Wing and method for manufacturing the same

Abstract

The present invention relates to the technical field of unmanned aerial vehicles, and discloses a wing and a method for manufacturing the same, wherein the wing comprises a wing body, a driving member arranged at the root of the wing body, and a wingtip arranged at the wingtip of the wing body, the wing body comprises an upper skin and a lower skin arranged relatively to each other, and a foam core material filled between the upper skin and the lower skin, and the upper skin, the lower skin, and the foam core material are solidified and formed as one piece, the driving member and the wingtip are both connected between the upper skin and the lower skin, and the driving member is used to connect to a servo. The present invention effectively reduces the weight of the wing, increases the range and endurance, reduces energy consumption, and the foam core material is cheap, which reduces costs. At the same time, the upper skin, the lower skin, and the foam core material are solidified and formed as one piece, which not only improves processing efficiency, but also has small overall deformation and high strength, thereby effectively ensuring the strength of the wing.

Description

Technical Field

[0001] The present invention relates to the technical field of unmanned aerial vehicles (UAVs), and in particular to a wing and a manufacturing method thereof. Background Art

[0002] The performance of drones is largely determined by their wing structure. Therefore, existing designs strive to reduce wing weight and flight resistance. However, due to the heavy loads borne by wings and their complex structure and connections, existing tail wing structures are severely limited in their ability to reduce weight. Furthermore, wings are susceptible to damage during use, resulting in high replacement costs. Summary of the Invention

[0003] To solve the above technical problems, the present invention provides a wing and a manufacturing method thereof, which can reduce the weight of the wing while ensuring the structural strength of the wing, thereby improving the flight performance and reducing the cost of the wing.

[0004] The technical solution adopted by the present invention to solve its technical problem is:

[0005] A wing includes a wing body, a drive member arranged at the root of the wing body, and a wingtip arranged at the wingtip of the wing body, the wing body includes an upper skin and a lower skin arranged opposite to each other, and a foam core material filled between the upper skin and the lower skin, and the upper skin, the lower skin and the foam core material are solidified and formed as a whole, the drive member and the wingtip are both connected between the upper skin and the lower skin, and the drive member is used to connect to a servo.

[0006] Preferably, the driving member includes a driving body arranged between the upper skin and the lower skin, a boss is provided at one end of the driving member body away from the wing body, a half gear for connecting to the servo is provided on the side of the boss away from the driving body, and a wing mounting hole is also provided on the boss.

[0007] Preferably, the upper skin, the lower skin and the foam core material are solidified into one body by resin, and the other end of the driving body is provided with a groove for filling with resin.

[0008] Preferably, the upper skin and the lower skin have the same structure, their outer surfaces are smooth, and their inner sides include two leading edge portions and a rear wing portion that are relatively arranged in the width direction. The rear wing portion is located between the two leading edge portions, and the leading edge portions have the same thickness. The rear wing portion includes multiple rear wing units along the length direction of the upper skin or the lower skin, and the thickness of the rear wing units decreases successively from the wing root to the wingtip.

[0009] Preferably, in the direction from the wing root to the wing tip, the rear aerofoil units are respectively the first rear aerofoil unit, the second rear aerofoil unit, the third rear aerofoil unit... the nth rear aerofoil unit, and the thickness of the first rear aerofoil unit is H1, the thickness of the second rear aerofoil unit is H2, and from the third rear aerofoil unit to the nth rear aerofoil unit, the adjacent ones are H1, H2, H3...H n The thickness of the two adjacent rear airfoil units is H m 、H m-1 , and n≥m≥3, then they satisfy:

[0010] H1-H2=(2~2.5);

[0011] H m -H m-1 =(0.7~1);

[0012] The value range of H1 is 7~8mm, H n The value range is 2 to 3 mm;

[0013] The thickness of the leading edge is h, then h=H n .

[0014] A method for manufacturing a wing, comprising the following steps:

[0015] Lay out the carbon fiber cloth of the lower skin in the mold according to the direction;

[0016] Filling the lower skin with a foam core, and arranging a driving member and a wing tip at both ends of the foam core respectively;

[0017] Lay out the carbon fiber cloth for the upper skin according to the design direction;

[0018] Injecting resin into the mold, and after the resin soaks into the carbon fiber cloth, heating the mold to cure the resin after heating, and solidifying the upper skin, lower skin, foam core, drive member and wing tip into one part as a whole to form a wing;

[0019] Verify the strength of the wing.

