Solar aircraft

Abstract

The invention relates to the technical field of flight equipment, in particular to a solar aircraft which comprises wings, an aircraft body and a reinforced fiber material, the wings are connected with the aircraft body, each wing comprises a main beam body, outer main beams and carbon plates, the outer main beams are arranged on the two sides of the main beam body in the extending direction, and the carbon plates are arranged on the outer main beams. Carbon plates are arranged on the main beam body and the outer main beam in the height direction of the main beam body, the outer main beam is made of a composite material, the main beam body, the outer main beam and the carbon plates form a main beam, and a reinforced fiber material is wound on the peripheral surface of the main beam. The invention provides an aircraft which can reduce the weight, improve the structural strength and shorten the manufacturing period.

Description

Technical Field

[0001] The present application relates to the technical field of flight equipment, and in particular to a solar-powered aircraft. Background Art

[0002] Solar aircraft is a new type of aircraft that uses solar radiation as a propulsion energy source. In recent years, it has attracted widespread attention in the aviation field. As the world pays more and more attention to environmental protection and sustainable energy utilization, solar aircraft, with its significant advantages of zero pollution and renewable energy, has shown great development potential and application prospects.

[0003] In order to lay a sufficient number of solar panels, the wing area of solar aircraft is usually large, which makes the wing face severe challenges when bearing flight and load. It is necessary to ensure that the wing has sufficient strength to support the overall weight of the aircraft and cope with various forces during flight, and to reduce the weight of the wing as much as possible to improve the flight performance and load capacity of the aircraft.

[0004] Related art proposes a high-strength wing structure using light wood as the wing beam, but the structure has limited strength and a long production cycle. Summary of the invention

[0005] The present application provides an aircraft that can reduce weight, improve structural strength, and shorten production cycle.

[0006] The solar powered aircraft according to the embodiment of the present invention comprises:

[0007] The wing and the fuselage are connected, the wing comprises a main beam body, an outer main beam and a carbon sheet, both sides of the main beam body in the extension direction are provided with an outer main beam, the main beam body and the outer main beam are provided with a carbon sheet in the height direction of the main beam body, wherein the material of the outer main beam is a composite material;

[0008] Reinforced fiber material, the main beam body, the outer main beam and the carbon sheet constitute the main beam, and the outer peripheral surface of the main beam is wound with reinforced fiber material.

[0009] The present application provides an aircraft that can reduce weight, improve structural strength, and shorten production cycle.

[0010] In some embodiments, the main beam body is made of polymethacrylimide foam, the outer main beam is made of polyaramid fiber paper honeycomb material, and the reinforcing fiber material is Kevlar fiber.

[0011] In some embodiments, the outer main beam includes a web, a plurality of strut members, and main beam flanges. The web is provided with main beam flanges on both sides in the height direction of the main beam body. A plurality of the strut members are arranged on the web and are sequentially connected in the extending direction of the main beam body. Two ends of the strut members in the height direction of the main beam are respectively connected to the corresponding main beam flanges.

[0012] The web is provided with a plurality of first weight-reducing holes. The plurality of first weight-reducing holes are arranged at intervals in the extending direction of the main beam body on the web, and the cross-sectional area of the first weight-reducing holes gradually increases in the direction away from the fuselage.

[0013] In some embodiments, the outer main beam further includes a connecting member and a web. The connecting member is arranged between the web and the main beam body. The connecting member includes a sleeve and a connecting lug. One end of the sleeve extends into the web and is connected to the web, and the other end of the sleeve extends into the main beam body and is connected to the main beam body. The main beam flanges and the main beam body are both provided with the connecting lugs, and the connecting lugs are provided with threaded holes.

[0014] The material of the connecting lug is carbon fiber, and the sleeve is a carbon nanotube.

[0015] In some embodiments, the wing further includes a plurality of wing ribs, a leading edge strip, a trailing edge strip, and a wing skin. The wing ribs are arranged at intervals in the extending direction of the main beam body, and the wing ribs are sleeved on the main beam body. One end of the wing rib facing the fuselage is provided with a leading edge strip, and the other end of the wing rib is provided with a trailing edge strip.

[0016] The wing skin is arranged on the top and bottom of the wing rib. The wing skin on the top of the wing rib is connected to the wing skin on the bottom at the leading edge, and the connection part is processed into an arc. The material of the wing skin is balsa wood.

[0017] In some embodiments, the wing further includes a strengthening beam and a D-shaped box. The D-shaped box is connected to the strengthening beam and the wing rib, and the bottom of the D-shaped box is connected to the bottom of the fuselage.

[0018] The wing rib is provided with a through hole that cooperates with the strengthening beam. The wing rib cooperates with the through hole to be sleeved on the strengthening beam. The strengthening beam is provided with second weight-reducing holes, and the second weight-reducing holes are provided on both sides of the strengthening beam away from the fuselage.

