EVTOL aircraft fuselage structure

Abstract

The invention relates to the technical field of airplanes, in particular to an eVTOL aircraft fuselage structure which comprises a fuselage framework, fuselage skin is arranged on the surface of the fuselage framework, fuselage partition plates are arranged on the inner wall of the fuselage skin, and the fuselage framework, the fuselage skin and the fuselage partition plates jointly form a fuselage. The aircraft body structure has the advantages of inner and outer skin design, visible range expansion, integrated sliding rail, supporting effect improvement and inner partition plate design, an innovative inner and outer double-layer skin aircraft body structure design is provided, and the purposes of improving the overall strength, rigidity and lightweight effect of the aircraft body are achieved; the outer skin is made of carbon fiber prepreg, the inner skin is made of 7075 high-strength aluminum alloy so as to ensure high strength of the outer layer and light weight of the inner layer, and the two layers of skins are connected through glue joint and bolts so as to ensure tight combination and load transmission efficiency between the two layers of skins; and then, under the action of the driver view window and the passenger view window, the structural design of the vehicle body in 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's passengers and equipment. On the one hand, it protects the safety of passengers, and on the other hand, it transfers the aircraft's load, maintains the aircraft's streamlined shape, and reduces automatic resistance. At the same time, the fuselage design 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 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 flight performance and load-carrying capacity of the aircraft; 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 and heavy 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 and contains 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, which has the advantages of inner and outer skin design, expanded visual range, integrated slide rails, increased support effect and internal partition design, and solves the problems raised by the above-mentioned background technology.

[0006] To achieve the above-mentioned purpose, the present invention provides the following technical solutions: an eVTOL aircraft fuselage structure, comprising: a fuselage frame, a fuselage skin is arranged on the surface of the fuselage frame, a fuselage partition is arranged 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, and 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, and 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, most of which are 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 main beam is arranged below the fuselage longitudinal reinforcement beam, a cockpit reinforcement beam is arranged on the front side of the fuselage longitudinal main 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 body, and the central wing longitudinal reinforcement beams are fixedly connected by bolts.

[0009] The number of the longitudinal fuselage beams is two and they are distributed at the bottom of the fuselage. The longitudinal fuselage beams are composed 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, the transverse beam 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-body front beam, a wing-body rear beam is arranged on one side of the wing-body front beam, a plurality of longitudinal frames are fixedly connected to one side of the wing-body front beam and the wing-body rear beam, a transverse frame is arranged on the surface of the longitudinal frame, and a wing mounting hole is opened on one side of the wing-body front beam and the wing-body rear beam.

[0012] The annular frame comprises 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, 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 in an autoclave, and a landing gear connection hole is opened on the surface of the lower combined frame.

[0014] The fuselage skin includes an outer skin, the inner cavity of the outer skin is provided with an inner skin, the combined frame is fixedly connected with the outer skin and the inner skin by bolts, and 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 glue-joined bolts, and the surfaces of the outer skin and the inner skin are both provided with opening components;

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

[0017] The opening assembly includes an aircraft door opening opened on the surfaces of an outer skin and an inner skin, an aircraft glass opening is commonly opened on the surfaces 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 through 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 inner cavity of the fuselage skin from front to back in sequence, 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 connected thereto to form an inner 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 at 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 opened on one side of the annular skeleton two, and a plurality of skin connection holes are opened on the surface of the annular skeleton two.

[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 lightness 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's field of view window, the fuselage structure design of the visual range is optimized to improve the observation range and safety of the driver and passengers. Furthermore, the integrated slide rail is large The beam design closely 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 one-piece 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, an 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 It 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 It is a bottom view of the local structure of the fuselage skin of the present invention;

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

[0033] Figure 5 It 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 local 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 It is a three-dimensional schematic diagram of the local structure of the cockpit reinforcement beam of the present invention;

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

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

[0039] Fig.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, longitudinal beam of fuselage; 112, longitudinal reinforcing beam of fuselage; 113, cockpit reinforcing beam; 1131, transverse beam; 1132, longitudinal beam; 1133, front landing gear mounting hole; 114, central wing longitudinal reinforcing 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 one; 122, circumferential frame two; 123, reinforcing 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 glass 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 vision window; 5. Passenger field of vision 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, in which many illustrative embodiments are shown. 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 in 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 Attachment Fig.11 As shown: This embodiment provides an eVTOL aircraft fuselage structure, including: a fuselage frame 1, a fuselage skin 2 is arranged on the surface of the fuselage frame 1, a fuselage partition 3 is arranged 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 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, and the longitudinal frames 11 at other positions are formed by paving carbon fiber composite materials and then hot-autoclaving. The circumferential frame 12 is located at the tail of the fuselage, and 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 7075, which 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 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. 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 .