[0020] Preferably, the verification of the strength of the wing structure includes a dynamic strength test and a static test, wherein:

[0021] Dynamic strength test method is:

[0022] By simulating and analyzing the wing structure of the UAV in cruising state, the maximum stress, maximum deformation of the wingtip, and maximum torsion angle of the wing at this time were obtained, and compared with the maximum stress, maximum deformation, and maximum torsion angle of the wing required by the design input.

[0023] The wing structural strength was analyzed under overload conditions to obtain the maximum stress position of the wing under the action of the wing limit load, and to compare whether the maximum stress at this position meets the stress strength of the material itself;

[0024] When the wing is in the above-mentioned cruising state and overload state, the Von Misses stress of each wing component is less than the strength limit value of the material used, then the strength of each wing component meets the technical requirements and the static test is carried out;

[0025] The method of static test is:

[0026] Strain gauges are applied on the surface of the drive component and at the connection between the skin and the drive component. Loading connectors are installed on the wing body to apply concentrated loads to simulate the working force of the wing.

[0027] Inspect the wings during and after the test. If no peeling or tearing of the skin is found and no deformation of the drive parts is found, the wings are qualified.

[0028] Preferably, the cruising state of the UAV is a state in which the wings are cruising and level at a speed of Mach 0.8;

[0029] The overload state is a load that meets the wing 3G overload and 1.5 times the structural safety factor at an angle of attack of 12 degrees.

[0030] Preferably, in the cruising state, the maximum deformation of the wingtip is 10 mm and the maximum torsion angle is -0.5 degrees.

[0031] Preferably, the loading test is carried out in seven levels, with loading forces of 0N, 550N, 1100N, 1650N, 2200N, 2750N, and 3300N respectively;

[0032] The position where the loading connector is installed on the wing body is the installation point, and the distance between the installation point and the center of the wing installation hole is 400 to 500 mm.

[0033] Compared to existing technologies, the wing and its manufacturing method according to embodiments of the present invention offer the following advantages: the wing body utilizes an upper and lower skin arranged oppositely, with a foam core material filling the space between the upper and lower skins. This effectively reduces the wing's weight, increases its range and endurance, and reduces energy consumption. Furthermore, the foam core material is inexpensive, reducing costs. Furthermore, the upper and lower skins, along with the foam core material, are integrally cured and molded, improving processing efficiency while minimizing overall deformation and increasing strength, thereby effectively ensuring the wing's strength. BRIEF DESCRIPTION OF THE DRAWINGS

[0034] Figure 1 It is a structural schematic diagram of the wing of the present invention.

[0035] Figure 2 for Figure 1 Schematic cross-section diagram.

[0036] Figure 3 Schematic diagram of the structure of the driving member of the present invention.

[0037] Figure 4 This is a top view of the lower skin of this application.

[0038] Figure 5 This is a wing displacement cloud diagram of the wing of the present invention in the cruising state.

[0039] Figure 6 This is a wing misses stress cloud diagram of the wing of the present invention in the cruising state.

[0040] Figure 7 This is a wing displacement cloud diagram of the wing of the present invention under overload condition.

[0041] Figure 8 This is a stress cloud diagram of the wing misses of the present invention under an overload state.

[0042] Figure 9 This is a stress cloud diagram of the wing skin misses of the wing of the present invention under overload condition.

[0043] Figure 10 This is a stress cloud diagram of the wing embedded parts misses under the overload state of the wing of the present invention.

[0044] Among them: 1-wing body, 11-upper skin, 12-lower skin, 121-leading edge, 122-rear wing unit, 13-foam core material, 2-drive member, 21-drive member body, 22-boss, 23-half gear, 24-wing mounting hole, 25-groove, 3-wingtip. DETAILED DESCRIPTION

[0045] The following embodiments of the present invention are described in further detail with reference to the accompanying drawings and examples. The following examples are used to illustrate the present invention but are not intended to limit the scope of the present invention.