[0019] In some embodiments, the wing and the fuselage have a preset angle in the height direction of the main beam. The preset angle is 2°. The ratio of the size of the wing in the extending direction of the main beam body to the size of the wing in the width direction of the main beam body is 10.92.

[0020] The dimension of the wing in the extending direction of the main beam body is 4320 mm, and the dimension of the fuselage in the width direction of the main beam body is 996 mm.

[0021] In some embodiments, the fuselage includes a plurality of frames, a plurality of fuselage skins and fiber reinforced ribs. The plurality of frames are spaced apart in the width direction of the main beam body, and the outer peripheral surfaces of the frames are all provided with the fuselage skins. Fiber reinforced ribs are provided inside the frames.

[0022] Wherein, the shape of the frame is square, and the side length of the frame gradually decreases in the direction away from the fuselage.

[0023] In some embodiments, the solar aircraft further includes a carbon fiber board, an aluminum alloy column, a power component and a tail wing component. One end of the frame away from the wing is connected to one end of the carbon fiber board, the other end of the carbon fiber board is connected to one end of the aluminum alloy column, the end of the aluminum alloy column away from the carbon fiber board is connected to the power component, and one end of the tail wing component is connected to the wing.

[0024] In some embodiments, it further includes solar panels and ailerons. The number of the solar panels is multiple. The solar panels are arranged on the top of the wing. Ailerons are provided on both sides of the wing in the extending direction of the main beam body, and the ailerons are arranged on the side of the wing away from the fuselage. BRIEF DESCRIPTION OF THE DRAWINGS

[0025] In order to more clearly illustrate the embodiments of the present application, the following will briefly introduce the drawings required for use in the embodiments. Obviously, the drawings described below are only some embodiments of the present application. For those of ordinary skill in the art, other drawings can be obtained based on these drawings without creative efforts.

[0026] Figure 1 Schematic diagram of the solar aircraft according to the embodiment of the present application.

[0027] Figure 2 Schematic diagram of the main beam body and the outer main beam according to the embodiment of the present application;

[0028] Figure 3 Schematic diagram of the outer main beam according to the embodiment of the present application.

[0029] Figure 4 Schematic diagram of the fuselage according to the embodiment of the present application.

[0030] Figure 5 Schematic diagram of the frame according to the embodiment of the present application.

[0031] Figure 6Schematic diagram of the wing of the embodiment of the present application.

[0032] Figure 7 Schematic diagram of the power component of the embodiment of the present application.

[0033] Figure 8 Schematic diagram of the fuselage and wing of the embodiment of the present application.

[0034] Figure 9 Top view schematic diagram of the fuselage and wing of the embodiment of the present application.

[0035] Figure 10 Schematic diagram of the solar panel on the wing of the embodiment of the present application.

[0036] Among them, the above-mentioned drawings include the following reference numerals:

[0037] Wing 1, main beam body 11, outer main beam 12, web 121, strut component 122, main beam flange 123, first weight reduction hole 124, connecting component 125, sleeve and connecting lug, web 126, carbon sheet 13, rib 13, leading edge strip 14, trailing edge strip 15, wing skin 16, strengthening beam 17, second weight reduction hole 171, D-shaped box 18,

[0038] Fuselage 2, fuselage skin 21, frame 22, fiber reinforced rib 23, carbon fiber board 24, aluminum alloy column 25,

[0039] Power component 3, tail wing component 4,

[0040] Solar panel 5, aileron 6. Detailed implementation manners

[0041] Next, the technical solutions in the embodiments of the present application will be clearly and completely described in conjunction with the drawings in the embodiments of the present application. Obviously, the described embodiments are only a part of the embodiments of the present application, rather than all the embodiments. Based on the embodiments in the present application, all other embodiments obtained by those of ordinary skill in the art without creative efforts belong to the protection scope of the present application.

[0042] It should be noted that the orientation or positional relationship indicated by terms such as "center", "longitudinal", "lateral", "length", "width", "thickness", "upper", "lower", "front", "rear", "left", "right", "vertical", "horizontal", "top", "bottom", "inner", "outer", "clockwise", "counterclockwise", "axial", "radial", "circumferential", etc. is based on the orientation or positional relationship shown in the drawings. It is only for the convenience of describing the present application and simplifying the description, rather than indicating or implying that the device or element referred to must have a specific orientation, be constructed and operated in a specific orientation. Therefore, it should not be construed as a limitation to the present application. The terms "mounted", "connected", and "coupled" should be understood in a broad sense. For example, it can be a fixed connection, a detachable connection, or an integral connection; it can be a mechanical connection or an electrical connection; it can be directly connected or indirectly connected through an intermediate medium, and it can be the communication inside two elements. The terms "parallel", "perpendicular", and "equal" include the described situations and situations similar to the described situations, and the range of the similar situations is within the acceptable deviation range, where the acceptable deviation range is determined by those of ordinary skill in the art considering the measurements being discussed and the errors associated with the measurements of specific quantities (i.e., the limitations of the measurement system). For example, "parallel" includes absolute parallel and approximate parallel, and the acceptable deviation range of approximate parallel can be, for example, within 5° deviation; "perpendicular" includes absolute perpendicular and approximate perpendicular, and the acceptable deviation range of approximate perpendicular can also be, for example, within 5° deviation. "Equal" includes absolute equality and approximate equality, and the acceptable deviation range of approximate equality can be, for example, that the difference between the two equal ones is less than or equal to 5% of any one of them. For those of ordinary skill in the art, the specific meanings of the above terms in the present application can be understood according to specific circumstances.