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

[0048] The annular frame 12 includes two annular frames 121 , two annular frames 122 are arranged at the rear side of the annular frame 121 , and 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, a general 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 general combined frame 143 is formed by paving carbon fiber composite materials and then curing in an autoclave, and a landing gear connection hole 144 is opened on the surface of the lower combined frame 142.

[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 cavity of the outer skin 21 is provided with an inner skin 22, 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, the inner skin 22 covers the inner side of the fuselage frame 1, and the inner skin 22 and the fuselage frame 1 are connected by glue-joined bolts, and 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 disposed on one side of the left outer skin 211 , and a fuselage radome 213 is disposed on one side of 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 disposed on one side of the left inner skin 221 .

[0053] The opening assembly 23 includes an aircraft 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 through bolts and films. The protruding frame 233 is used in conjunction with the aircraft 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 port 237 is opened at the bottom of the outer skin 21 and located on the right side of the rear landing gear opening 236. A tail disassembly and maintenance opening 238 is opened at 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 connected thereto to form an internal platform of the cabin.

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

[0056] In this embodiment: through the coordinated use of 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 limit 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 the 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 paving 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 materials 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 transferring loads. The carbon fiber composite materials provide 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 rigidity, while the foam core achieves lightweight, and the combination of the two forms a fuselage skin 2 that is both strong and light.

[0065] Specifically, a plurality of tail wing docking holes 124 are provided on one side of the annular frame 122 , and a plurality of skin connection holes 125 are provided 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 connecting hole 125, under the connection of the tail docking hole 124, it is connected to the tail section through a bolt connection, and the load of the vertical tail section is transferred to the fuselage. Under the setting of the skin connecting hole 125, the annular 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: by providing the aircraft charging opening 235, a channel is provided for charging the aircraft, thereby improving the rationality of the structural layout.

[0069] The working principle and use process of the present invention: The fuselage skin 2 is the key protective layer of the aircraft, which is composed of an outer skin 21 and an inner skin 22. Both are made of carbon fiber prepreg and foam sandwich material, which are laid on the mold and cured. This structure not only achieves lightweight, but also ensures high strength and excellent impact resistance. The outer skin 21 fits tightly to the outside of the fuselage frame 1, while the inner skin 22 covers the inside of the frame. The two are closely connected by gluing and screwing to form a closed frame structure, which provides comprehensive protection for the fuselage occupants and equipment. The fuselage skin 2 also plays a role in maintaining the shape of the fuselage and reducing air resistance.

