide and be fixed as required, thereby improving the flexibility of use.

[0059] Specifically, the fuselage longitudinal reinforcement beams 112 are distributed at the top and bottom of the fuselage, and the fuselage longitudinal reinforcement beams 112 are fixedly connected to the central wing longitudinal reinforcement beams 114 by bolts.

[0060] In this embodiment, by disposing the fuselage longitudinal reinforcement beam 112 and utilizing its position setting and strengthening its connection with the central wing longitudinal reinforcement beam 114, it can be used to bear the fuselage load and improve the fuselage stability.

[0061] Specifically, the annular frame 121 is formed by laying carbon fiber composite materials and then curing them in an autoclave, and the annular frame 122 is machined from 7075 high-strength aluminum alloy.

[0062] In this embodiment: through the coordinated use of annular frame 121 and annular frame 2 122, annular frame 121 uses carbon fiber composite material to reduce the overall weight, and annular frame 2 122 uses 7075 high-strength aluminum alloy to ensure the strength and reliability of connecting the tail wing and transmitting loads. The carbon fiber composite material provides excellent mechanical properties and thermal stability, while the 7075 aluminum alloy meets specific structural requirements with its high strength and good machinability.

[0063] Specifically, the fuselage skin 2 is formed by laying carbon fiber prepreg and foam core on a mold and curing them.

[0064] In this embodiment: through the arrangement of the fuselage skin 2, the carbon fiber prepreg provides excellent strength and stiffness, while the foam core achieves lightweighting, and the combination of the two forms a fuselage skin 2 that is both strong and lightweight.

[0065] Specifically, a plurality of tail wing docking holes 124 are opened on one side of the annular frame 122 , and a plurality of skin connection holes 125 are opened on the surface of the annular frame 122 .

[0066] In this embodiment: through the coordinated use of the tail docking hole 124 and the skin connection hole 125, under the connection of the tail docking hole 124, it is connected to the tail section through bolt connection, and the load of the horizontal vertical tail section is transferred to the fuselage. Under the setting of the skin connection hole 125, the circumferential frame 12 is connected to the fuselage skin 2, thereby improving the stability of the fuselage skin 2.

[0067] Specifically, aircraft charging openings 235 are provided on the surfaces of the outer skin 21 and the inner skin 22 .

[0068] In this embodiment, the aircraft charging opening 235 is provided to provide a channel for charging the aircraft, thereby improving the rationality of the structural layout.

[0069] The working principle and usage process of this invention: The fuselage skin 2, serving as the aircraft's key protective layer, consists of an outer skin 21 and an inner skin 22. Both are constructed from carbon fiber prepreg and foam sandwich material, laid on a mold and cured. This structure not only achieves lightweighting but also ensures high strength and excellent impact resistance. The outer skin 21 fits tightly against the outside of the fuselage frame 1, while the inner skin 22 covers the inside. The two are tightly connected by gluing and screwing, forming a closed frame structure that provides comprehensive protection for the fuselage's occupants and equipment. The fuselage skin 2 also maintains the fuselage's shape and reduces air resistance.

[0070] Various openings are designed on the surface of the outer skin 21 and the inner skin 22 to meet the various functional requirements of the fuselage, such as installing glass, doors, landing gear placement, exhaust vents, inspection holes and charging ports. The ingenious design of these openings not only ensures the integrity of the fuselage, but also meets various requirements in actual operation.

[0071] To further enhance the load-bearing capacity of the fuselage, the protruding frame 233 is tightly attached to the outer skin 21 and then firmly connected to the fuselage through bolt connection and bonding technology. This design not only improves the overall strength of the fuselage, but also provides a wider field of view for the pilot and passengers. The pilot's field of view window 4 provides the pilot with a three-window visual range, while the passenger field of view window 5 allows passengers to enjoy a wide field of view through five windows.