[0043] In order to enable those skilled in the art of the present technology to better understand the solution of the present application, the following further elaborates on the present application in conjunction with the Figures 1 to 10 accompanying drawings and specific embodiments.

[0044] The solar aircraft of the embodiment of the present invention includes a wing 1, a fuselage 2, and a reinforcing fiber material. The wing 1 and the fuselage 2 are connected. The wing 1 includes a main beam body 11, an outer main beam 12, and carbon sheets 13. Outer main beams 12 are provided on both sides of the extending direction of the main beam body 11, and carbon sheets 13 are provided on the main beam body 11 and the outer main beam 12 in the height direction of the main beam body 11. Among them, the material of the outer main beam 12 is a composite material. The main beam body 11, the outer main beam 12, and the carbon sheets 13 form a main beam, and a reinforcing fiber material is wound around the outer peripheral surface of the main beam.

[0045] Specifically, the fuselage 2 extends in the front-rear direction, the wing 1 extends in the left-right direction, the height direction of the main beam body 11 is the up-down direction, the extending direction of the main beam body 11 is the left-right direction, and the width direction of the main beam body 11 is the front-rear direction.

[0046] An outer main beam 12 is provided on each of the left and right sides of the main beam body 11, that is, the left end of the main beam body 11 is connected to an outer main beam 12, and the right end of the main beam body 11 is connected to another outer main beam 12. The outer main beam 12 is made of composite material. At the same time, carbon sheets 13 are provided on both the upper and lower sides of the outer main beam 12, and carbon sheets 13 are provided on both the upper and lower sides of the main beam body 11. That is, a carbon sheet 13 is provided at the upper end of the outer main beam 12, a carbon sheet 13 is provided at the lower end of the outer main beam 12, a carbon sheet 13 is connected to the upper end of the main beam body 11, and a carbon sheet 13 is connected to the lower end of the main beam body 11.

[0047] The main beam body 11, the outer main beam 12 and the carbon sheet 13 form a main beam, and the outer peripheral surface of the main beam is wound with a reinforcing fiber material.

[0048] The main beam body 11 is made of composite material, which is convenient for processing and manufacturing with a mold, shortens the manufacturing cycle, uses the composite material to take into account a certain structural strength, and reduces the weight, thereby improving the load capacity of the wing 1 and reducing the weight during flight.

[0049] The present application provides a solar aircraft, which can reduce weight, improve structural strength, and shorten the manufacturing cycle.

[0050] In the solar aircraft of the embodiment of the present invention, the main beam body 11 is made of composite material, which is convenient for manufacturing and processing the main beam body 11 with a mold, shortens the manufacturing cycle, and reduces the manufacturing cost. The main beam body 11, the outer main beam 12 and the carbon sheet 13 form a main beam, and the main beam is wound with a reinforcing fiber material to improve the structural strength, taking into account the strength and avoiding the increase in weight caused by using materials such as metal. At the same time, winding on the main beam body 11 also enhances the torsional resistance.

[0051] In some embodiments, the material of the main beam body 11 is polymethacrylimide foam, the material of the outer main beam 12 is polyaramide fiber paper honeycomb material, and the reinforcing fiber material is Kevlar fiber.

[0052] Specifically, polymethacrylimide foam is a closed-cell foam material that can reduce weight while providing a certain structural strength. Compared with using wood materials, the polymethacrylimide foam has higher structural strength and lower weight. The material of the outer main beam 12 is polyaramid fiber paper honeycomb material, which is made of aramid fiber paper. Compared with wood materials, it improves the compressive strength and shear strength of the outer main beam 12 and reduces the weight. And it has certain heat insulation performance. Since most of the solar panels 5 are arranged above the wing ribs 13 corresponding to the outer main beam 12, the polyaramid fiber paper honeycomb material used for the outer main beam 12 has certain heat insulation performance, which can prevent heat from being transferred to the electronic components below the outer main beam 12 and improve the stability of the aircraft during use.

[0053] At the same time, the polyaramid fiber paper of the honeycomb material can be processed by processes such as cutting, bonding, and machining, which is convenient for integration with other structural components. It improves the convenience of production and reduces the production cost.