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

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

[0072] The fuselage frame 1 is 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 connected in a precise manner to form the aircraft's frame system, which improves the fuselage's torsion resistance. Due to different load distributions, 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 formed by hot-autoclave curing after carbon fiber composite materials are laid. There are two longitudinal frames 11 distributed at the bottom of the fuselage, and they have four beams consisting of a, b, c, and d, connected by films and bolts. The overall structure of the beam is formed. Since seats and some equipment need to be installed, rail grooves 6 and limit grooves 7 are provided on beams a and b to allow seats to be connected thereto and slide and fix as required. The transverse beam 1131 and the longitudinal beam 1132 are connected by bolts to form a closed structure to bear the load of the cockpit door equipment and enhance 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 transverse frame 1144. The central wing longitudinal reinforcement beam 114 is mainly located at the junction of the wing and the body. After being connected by bolts, a stable frame structure is formed. The surfaces of the beam 1141 and the beam 1142 behind 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 joint of the wing and the body, and is connected to the longitudinal frame 11 to provide stable support for the bottom of the fuselage, bear the load of the occupants, and provide a landing gear installation position and transfer 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 formed by hot-autoclave curing after paving carbon fiber composite materials. 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 skeleton 12 is sparsely distributed and connected with the outer skin 21 to bear the load, prevent the skin from buckling, and maintain the aerodynamic shape of the aircraft. In view of the small load, the annular skeleton 121 is formed by hot-pressing and curing after paving carbon fiber composite materials. The annular skeleton 122 is machined from 7075 high-strength aluminum alloy because it is required to connect the tail and transfer the load. Two reinforcing ribs 123 are connected in the middle to increase the bending and torsion resistance of the structure. The annular skeleton 122 is provided with a plurality of tail docking holes 124 and skin connection holes 125, which are connected to the tail section by bolts to transfer the load of the horizontal vertical tail section to the fuselage.The combined frame 14 is mainly located at the front section and the center of the fuselage, and is connected to the outer skin 21 and the inner skin 22 to form a closed frame structure. Taking the position where the landing gear is connected 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 to withstand torsion load and shear load. Since the upper and lower positions need to connect the wings and the landing gear, the load transmitted is relatively large, so the upper combined frame 141 and the lower combined frame 142 are mostly machined from 7075 high-strength aluminum alloy, and the total combined frame 143 is made of 7075 high-strength aluminum alloy. The frame 143 is mostly made of carbon fiber composite materials and then autoclaved for curing; finally, the fuselage partition 3 includes the fuselage ground partition 31, the fuselage cabin partition 32, the central wing partition 33, the equipment cabin partition 34 and the cockpit partition 35, among which the fuselage ground partition 31, the fuselage cabin partition 32, the central wing partition 33 and the equipment cabin partition 34 are not the main load-bearing parts, so they are mostly made of carbon fiber prepreg and foam sandwich for curing. The fuselage bottom partition 31 is laid flat on the transverse frame 13 and connected to it to form a platform in the cabin, and at the same time, space is reserved for the slide rail channel of the fuselage longitudinal beam 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] Although the present invention has been disclosed as above with preferred embodiments, it is not intended to limit the present invention. A person with ordinary knowledge in the technical field to which the present invention belongs may make various changes and modifications without departing from the spirit and scope of the present invention. Therefore, the protection scope of the present invention shall be determined by the definition of the claims.

Claims

1. An eVTOL aircraft fuselage structure, comprising: A fuselage frame (1), characterized in that: a fuselage skin (2) is arranged on the surface of the fuselage frame (1), a fuselage partition (3) is arranged 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) comprises 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 hot-pressing and curing them; 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 paving carbon fiber composite materials and then hot-pressing and curing them; the combined frame (14) is located at the front end and the center of the fuselage; The longitudinal skeleton (11) comprises 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), 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; The number of the longitudinal fuselage beams (111) is two and they are 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 a film and bolts. The cockpit reinforcement beam (113) comprises 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-body front beam (1141), a wing-body rear beam (1142) is arranged on one side of the wing-body front beam (1141), a plurality of longitudinal frames (1143) are fixedly connected to one side of each of the wing-body front beam (1141) and the wing-body rear beam (1142), a transverse frame (1144) is arranged 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-body front beam (1141) and the wing-body rear beam (1142); The annular frame (12) comprises two annular frames one (121), two annular frames two (122) are arranged on the rear side of the annular frame one (121), and two reinforcing ribs (123) are fixedly connected between the two annular frames two (122); The combined frame (14) comprises an upper combined frame (141), a lower combined frame (142) is arranged below the upper combined frame (141), a main combined frame (143) is arranged between the upper combined frame (141) and the lower combined frame (142), the combined frame (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 in an autoclave, and a landing gear connection hole (144) is opened on the surface of the lower combined frame (142); The fuselage skin (2) includes an outer skin (21), the inner cavity of the outer skin (21) is provided with an inner skin (22), 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), the inner skin (22) covers the inner side of the fuselage frame (1), and the inner skin (22) and the fuselage frame (1) are connected by adhesive bolts, 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 disposed on one side of the left outer skin (211), and a fuselage radome (213) is disposed 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) comprises an aircraft cabin door opening (231) opened on the surface of an outer skin (21) and an 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 a film. The protruding frame (233) is used in conjunction with the aircraft cabin 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 formed by paving and curing carbon fiber prepreg and foam sandwich. The fuselage ground partition (31) is placed above the transverse frame (13) and 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: A track groove (6) is provided on the top of each of a and b, and a plurality of limit grooves (7) are provided on both sides of the inner cavity of the track groove (6).

4. 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) by bolts.

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

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

7. 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).

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

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

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