[0072] The fuselage frame 1 serves as the "backbone" supporting the entire fuselage and is composed of a longitudinal frame 11, a circumferential frame 12, a transverse frame 13 and a combined frame 14. These frames are precisely connected to form the aircraft's skeleton system, which improves the fuselage's torsional resistance. Due to different load distribution, the longitudinal frames 11 at the bottom and top of the fuselage are mostly machined from 7075 high-strength aluminum alloy, while the longitudinal frames 11 at other locations are made of carbon fiber composite materials that are laid out and then autoclaved for curing. There are two longitudinal frames 11, which are distributed at the bottom of the fuselage. They have four beams consisting of a, b, c, and d, which are connected by films and bolts. The crossbeam 1131 and the longitudinal beam 1132 are connected by bolts to form a closed structure, which bears the load of the cockpit door equipment and strengthens the bearing capacity of the fuselage. The longitudinal reinforcement beam 112 of the fuselage is mainly distributed at the top and bottom of the fuselage, and bears the fuselage load by connecting with the cross frame 1144. The longitudinal reinforcement beam 114 of the central wing is mainly located at the junction of the wing and the body, and a stable frame structure is formed after being connected by bolts. The surfaces of the beam 1141 and the rear beam 1142 of the wing are respectively provided with wing mounting holes 1145. After being connected with the wing bolts, the load is transferred from the wing to the fuselage. The transverse frame 13 is mainly located at the bottom of the fuselage, at the junction of the wing and the body, and is connected to the longitudinal frame 11. It provides stable support for the bottom of the fuselage, bears the load of the occupants, and provides a landing gear installation position and transfers the landing gear load. In view of the different load distribution, the transverse frame 13 is mostly machined from 7075 high-strength aluminum alloy, and some are made of carbon fiber composite materials after paving and autoclaving. The circumferential frame 12 is mainly located at the tail of the fuselage because the aerodynamic load of the tail and the load of the tail are relatively small. Therefore, the annular skeletons 12 are relatively sparsely distributed and connected to the outer skin 21 to bear the load, prevent the skin from buckling, and maintain the aircraft's aerodynamic shape. Given the small load, the annular skeleton 121 is formed by laying out carbon fiber composite materials and then curing them in an autoclave. The annular skeleton 122 is machined from 7075 high-strength aluminum alloy to connect to the tail and transfer the load. Two reinforcing ribs 123 are connected in the middle to increase the structural bending and torsional resistance. The annular skeleton 122 is provided with multiple tail docking holes 124 and skin connection holes 125. It is connected to the tail section by bolts, transferring the load of the horizontal vertical tail section to the fuselage.The combined frame 14 is mainly located in the front section and the center of the fuselage, and is connected to the outer skin 21 and the inner skin 22. It is a closed frame structure. Taking the position of connecting the landing gear as an example: the combined frame 14 includes an upper combined frame 141, a lower combined frame 142 and a total combined frame 143, which are then connected to the outer skin 21 and the inner skin 22 by adhesive bonding bolts to form a closed frame structure that withstands torsional loads and shear loads. Since the upper and lower positions need to connect the wings and the landing gear, the load transmitted is relatively large. Therefore, the upper combined frame 141 and the lower combined frame 142 are mostly made of 7075 high-strength aluminum alloy by machining, and the total combined frame 143 is made of 7075 high-strength aluminum alloy by machining. The assembly frame 143 is typically formed by laying out carbon fiber composite materials and then autoclaving and curing them. Finally, the fuselage bulkhead 3 includes the fuselage floor bulkhead 31, fuselage cabin bulkhead 32, center wing bulkhead 33, equipment bay bulkhead 34, and cockpit bulkhead 35. The fuselage floor bulkhead 31, fuselage cabin bulkhead 32, center wing bulkhead 33, and equipment bay bulkhead 34 are not primary load-bearing components and are therefore typically constructed by laying out and curing carbon fiber prepreg and foam core. The fuselage floor bulkhead 31 is laid flat on and connected to the transverse frame 13 to form an in-cabin platform, while also leaving space for the slide rails of the fuselage longitudinal beams 111.

