Wing and manufacturing method thereof

Abstract

The invention relates to the technical field of unmanned aerial vehicles, and discloses a wing and a manufacturing method thereof.The wing comprises a wing body, a driving part arranged at the root of the wing body and a wingtip arranged at the wingtip of the wing body, and the wing body comprises an upper skin and a lower skin which are oppositely arranged; the upper skin and the lower skin are integrally formed, the foam core material is filled between the upper skin and the lower skin, the upper skin, the lower skin and the foam core material are integrally formed in a curing mode, the driving part and the wingtip are connected between the upper skin and the lower skin, and the driving part is used for being connected with a steering engine. According to the invention, the weight of the wing is effectively reduced, the voyage and cruising ability are increased, the energy consumption is reduced, the price of the foam core material is low, and the cost is reduced. Meanwhile, the upper skin, the lower skin and the foam core material are integrally cured and formed, so that the processing efficiency is improved, the overall deformation is small, the strength is high, and the strength of the wing is effectively ensured.

Description

Technical Field

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

[0002] The performance of drones is largely determined by the wing structure, so in existing designs, the weight of the wings is reduced as much as possible and their flight resistance is reduced. However, since the wings bear a large load and have complex structural forms and connection relationships, the weight reduction of existing tail structures is very limited. In addition, the wings are easily damaged during use, and the overall replacement cost is high. Summary of the invention

[0003] In order 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 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, and the upper skin, the lower skin and the foam core material are solidified and formed as a whole, 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.

[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 a 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 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 relatively arranged leading edge portions and a rear wing surface portion in the width direction, the rear wing surface portion is located between the two leading edge portions, the leading edge portions have the same thickness, and the rear wing surface portion includes a plurality of rear wing surface units along the length direction of the upper skin or the lower skin, and the thickness of the rear wing surface units decreases successively from the wing root to the wing tip.

[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 respectively 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 comprises the following steps:

[0015] Lay 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 up the carbon fiber cloth for the upper skin according to the design direction;

[0018] Injecting resin into the mold, and after the resin soaks the carbon fiber cloth, heating the mold to cure the resin after heating, and curing the upper skin, lower skin, foam core, drive member and wing tip as a whole into one part 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 the cruising state, the maximum stress, maximum deformation of the wing tip and the maximum torsion angle of the wing tip are obtained, and compared with the maximum stress, maximum deformation and maximum torsion angle of the wing tip required by the design input.

[0023] The wing structure strength analysis was conducted under overload conditions to obtain the maximum stress position of the wing under 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 component of the wing is less than the strength limit value of the material used, then the strength of each component of the wing meets the technical requirements and the static test is carried out;

[0025] The method of static test is:

[0026] Strain gauges are laid on the surface of the driving part and the connection between the skin and the driving part, and 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 drone is a state in which the wings are cruising and flying 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 wing tip is 10 mm and the maximum twist angle is -0.5 degrees.

[0031] Preferably, the loading test is carried out in seven levels, with the loading forces being 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 400-500 mm.

[0033] Compared with the prior art, the wing and the manufacturing method thereof according to the embodiment of the present invention have the following beneficial effects: the wing body adopts an upper skin and a lower skin arranged relatively to each other, and a foam core material is filled between the upper skin and the lower skin, thereby effectively reducing the weight of the wing, increasing the range and endurance, reducing energy consumption, and the foam core material is cheap, thus reducing the cost. At the same time, the upper skin, the lower skin and the foam core material are integrally solidified and formed, which not only improves the processing efficiency, but also has small overall deformation and high strength, thereby effectively ensuring the strength of the wing. BRIEF DESCRIPTION OF THE DRAWINGS

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

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

[0036] Figure 3 It is a schematic structural diagram of the driving member of the present invention.

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

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

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

[0040] Figure 7 It is a wing displacement cloud diagram of the wing of the present invention under an overload state.

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

[0042] Fig. 9 It is a stress cloud diagram of the wing skin misses of the wing of the present invention under an overload state.