[0046] like Figure 1-3As shown, a wing of a preferred embodiment of the present invention includes a wing body 1, a driving member 2 arranged at the root of the wing body 1 and a wingtip 3 arranged at the wingtip of the wing body 1, the wing body 1 includes an upper skin 11 and a lower skin 12 arranged opposite to each other, and a foam core material 13 filled between the upper skin 11 and the lower skin 12, and the upper skin 11, the lower skin 12 and the foam core material 13 are solidified and formed as a whole, the driving member 2 and the wingtip 3 are both connected between the upper skin 11 and the lower skin 12, and the driving member 2 is used to connect to the servo. Specifically, the servo on the fuselage is connected to a driving gear, and the driving member 2 includes a driving body 21 arranged between the upper skin 11 and the lower skin 12. The driving body 21 is provided with a boss 22 at one end away from the wing body 1, and a half gear 23 for connecting to the servo is provided at the end of the boss 22 away from the driving body 21. The boss 22 is also provided with a wing mounting hole 24, and the wing mounting hole 24 is used to connect the rotating shaft. The half gear 23 is engaged with the driving gear, that is, the wing body 1 is driven to expand or close by the servo.

[0047] The skin is made of carbon fiber cloth, with both the upper skin 11 and the lower skin 12 formed from multiple layers of laminated carbon fiber cloth. The wingtip 3 is integrally milled from aluminum alloy, primarily maintaining its shape. It is secured between the upper and lower skins 11, 12 via resin bonding. The actuator 2 is also made of aluminum alloy. Aluminum alloy is lightweight, and by making the actuator 2 and wingtip 3 from aluminum alloy, the overall weight of the wing can be reduced.

[0048] Based on the above-mentioned technical features, the wing body 1 employs an upper skin 11 and a lower skin 12 arranged opposite each other, with a foam core 13 filled between the upper and lower skins 11, 12. This effectively reduces the wing's weight, increases its range and endurance, and reduces energy consumption. Furthermore, the foam core is inexpensive, reducing costs. Furthermore, the upper and lower skins 11, 12, and foam core 13 are integrally solidified, improving processing efficiency while minimizing overall deformation and increasing strength, thereby effectively ensuring the wing's strength.

[0049] In this embodiment, the upper skin 11, the lower skin 12 and the foam core material 13 are solidified into one body by resin, and a groove 25 for filling resin is provided at the end of the driving member body 21 away from the boss 22. After installation, the outer surface of the driving member body 21 is fixedly connected to the upper skin 11 and the lower skin 12 by resin, and at the same time forms a whole with the wing body 1 through the groove 25, thereby ensuring the connection strength between the driving member 2 and the wing body 1.

[0050] In this embodiment, since the driver body 21 is disposed between the upper skin 11 and the lower skin 12, that is, the skin completely covers the driver body 21, the wing root portion of the skin needs to be relatively thicker, while the remaining portions can be successively thinner to reduce stress concentration. Specifically:

[0051] In this application, the attached Figure 1 The L direction is the length direction of the skin, and the W direction is the width direction of the skin.

[0052] like Figure 2 、 4 As shown, the upper skin 11 and the lower skin 12 have the same structure, and their outer side surfaces (i.e., the surface through which air flows) are smooth to ensure smooth airflow outside. The inner sides of the upper skin 11 and the lower skin 12 include two leading edge portions 121 and a rear wing face portion oppositely disposed in the width direction. The rear wing face portion is located between the two leading edge portions 121. The rear wing face portion includes multiple rear wing units 122 along the length direction of the upper skin 11 or the lower skin 12, and the thickness of the rear wing units 122 decreases from the wing root to the wingtip.