[0054] In the solar aircraft of the embodiment of the present invention, the material of the main beam body 11 is PMI composite material, carbon sheets 13 are pasted on the upper and lower sides and Kevlar fibers are wound integrally to enhance the structural strength during flight, support, and load bearing. The outer main beam 12 is made of polyaramid fiber paper honeycomb material to form a honeycomb beam, carbon sheets 13 are pasted on both sides and Kevlar fibers are wound integrally to strengthen, achieving weight reduction while ensuring strength, so that the anti-torsion and anti-bending effects meet the load requirements.

[0055] In some embodiments, the outer main beam 12 includes a web 121, a plurality of strut members 122, and main beam flanges 123. The web 121 is provided with main beam flanges 123 on both sides in the height direction of the main beam body 11. A plurality of strut members 122 are arranged on the web 121 and are connected in sequence in the extending direction of the main beam body 11. The two ends of the strut member 122 in the height direction of the main beam are respectively connected to the corresponding main beam flanges 123.

[0056] A plurality of first weight reduction holes 124 are provided on the web 121. The plurality of first weight reduction holes 124 are arranged at intervals on the web 121 in the extending direction of the main beam body 11, and the cross-sectional area of the first weight reduction holes 124 gradually increases in the direction away from the fuselage 2.

[0057] Specifically, the outer main beam 12 includes a web 121 arranged in the middle and main beam flanges 123 on the upper and lower sides of the web 121. The strut members 122 are arranged on the web 121 to support the web 121 in the up and down directions. At the same time, the upper end of the strut member 122 is connected to the main beam flange 123 at the upper end of the web 121, and the lower end of the strut member 122 is connected to the wing flange at the lower end of the web 121.

[0058] A plurality of first weight reduction holes 124 are provided on the web 121 to reduce the weight of the web 121. At the same time, in the direction away from the fuselage 2, the cross-sectional area of the first weight reduction holes 124 gradually increases, thereby reducing the weight of the main beam while ensuring the strength of the main beam. At the same time, it can also make the overall center of gravity of the aircraft concentrate on the middle part of the wing 1, so as to facilitate the control of the aircraft and improve the stability during flight.

[0059] Further, the strut member 122 includes inclined struts and straight struts, and the inclined struts and straight struts are alternately arranged in the extending direction of the main beam body 11, and the inclined struts on both sides of the straight strut are symmetrically arranged with respect to the straight strut.

[0060] The strut member 122 is composed of inclined struts and straight struts, and these two types of struts are alternately arranged in the extending direction of the main beam body 11. This alternating arrangement optimizes the structural strength and stiffness, and there is no need to increase the thickness of the web 121 to increase the structural strength, thereby reducing the weight. In other words, the inclined struts are symmetrically arranged on both sides of the straight strut. This symmetrical arrangement not only enhances the structural stability but also helps to disperse and resist various forces and torques generated during flight.

[0061] By alternately arranging the inclined struts and straight struts, and symmetrically arranging the inclined struts, the overall structural strength of the outer main beam 12 can be improved. This enables the outer main beam 12 to withstand various loads during flight, including lift, drag, and torque.

[0062] The symmetrically arranged and alternately arranged strut member 122 helps to concentrate the overall center of gravity of the aircraft on the middle part of the wing 1, thereby improving the stability and controllability during flight.

[0063] In some embodiments, the outer main beam 12 further includes a connecting member 125 and a web 121. The connecting member 125 is disposed between the web 121 and the main beam body 11. The connecting member 125 includes a sleeve and a connecting lug. One end of the sleeve extends into the web 121 and is connected to the web 121, and the other end of the sleeve extends into the main beam body 11 and is connected to the main beam body 11. Both the main beam flange 123 and the main beam body 11 are provided with connecting lugs, and the connecting lugs are provided with threaded holes.

[0064] The material of the connecting lug is carbon fiber, and the sleeve is a carbon nanotube.

[0065] Specifically, the sleeve is disposed between the main beam body 11 and the web 121. One end of the sleeve is connected to the main beam body 11, and the other end of the sleeve is connected to the web 121. A connecting lug is provided at the upper end of the web 121, and a connecting lug is provided at the upper end of the main beam body 11, thereby facilitating the fixing by passing a bolt through the two connecting lugs.

[0066] In the solar aircraft according to the embodiments of the present invention, the connecting lugs are respectively arranged on the main beam flange 123 and the main beam body 11 and are connected by bolts. The connecting lugs are made of carbon fiber material, which reduces the overall weight and improves the structural strength. The sleeve is a carbon nanotube. One end of the sleeve extends into the inside of the web 121 and is connected thereto, and the other end extends into the inside of the main beam body 11 and is connected thereto. To ensure the continuity and stability of the connection of the wing 1. The sleeve is made of carbon nanotube material to reduce the weight while improving the connection strength between the components of the wing 1.

[0067] In some embodiments, the wing 1 further includes a plurality of wing ribs 13, a leading edge strip 14, a trailing edge strip 15, and a wing skin 16. The plurality of wing ribs 13 are arranged at intervals along the extending direction of the main beam body 11, and the wing ribs 13 are sleeved on the main beam body 11. A leading edge strip 14 is provided at one end of the wing rib 13 facing the fuselage 2, and a trailing edge strip 15 is provided at the other end of the wing rib 13.