[0073] It should be noted that, in this document, relational terms such as first and second, etc. are merely used to distinguish one entity or operation from another entity or operation, and do not necessarily require or imply any actual relationship or order between these entities or operations.

[0074] While the present invention has been disclosed above with reference to preferred embodiments, this is not intended to limit the present invention. Persons skilled in the art will readily appreciate that various modifications and variations can be made without departing from the spirit and scope of the present invention. Therefore, the scope of protection of the present invention shall be determined by the claims.

Claims

1. An eVTOL aircraft fuselage structure, comprising: A fuselage frame (1), characterized in that: a fuselage skin (2) is provided on the surface of the fuselage frame (1), a fuselage partition (3) is provided on the inner wall of the fuselage skin (2), and the fuselage frame (1), the fuselage skin (2) and the fuselage partition (3) together constitute a fuselage; The fuselage frame (1) includes a longitudinal frame (11), a circumferential frame (12), a transverse frame (13) and a combined frame (14). The longitudinal frame (11) is located at the bottom of the fuselage, the upper part of the fuselage and the middle and rear part of the fuselage. The longitudinal frame (11) located at the bottom of the fuselage and the longitudinal frame (11) located at the upper part of the fuselage are both machined from 7075 high-strength aluminum alloy. The longitudinal frames (11) at other positions are formed by paving carbon fiber composite materials and then autoclaving. The circumferential frame (12) is located at the tail of the fuselage. The transverse frame (13) is located at the bottom of the fuselage and the joint between the wing and the fuselage. The longitudinal skeleton (11) includes a fuselage longitudinal reinforcement beam (112), a fuselage longitudinal beam (111) is provided below the fuselage longitudinal reinforcement beam (112), a cockpit reinforcement beam (113) is provided on the front side of the fuselage longitudinal beam (111), a central wing longitudinal reinforcement beam (114) is provided above the fuselage longitudinal reinforcement beam (112), the central wing longitudinal reinforcement beam (114) is located at the wing-body junction, and the central wing longitudinal reinforcement beams (114) are fixedly connected by bolts; The number of the fuselage longitudinal beams (111) is two and they are distributed at the bottom of the fuselage; The cockpit reinforcement beam (113) includes a transverse beam (1131), the transverse beam (1131) is fixedly connected to the fuselage longitudinal reinforcement beam (112) by bolts, the inner cavity of the transverse beam (1131) is provided with a longitudinal beam (1132), and one side of the longitudinal beam (1132) is provided with a front landing gear mounting hole (1133); The central wing longitudinal reinforcement beam (114) includes a wing front beam (1141), a wing rear beam (1142) is provided on one side of the wing front beam (1141), a plurality of longitudinal frames (1143) are fixedly connected to one side of each of the wing front beam (1141) and the wing rear beam (1142), a transverse frame (1144) is provided on the surface of each of the longitudinal frames (1143), and a wing mounting hole (1145) is provided on one side of each of the wing front beam (1141) and the wing rear beam (1142); The annular frame (12) includes two annular frames (121), two annular frames (122) are provided on the rear side of the annular frame (121), and two reinforcing ribs (123) are fixedly connected between the two annular frames (122); The combined skeleton (14) includes an upper combined frame (141), a lower combined frame (142) is provided below the upper combined frame (141), a main combined frame (143) is provided between the upper combined frame (141) and the lower combined frame (142), the upper combined frame (141) and the lower combined frame (142) are machined from 7075 high-strength aluminum alloy, the main combined frame (143) is formed by paving carbon fiber composite materials and then autoclaving and curing, and a landing gear connection hole (144) is provided on the surface of the lower combined frame (142); The fuselage skin (2) includes an outer skin (21), an inner skin (22) is provided in an inner cavity of the outer skin (21), the combined frame (14) is fixedly connected to the outer skin (21) and the inner skin (22) by bolts, and the outer skin (21) covers the outer side of the fuselage frame (1), and the inner skin (22) covers the inner side of the fuselage frame (1), and the surfaces of the outer skin (21) and the inner skin (22) are both provided with an opening component (23); The outer skin (21) includes a left outer skin (211), a right outer skin (212) is provided on one side of the left outer skin (211), and a fuselage radome (213) is provided on one side of the left outer skin (211) and the right outer skin (212); The inner skin (22) includes a left inner skin (221), and a right inner skin (222) is provided on one side of the left inner skin (221); The opening assembly (23) includes an aircraft cabin door opening (231) opened on the surface of the outer skin (21) and the inner skin (22), the surfaces of the outer skin (21) and the inner skin (22) are jointly provided with an aircraft glass opening (232), the surface of the inner skin (22) is provided with a protruding frame (233), the protruding frame (233) is used in conjunction with the outer skin (21), and the protruding frame (233) is connected to the outer skin (21) by bolts and films. The protruding frame (233) is used in conjunction with the aircraft door opening (231); a front landing gear opening (234) is provided at the bottom of the outer skin (21); two rear landing gear openings (236) are provided at the bottom of the outer skin (21); a cooling system exhaust port (237) is provided at the bottom of the outer skin (21) and on the right side of the rear landing gear opening (236); and a tail disassembly and maintenance opening (238) is provided at the right side of the bottom of the inner skin (22); The fuselage partition (3) comprises a fuselage ground partition (31), a fuselage cabin partition (32), a central wing partition (33), an equipment cabin partition (34) and a cockpit partition (35). The cockpit partition (35), the fuselage ground partition (31), the fuselage cabin partition (32) and the equipment cabin partition (34) are sequentially arranged at the bottom of the inner cavity of the fuselage skin (2) from front to back. The central wing partition (33) is arranged above the fuselage skin (2). The fuselage ground partition (31), the fuselage cabin partition (32), the central wing partition (33) and the equipment cabin partition (34) are all made of carbon fiber prepreg and foam sandwich paving and curing. The fuselage ground partition (31) is placed above the transverse frame (13) and is connected thereto to form an inner platform of the cabin.