[0053] In the specific setting, from the wing root to the wing tip, the rear aerofoil unit 122 is defined as the first rear aerofoil unit, the second rear aerofoil unit, the third rear aerofoil unit... the nth rear aerofoil unit, and the thickness of the first rear aerofoil unit is H1, the thickness of the second rear aerofoil unit is H2, and from the third rear aerofoil unit to the nth rear aerofoil unit, the adjacent ones are H1, H2, H3...H n The thickness of the two adjacent rear airfoil units is H m 、H m-1 , and n≥m≥3, then they satisfy:

[0054] H1-H2=(2-2.5), such as 2.1mm, 2.2mm, 2.3mm, 2.4mm, etc., preferably 2.2mm;

[0055] H m -H m-1 =(0.7~1), such as 0.75mm, 0.8mm, 0.81mm, 0.82mm, 0.84mm,

[0056] 0.86mm, 0.88mm, 0.9mm, 0.91mm, 0.93mm, 0.95mm, 0.98mm, 0.99mm, etc., preferably 0.88mm;

[0057] The value range of H1 is 7-8 mm, such as 7.1 mm, 7.2 mm, 7.3 mm, 7.4 mm, 7.5 mm, 7.6 mm, 7.7 mm, 7.8 mm, 7.9 mm, etc., preferably 7.51 mm. n The value range is 2 to 3 mm, such as 2.1 mm, 2.2 mm, 2.3 mm, 2.4 mm, 2.5 mm, 2.6 mm, 2.7 mm, 2.8 mm, 2.9 mm, etc., preferably 2.67 mm;

[0058] The thickness of the front edge portion 121 is h, then h=H n , thereby maintaining the integrity of the upper skin 11 and the lower skin 12.

[0059] Specifically, the number of rear airfoil units 122, the maximum thickness of the rear airfoil units 122, the minimum thickness of the rear airfoil units 122, and the thickness difference between adjacent rear airfoil units 122 are interrelated and interact with each other. During fabrication, these values ​​must be fully considered to ensure that all of these relationships are met. This ensures both stable connection between the driver 2 and the wing body 1 and avoids stress concentration, thereby maintaining wing performance. During specific configuration, the thickness of the first rear airfoil unit can be determined based on the thickness of the driver body 21, and the remaining parameters can then be determined accordingly.

[0060] In a specific example, the number of the rear airfoil units 122 is five, i.e., n=5. The thickness of the first rear airfoil unit 122 is 7.51 mm, the thickness of the fifth rear airfoil unit is 2.67 mm, H1-H2=2.2 mm, and the thickness difference between the three rear airfoil units is 0.88 mm. The thickness of each of the five rear airfoil units can be obtained as shown in the following table:

[0061]

[0062] To solve the above technical problems, the present application also provides a wing manufacturing method, which includes the following steps:

[0063] Lay down the carbon fiber cloth of the lower skin 12 in the mold according to the direction;

[0064] A foam core 13 is filled in the lower skin 12, and a driving member 2 and a wing tip 3 are respectively provided at both ends of the foam core 13;

[0065] Lay up the carbon fiber cloth of the upper skin 11 according to the design direction;

[0066] Resin is injected into the mold, and after the resin soaks into the carbon fiber cloth, the mold is heated to cure the resin after heating, and the upper skin 11, the lower skin 12, the foam core 13, the drive member 2 and the wing tip 3 are integrally cured into one part to form a wing;

[0067] Verify the strength of the wing structure, including dynamic strength tests and static tests.

[0068] Dynamic strength test method is:

[0069] Based on the distributed aerodynamic loads of the wing during cruise level flight at a speed of Mach 0.8, and the distributed aerodynamic loads that meet the wing 3G overload and 1.5 times the structural safety factor load requirements at a 12-degree angle of attack, the infinite plate interpolation (IPS) method is used to interpolate the aerodynamic surface node loads to the structural nodes through the compiled aerodynamic node load and structural node load interpolation calculation program. According to the two load conditions of the UAV wing in the cruise state and overload state, under the action of its corresponding aerodynamic loads, the corresponding maximum displacement cloud map and maximum Von Mises stress map are obtained. The corresponding maximum values ​​can be obtained based on the above figures, and the corresponding values ​​can be compared. Specifically:

[0070] By simulating and analyzing the wing structure of the UAV in the cruising state, the maximum stress, maximum deformation of the wingtip and the wingtip torsion angle at this time are obtained, and compared with the maximum stress, maximum deformation and maximum torsion angle of the wing required by the design input.