[0068] The wing skin 16 is arranged at the top and bottom of the wing rib 13. The wing skin 16 at the top of the wing rib 13 is connected to the wing skin 16 at the bottom at the leading edge, and the connection part is processed into an arc shape. And the material of the wing skin 16 is balsa wood.

[0069] Specifically, the wing skin 16 is arranged at the lower end or the upper end of the wing rib 13, and the wing skins 16 at the upper end and the lower end are connected at the leading edge strip 14, that is, the connection part is processed into an arc shape, and the wing skin 16 at the connection part is chamfered to reduce the flight resistance and avoid the wing skin 16 at the upper end or the wing skin 16 at the lower end falling off due to air resistance during flight.

[0070] In other words, the wing skin 16 at the connection part is designed into an arc shape, which reduces the flight resistance and improves the flight efficiency. The arc shape can guide the air flow more smoothly, reduce the generation of turbulence and eddy currents, thereby reducing the air resistance, and at the same time avoid stress concentration and improve the service life.

[0071] In some embodiments, the wing 1 further includes a strengthening beam 17 and a D-shaped box 18. The D-shaped box 18 is connected to the strengthening beam 17 and the wing rib 13, and the bottom of the D-shaped box 18 is connected to the bottom of the fuselage 2.

[0072] The rib 13 is provided with a through hole that mates with the reinforcing beam 17. The rib 13 mates with the through hole to sleeved on the reinforcing beam 17. The reinforcing beam 17 is provided with a second weight reduction hole 171, and the second weight reduction holes 171 are provided on both sides of the reinforcing beam 17 that are away from the fuselage 2. The reinforcing beam 17 is arranged at the rear side of the rib 13 to strengthen the overall structural strength of the wing 1. At the same time, the reinforcing beam 17, the main beam body 11, and the outer main beam 12 are arranged at intervals in the front-back direction to improve the overall stiffness and strength of the wing 1. And the spaced arrangement of the main beam and the reinforcing beam 17 is to disperse and resist various forces and torques generated during flight, improving the safety and reliability of the aircraft.

[0073] The D-shaped box 18 connects the reinforcing beam 17, the rib 13, and the bottom of the fuselage 2 together to form an integral structure, improving the stiffness and strength of the wing 1, enabling the wing 1 to better withstand various loads during flight. The bottom of the D-shaped box 18 is connected to the bottom of the fuselage 2. This connection method ensures a firm connection and stable support between the wing 1 and the fuselage 2, optimizing the aerodynamic performance of the aircraft.

[0074] In some embodiments, the wing 1 and the fuselage 2 have a preset angle in the height direction of the main beam. The preset angle is 2°. The ratio of the dimension of the wing 1 in the extending direction of the main beam body 11 to the dimension of the wing 1 in the width direction of the main beam body 11 is 10.92.

[0075] The dimension of the wing 1 in the extending direction of the main beam body 11 is 4320 mm, and the dimension of the fuselage 2 in the width direction of the main beam body 11 is 996 mm.

[0076] Specifically, the wing 1 and the fuselage 2 have a preset angle A in the up-down direction, and A is 2°. This can optimize the angle of attack, enabling a suitable angle of attack to be formed between the wing 1 and the airflow, which helps to increase the lift coefficient of the wing 1 and can more effectively generate lift during flight, ensuring the flight altitude and stability of the aircraft.

[0077] It can also reduce airflow interference. To a certain extent, it can reduce the airflow interference between the wing 1 and the fuselage 2, reduce aerodynamic noise and drag, improve the overall aerodynamic performance of the aircraft, make the airflow flow more smoothly over the wing 1 and the fuselage 2, and improve energy utilization efficiency. The preset angle A can be understood as being in a certain plane in the up-down direction and the front-back direction. That is, in this plane, a straight line parallel to the direction in which the fuselage extends in the front-back direction and another straight line parallel to the direction in which the wing extends in the front-back direction form an angle of 2° in this plane in the up-down direction and the front-back direction.

[0078] The ratio of the dimension of the wing 1 in the extending direction of the main beam body 11 to the dimension of the wing 1 in the width direction of the main beam body 11 is 10.92. The dimension of the wing 1 in the extending direction of the main beam body 11 is 4320 mm, and the upper surface area is 1.711 ㎡. Reducing induced drag, a high aspect ratio wing 1 can effectively reduce induced drag and improve the lift-to-drag ratio of the aircraft. This means that under the condition of consuming the same energy, the aircraft can fly a longer distance, or under the same flight range, consume less energy, which is beneficial to improving the energy utilization efficiency and endurance of the solar aircraft. Improving gliding performance, enhancing the gliding ability of the aircraft. In the case of insufficient solar energy or when energy needs to be saved, such as during night flight or in case of special situations, the aircraft can use good gliding performance to stay in the air, extend the flight time, and wait for the right time to obtain more solar energy or perform other operations.