2. The eVTOL aircraft fuselage structure according to claim 1, characterized in that: The surface of the fuselage skin (2) is provided with three pilot vision windows (4), and the surface of the fuselage skin (2) is provided with five passenger vision windows (5).

3. The eVTOL aircraft fuselage structure according to claim 1, characterized in that: The fuselage longitudinal reinforcement beams (112) are distributed at the top and bottom of the fuselage, and the fuselage longitudinal reinforcement beams (112) are fixedly connected to the central wing longitudinal reinforcement beams (114) via bolts.

4. The eVTOL aircraft fuselage structure according to claim 1, characterized in that: The annular frame 1 (121) is formed by laying carbon fiber composite materials and then curing them in an autoclave, and the annular frame 2 (122) is formed by machining 7075 high-strength aluminum alloy.

5. The eVTOL aircraft fuselage structure according to claim 1, characterized in that: The fuselage skin (2) is formed by paving carbon fiber prepreg and foam core on a mold and curing.

6. The eVTOL aircraft fuselage structure according to claim 1, characterized in that: A plurality of tail wing docking holes (124) are provided on one side of the annular frame 2 (122), and a plurality of skin connection holes (125) are provided on the surface of the annular frame 2 (122).

7. The eVTOL aircraft fuselage structure according to claim 1, characterized in that: Aircraft charging openings (235) are provided on the surfaces of the outer skin (21) and the inner skin (22).

8. The eVTOL aircraft fuselage structure according to claim 1, characterized in that: The annular skeleton (12) is fixedly connected to the outer skin (21) via bolts.

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