[0071] By simulating and analyzing the drone's wing structure during cruising, the maximum stress, maximum wingtip deformation, and wingtip torsion angle were determined. These stresses, deformations, and torsion angles were compared with the design input requirements. In cruising mode, the maximum wingtip deformation was 10 mm, and the maximum torsion angle was -0.5 degrees.

[0072] A specific embodiment, such as Figure 5 、 6 As shown in the figure, by simulating and analyzing the wing structure of the UAV in the cruising state (the lift of the wing on one side is 787.129N), it is found that the maximum stress of the wing is 50.3MPa at the wing root, the maximum deformation of the wingtip is 8.13mm, and the wingtip torsion angle is -0.26 degrees. In the level flight state, the maximum deformation and torsion angle of the wingtip are both less than the maximum deformation of 10mm and the maximum torsion angle of -0.5 degrees required by the design input, which meets the stiffness requirements of the wing.

[0073] The wing structural strength analysis was carried out on the wing under overload conditions to obtain the maximum stress position of the wing under the action of the wing limit load, and to compare whether the maximum stress at this position meets the stress strength of the material itself.

[0074] A specific embodiment, such as Figure 7-10As shown in the figure, the wing structural strength analysis was conducted under the load conditions of 3G overload and 1.5 times safety factor (the lift of the wing on one side is 3245.296N). It can be seen that the maximum stress of the wing under the action of the limit load appears on the step mating surface where the lower skin of the wing root contacts the drive component. Figure 7 It can be seen that the maximum stress here occurs on the composite skin, and the maximum σ of the skin here is mises =207MPa stress is much smaller than the minimum σ of carbon fiber woven prepreg b =500MPa, meeting the strength requirements. Figure 8 It can be seen that the maximum stress of the wing drive part 2 occurs at the groove 25 where it is glued to the composite material. mises =150MPa stress is much smaller than the σ of aluminum alloy material b =440MPa, meeting the strength requirements.

[0075] When the wing is in the above-mentioned cruising state and overload state, the Von Misses stress of each wing component is less than the strength limit value of the material used, then the strength of each wing component meets the technical requirements and the static test is carried out;

[0076] The method of static test is:

[0077] By means of strain electrical measurement, strain gauges are applied on the surface of the driver body 21 not covered with the skin and at the connection between the skin and the driver body 21. The strain gauges are 90° strain gauges with a total of 28 strain measurement points. Loading connectors (airfoil pads, hinges) are installed on the wing body 1, and concentrated loads are applied to simulate the working force of the wing. The position where the loading connector is installed on the wing body is the installation point, and the distance between the installation point and the center of the wing mounting hole is 400 to 500 mm, such as 410 mm, 420 mm, 430 mm, 440 mm, 450 mm, 460 mm, 470 mm, 480 mm, 490 mm, etc., preferably 441 mm. The installation point is located at a position offset from the center of the wing mounting hole along the wing span. The loading test is carried out in seven levels, with loading forces of 0N, 550N, 1100N, 1650N, 2200N, 2750N and 3300N respectively.

[0078] Inspect the wings during and after the test. If no peeling or tearing of the skin is found and no deformation of the drive parts is found, the wings are qualified.

[0079] The wings manufactured using the above method were subjected to static tests to examine the static strength of the connection between the metal drive components and the composite materials in the wing structure. The results showed that the strength calculation and analysis were not much different from the actual results, and met the strength requirements. At the same time, the weight of the wing was reduced, thereby improving flight performance and reducing wing costs.

[0080] The above is only a preferred embodiment of the present invention. It should be pointed out that for ordinary technicians in this technical field, several improvements and substitutions can be made without departing from the technical principles of the present invention. These improvements and substitutions should also be regarded as the scope of protection of the present invention.