[0079] The dimension of the fuselage 2 in the width direction of the main beam body 11 is 996 mm. The shorter length of the fuselage 2 can reduce the materials used, lower the overall weight of the aircraft, and improve fuel efficiency or battery endurance.

[0080] Furthermore, the overall top view of the wing 1 presents a rectangular shape, that is, the taper ratio is 1.0. Furthermore, the structure of the wing 1 is simple. The taper ratio of 1.0 means that the root and tip dimensions of the wing 1 are similar. This structure of the wing 1 is relatively simple and has certain advantages in manufacturing and assembly. At the same time, it is also beneficial to ensure the structural strength and stiffness of the wing 1, reduce the structural weight, make the aerodynamic load distribution of the wing 1 along the span relatively uniform, reduce the problem of local aerodynamic stress concentration, improve the overall aerodynamic performance of the wing 1, contribute to the stable flight of the aircraft, and the performance of the wing 1 is more consistent under different flight attitudes and working conditions.

[0081] In some embodiments, the fuselage 2 includes a plurality of frames 22, a plurality of fuselage skins 21, and fiber reinforced ribs 23. The plurality of frames 22 are spaced apart in the width direction of the main beam body 11, and the outer peripheral surfaces of the frames 22 are all connected with the fuselage skins 21. Fiber reinforced ribs 23 are provided inside the frames 22.

[0082] Among them, the shape of the frame 22 is square, and the side length of the frame 22 gradually decreases along the direction away from the fuselage 2.

[0083] Specifically, the frames 22 are arranged at intervals in the front-rear direction. Four fuselage skins 21 are connected to the outer peripheral surfaces of the frames 22. The frames 22 are square. Fiber reinforced ribs 23 are provided inside the frames 22 to improve the structural strength of the frames 22. At the same time, an arc transition is adopted between the skins on the four sides to enhance the load-bearing capacity. The size of the frames 22 gradually decreases from front to back, that is, the side length of the frames 22 gradually decreases.

[0084] The fuselage 2 is built with gradually changing basswood frames 22 and covered with balsa wood panels on them. The maximum height of the frame 22 is 67 mm, and the minimum height is 47 mm. The connections of the panels on each side are transitioned with arcs. The length of the fuselage 2 is 996 mm. While ensuring the structural strength, the weight is reduced, the load capacity is increased. At the same time, the shorter length of the fuselage 2 can reduce the materials used, lower the overall weight of the aircraft, and improve the fuel efficiency or battery endurance.

[0085] In some embodiments, the solar aircraft further includes a carbon fiber board 24, an aluminum alloy column 25, a power component 3, and a tail wing component 4. One end of the frame 22 away from the wing 1 is connected to one end of the carbon fiber board 24. The other end of the carbon fiber board 24 is connected to one end of the aluminum alloy column 25. The end of the aluminum alloy column 25 away from the carbon fiber board 24 is connected to the power component 3. One end of the tail wing component 4 is connected to the wing 1.

[0086] Specifically, the front end of the carbon fiber board 24 is connected to the rear end of the aluminum alloy column 25, and the power component 3 is installed at the front end of the carbon fiber board 24.

[0087] A power support composed of a carbon fiber board 24 and a double-headed aluminum alloy column 25 is installed at the leading edge of the fuselage 2 of the solar aircraft. The support is fixed to the frame 22 by bolts to ensure a rigid connection of the power system.

[0088] The power component 3 includes a power unit, a brushless motor, and a propeller. The power unit consists of a set of helical gear sets with a transmission ratio of 1:10. The small gear is driven by the brushless motor, and the driven gear drives the propeller to pull forward.

[0089] Further, the power unit of the solar aircraft uses a helical gear set with a transmission ratio of 1:10, a module of 0.5, and a pressure angle of 20°. The gearbox housing is made of 3D printed nylon material, and it is matched with a 2308C, 1150KV brushless motor. The no-load speed reaches 2300 r / min. Under the working condition of 16 m / s, the propeller thrust is 6 N, the transmission efficiency is 82%, and the noise is reduced by 15 dB.

[0090] Further, the propeller of the solar aircraft uses carbon fiber blades. The carbon fiber blades are formed by vacuum molding of 3k200g twill carbon cloth. The surface of the mold is coated with a release agent, and the vacuum pressure is -95 kPa. After clamping, it is cured for 24 hours. The weight of a single propeller is 45 g, and the static balance deviation ≤ 0.1 g. The overall weight of the propeller is reduced, and the balance deviation is very small, not exceeding 0.1 g. The blades are very stable during rotation and will not generate unnecessary vibrations, which can improve the flight stability of the aircraft. The propeller of the solar aircraft uses carbon fiber materials and vacuum molding technology to produce light and stable blades, providing strong guarantee for the flight of the aircraft.

[0091] The solar aircraft also includes a tail wing component 4. The tail wing component 4 of the solar aircraft is designed with an inverted V-shaped twin tail. The unilateral width increases from 125 mm to 208 mm from top to bottom. The wing 1 is equipped with a trapezoidal tail rudder, which is connected by a hinge. The tail rudder is controlled by a corresponding servo for pitch and yaw. The tail wing component 4 is fixed to the fuselage 2 through existing carbon tubes, and both ends of the carbon tubes are reinforced with laminates.

[0092] In some embodiments, it also includes solar panels 5 and ailerons 6. The number of solar panels 5 is multiple. The solar panels 5 are arranged on the top of the wing 1. Ailerons 6 are provided on both sides of the wing 1 in the extending direction of the main beam body 11, and the ailerons 6 are arranged on the side of the wing 1 away from the fuselage 2.

[0093] Specifically, ailerons 6 are provided on both sides of the wing 1 in the extending direction of the main beam body 11, and the ailerons 6 are arranged on the side of the wing 1 away from the fuselage 2, that is, the tail wing component 4 is arranged at the rear of the wing rib 13 and behind the strengthening beam 17. The ailerons 6 are rotatable relative to the wing 1 in the up and down direction. At the same time, multiple servos are provided on the wing 1, and the output ends of the servos corresponding to the tail wing are connected to the front end of the tail wing.

[0094] The control system of the solar aircraft includes an electronic speed controller directly connected to a brushless motor, a radio receiver, an open-source flight control board based on ArduPilot, and four digital servos. The four digital servos are respectively connected to the tail wing or the ailerons 6. The servos respectively control the left / right ailerons 6 and the left / right tail rudders. The servo stroke is ±60°, the response time is 0.12 s, the roll rate of the flight control board is set at 150° / s, the pitch rate is 120° / s, and it is suitable for the working condition of a load of 1 - 2 kg.

[0095] The solar panel 5 is a monocrystalline silicon flexible battery panel. 84 monocrystalline silicon flexible battery panels with an efficiency of 22.5% are laid on the surface of the wing 1 of the solar aircraft, covering 1.1315 ㎡ of the upper surface of the wing 1, accounting for 76.75% of the total area of the upper surface of the wing 1. High-temperature resistant tape is used to bond the mask to reduce light loss. The battery panels are divided into two groups of parallel branches, and each group is connected in series with an anti-reverse diode. The peak voltage is 27.5 V, and the average power output is 250 W. It uses the flight airflow for passive heat dissipation, and the battery operating temperature ≤ 45 °C, which is 15 °C lower than the existing internal battery design.

[0096] Furthermore, the control system of the solar aircraft includes an electronic speed controller directly connected to a brushless motor, a radio receiver, an open-source flight control board based on the ArduPilot flight control system, and four 9g digital servos. The servos respectively control the ailerons 6 on the left and right sides and the tail rudders on the left and right sides. The servo stroke is ±60°, the response time is 0.12 s, the roll rate of the flight control board is set at 150° / s, the pitch rate is 120° / s, and it is suitable for the working condition of a load of 1 - 2 kg.

[0097] Further defined, under the condition of solar irradiance of 100,000 LX, the maximum load of the solar-powered aircraft is 2.1 kg, the cruising speed is 12 m / s. Under a 6-level wind, when the aileron 6 deflects 15°, a stable flight path can be maintained, and when the rudder deflects 20°, the yaw can be corrected.

[0098] Further defined, for the wing 1 rib of the solar-powered aircraft, the visual method is adopted to reduce errors; for the fuselage 2 and the tail wing, rib platforms and desktop scanning are used for positioning to improve accuracy. A dispensing machine is used to precisely control the amount of epoxy resin (error ±0.1 g), and the thickness of the glue layer ≤ 0.2 mm. This can improve the assembly accuracy and reduce the performance degradation or safety hazards caused by errors. Rib platforms and desktop scanning for positioning can also improve production efficiency because once the positioning parameters are determined, the assembly work can be quickly completed. The dispensing machine can precisely control the amount of epoxy resin to ensure that the amount and position of each glue application are accurate. By precisely controlling the amount of epoxy resin, it can be ensured that the thickness of the glue layer is uniform and meets the requirements (≤ 0.2 mm), thereby improving the strength and durability of the connection part. The working mode of this solar-powered aircraft: Place the solar-powered aircraft disclosed in this application on a taxiing cart and take off by accelerating the taxiing through the propeller. In the climbing stage, turn on the maximum power climbing mode of the flight control board, that is, set the climbing angle of the aircraft according to the solar altitude angle obtained from the ground test to obtain the climbing mode with the maximum illuminance. When reaching an altitude of 100 m, switch to the autonomous cruise mode. After the cruise mission is completed, enter the landing route, and the flight control board switches to the stability augmentation mode.