Claims

1. A wing, characterized in that: The invention comprises a wing body, a driving member arranged at the root of the wing body, and a wingtip arranged at the wingtip of the wing body. The wing body comprises an upper skin and a lower skin arranged opposite to each other, and a foam core material filled between the upper skin and the lower skin. The upper skin, the lower skin, and the foam core material are integrally solidified. The driving member and the wingtip are both connected between the upper skin and the lower skin. The driving member is used to connect to a steering gear. The upper skin and the lower skin have the same structure, and their outer surfaces are smooth. Their inner sides include two leading edge portions and a rear wing portion that are oppositely arranged in the width direction. The rear wing portion is located between the two leading edge portions. The leading edge portions have the same thickness. The rear wing portion includes a plurality of rear wing units along the length direction of the upper skin or the lower skin, and the thickness of the rear wing units decreases sequentially from the wing root to the wing tip. In the direction from the wing root to the wing tip, the rear aerofoil units are respectively the first rear aerofoil unit, the second rear aerofoil unit, the third rear aerofoil unit... the nth rear aerofoil unit, and the thickness of the first rear aerofoil unit is H1, the thickness of the second rear aerofoil unit is H2, and from the third rear aerofoil unit to the nth rear aerofoil unit, the adjacent ones are H1, H2, H3...H n The thickness of the two adjacent rear airfoil units is H m 、H m-1 , and n≥m≥3, then they satisfy: H1-H2=(2~2.5); H m -H m-1 =(0.7~1); The value range of H1 is 7~8mm, H n The value range is 2~3mm; The thickness of the leading edge is h, then h= H n .

2. The wing according to claim 1, wherein: The driving component includes a driving body arranged between the upper skin and the lower skin. A boss is provided at one end of the driving body away from the wing body. A half gear for connecting to a servo is provided on the side of the boss away from the driving body. A wing mounting hole is also provided on the boss.

3. The wing according to claim 2, wherein: The upper skin, the lower skin and the foam core material are solidified into one body through resin, and the other end of the driving body is provided with a groove for filling with resin.

4. A method for manufacturing a wing, characterized in that: The wing is the wing according to any one of claims 1 to 3, and its manufacturing method comprises the following steps: Lay out the carbon fiber cloth of the lower skin in the mold according to the direction; Filling the lower skin with a foam core, and arranging a driving member and a wing tip at both ends of the foam core respectively; Lay out the carbon fiber cloth for the upper skin according to the design direction; Injecting resin into the mold, and after the resin soaks into the carbon fiber cloth, heating the mold to cure the resin after heating, and solidifying the upper skin, lower skin, foam core, drive member and wing tip into one part as a whole to form a wing; Verify the strength of the wing.

5. The wing manufacturing method according to claim 4, characterized in that: Verification of the wing structure strength includes dynamic strength tests and static tests, including: Dynamic strength test method is: By simulating and analyzing the wing structure of the UAV in cruising state, the maximum stress, maximum deformation of the wingtip, and maximum torsion angle of the wing at this time were obtained, and compared with the maximum stress, maximum deformation, and maximum torsion angle of the wing required by the design input. The wing structural strength was analyzed under overload conditions to obtain the maximum stress position of the wing under the action of the wing limit load, and to compare whether the maximum stress at this position meets the stress strength of the material itself; When the wing is in the above-mentioned cruising state and overload state, the Von Misses stress of each wing component is less than the strength limit value of the material used, then the strength of each wing component meets the technical requirements and the static test is carried out; The method of static test is: Strain gauges are applied on the surface of the drive component and at the connection between the skin and the drive component. Loading connectors are installed on the wing body to apply concentrated loads to simulate the working force of the wing. Inspect the wings during and after the test. If no peeling or tearing of the skin is found and no deformation of the drive parts is found, the wings are qualified.

6. The wing manufacturing method according to claim 4, wherein: The UAV cruise state is the state in which the wings are cruising and flying at a speed of 0.8 Mach; The overload state is a load that meets the wing 3G overload and 1.5 times the structural safety factor at an angle of attack of 12 degrees.

7. The wing manufacturing method according to claim 4, wherein: In cruising state, the maximum deformation of the wingtip is 10mm and the maximum torsion angle is -0.5 degrees.

8. The wing manufacturing method according to claim 5, wherein: The loading test was conducted at seven levels, with loading forces of 0 N, 550 N, 1100 N, 1650 N, 2200 N, 2750 N, and 3300 N; The position where the loading connector is installed on the wing body is the installation point, and the distance between the installation point and the center of the wing installation hole is 400-500 mm.

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