[0099] The above has introduced in detail a certain application provided by this application. Specific examples are used in this article to elaborate on the principle and implementation mode of this application. The description of the above embodiments is only used to help understand the method and its core idea of this application. It should be noted that for those of ordinary skill in the art in this technical field, without departing from the principle of this application, several improvements and modifications can still be made to this application, and these improvements and modifications also fall within the protection scope of the claims of this application.

Claims

1. A solar aircraft, characterized in that, Comprising: A wing and a fuselage, the wing and the fuselage being connected. The wing includes a main beam body, an outer main beam, and carbon sheets. Outer main beams are provided on both sides in the extending direction of the main beam body. Carbon sheets are provided on both the main beam body and the outer main beam in the height direction of the main beam body. Among them, the outer main beam is made of composite material; Reinforcing fiber material. The main beam body, the outer main beam, and the carbon sheets form a main beam, and the outer peripheral surface of the main beam is wound with reinforcing fiber material.

2. The solar aircraft according to claim 1, wherein The main beam body of the main beam is made of polymethacrylimide foam, the outer main beam is made of aramid fiber paper honeycomb material, and the reinforcing fiber material is Kevlar fiber.

3. The solar aircraft according to claim 1, characterized in that, The outer main beam includes a web, a plurality of strut members, and main beam flanges. Main beam flanges are provided on both sides of the web in the height direction of the main beam body. A plurality of the strut members are arranged on the web and are sequentially connected in the extending direction of the main beam body. The two ends of the strut members in the height direction of the main beam are respectively connected to the corresponding main beam flanges. A plurality of first weight-reducing holes are provided on the web. The plurality of first weight-reducing holes are arranged at intervals on the web in the extending direction of the main beam body, and the cross-sectional area of the first weight-reducing holes gradually increases in the direction away from the fuselage.

4. The solar aircraft according to claim 3, wherein The outer main beam further includes a connecting member and a web. The connecting member is arranged between the web and the main beam body. The connecting member includes a sleeve and a connecting lug. One end of the sleeve extends into the web and is connected to the web. The other end of the sleeve extends into the main beam body and is connected to the main beam body. The main beam flanges and the main beam body are both provided with the connecting lugs, and the connecting lugs are provided with threaded holes. The connecting lug is made of carbon fiber, and the sleeve is a carbon nanotube.

5. The solar aircraft according to claim 1, characterized in that, The wing further includes a plurality of wing ribs, a leading edge strip, a trailing edge strip, and a wing skin. The wing ribs are arranged at intervals in the extending direction of the main beam body, and the wing ribs are sleeved on the main beam body. A leading edge strip is provided at one end of the wing rib facing the fuselage, and a trailing edge strip is provided at the other end of the wing rib. The wing skin is arranged on the top and bottom of the wing rib. The wing skin at the top of the wing rib is connected to the wing skin at the bottom at the leading edge, and the connection part is processed into an arc. And the wing skin is made of balsa wood.

6. The solar aircraft according to claim 5, characterized in that The wing further includes a strengthening beam and a D-shaped box. The D-shaped box is connected to the strengthening beam and the wing rib, and the bottom of the D-shaped box is connected to the bottom of the fuselage. The wing rib is provided with a through hole that cooperates with the strengthening beam. The wing rib cooperates with the through hole to be sleeved on the strengthening beam. Second weight-reducing holes are provided on the strengthening beam, and second weight-reducing holes are provided on both sides of the strengthening beam away from the fuselage.

7. The solar aircraft according to claim 1, characterized in that The wing and the fuselage have a preset angle in the height direction of the main beam. The preset angle is 2°. The ratio of the size of the wing in the extending direction of the main beam body to the size of the wing in the width direction of the main beam body is 10.

92. The size of the wing in the extending direction of the main beam body is 4320 mm, and the size of the fuselage in the width direction of the main beam body is 996 mm.

8. The solar aircraft according to any one of claims 1-7, characterized in that, The fuselage includes a plurality of frames, a plurality of fuselage skins and fiber reinforced ribs. The plurality of frames are spaced apart in the width direction of the main beam body, and the fuselage skins are provided on the outer peripheral surfaces of the frames. Fiber reinforced ribs are provided inside the frames. Among them, the shape of the frame is square, and the side length of the frame gradually decreases in the direction away from the fuselage.

9. The solar aircraft according to claim 8, characterized in that, It further includes a carbon fiber plate, an aluminum alloy column, a power component and a tail wing component. One end of the frame away from the wing is connected to one end of the carbon fiber plate. The other end of the carbon fiber plate is connected to one end of the aluminum alloy column. One end of the aluminum alloy column away from the carbon fiber plate is connected to the power component. One end of the tail wing component is connected to the wing.

10. The solar aircraft according to claim 1, wherein It further includes solar panels and ailerons. The number of the solar panels is multiple. The solar panels are arranged on the top of the wing. Ailerons are provided on both sides of the wing in the extending direction of the main beam body, and the ailerons are arranged on the side of the wing away from the fuselage.

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