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

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

[0045] The specific implementation of the present invention is further described in detail below in conjunction with 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 comprises a wing body 1, a driving member 2 arranged at the root of the wing body 1 and a wing tip 3 arranged at the wing tip of the wing body 1, the wing body 1 comprises an upper skin 11 and a lower skin 12 arranged relatively 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 integrally solidified and formed, the driving member 2 and the wing tip 3 are both connected between the upper skin 11 and the lower skin 12, and the driving member 2 is used to connect to a 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, and a boss 22 is provided at one end of the driving body 21 away from the wing body 1, and a half gear 23 for connecting to the servo is provided at one end of the boss 22 away from the driving body 21, and a wing mounting hole 24 is also provided on the boss 22, and the wing mounting hole 24 is used to connect the rotating shaft, and the half gear 23 is meshed with the driving gear, that is, the wing body 1 is driven to be unfolded or closed by the servo.

[0047] The material of the skin is carbon fiber cloth, and the upper skin 11 and the lower skin 12 are both composed of multiple layers of carbon fiber cloth in a stacked shape. The wing tip 3 is integrally milled from aluminum alloy, and its main function is to maintain the shape. It is fixed between the upper skin 11 and the lower skin 12 by resin bonding. The material of the drive member 2 is aluminum alloy. Aluminum alloy is light in weight. By making the drive member 2 and the wing tip 3 from aluminum alloy, the weight of the wing can be reduced as a whole.

[0048] Based on the wing of the above technical features, the wing body 1 adopts a structure in which an upper skin 11 and a lower skin 12 are arranged relatively, and a foam core material 13 is filled between the upper skin 11 and the lower skin 12, thereby effectively reducing the weight of the wing, increasing the range and endurance, reducing energy consumption, and the foam core material is cheap, which reduces the cost. At the same time, the upper skin 11, the lower skin 12 and the foam core material 13 are integrally solidified and formed, which not only improves the processing efficiency, but also has small overall deformation and high strength, thereby effectively ensuring the strength of the wing.

[0049] In this embodiment, the upper skin 11, the lower skin 12 and the foam core material 13 are cured into one 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, it 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 of the skin needs to be relatively thicker, and the remaining parts can be successively thinner, so as 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 the airflow passes) are smooth to ensure the smoothness of the outer airflow. The inner sides of the upper skin 11 and the lower skin 12 include two relatively arranged leading edge portions 121 and a rear wing face portion in the width direction, and the rear wing face portion is located between the two leading edge portions 121. The rear wing face portion includes a plurality of rear wing surface units 122 along the length direction of the upper skin 11 or the lower skin 12, and the thickness of the rear wing surface units 122 decreases successively from the wing root to the wing tip.

[0053] In the specific setting, from the wing root to the wing tip, the rear wing unit 122 is defined as the first rear wing unit, the second rear wing unit, the third rear wing unit... the nth rear wing unit, and the thickness of the first rear wing unit is H1, the thickness of the second rear wing unit is H2, and from the third rear wing unit to the nth rear wing 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] That is, the number of the rear wing units 122, the maximum thickness of the rear wing units 122, the minimum thickness of the rear wing units 122, and the thickness difference between adjacent rear wing units 122 are interrelated and interact with each other. When making, each value must be fully considered to ensure that all of the above relationships are satisfied, so as to ensure the connection stability between the drive member 2 and the wing body 1 and avoid stress concentration, thereby ensuring the performance of the wing. When setting specifically, the thickness of the first rear wing unit can be determined according to the thickness of the drive member body 21, and then the remaining parameters can be determined based on this.

[0060] In a specific example, the number of the rear airfoil units 122 is five, that is, 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] In order to solve the above technical problems, the present application also provides a wing manufacturing method, which comprises the following steps:

[0063] Laying 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 arranged 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 when the resin soaks 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] According to the distributed aerodynamic loads of the wing when cruising and leveling at a speed of 0.8 Mach, and the distributed aerodynamic loads that meet the requirements of 3G overload and 1.5 times the structural safety factor load of the wing at an angle of attack of 12 degrees, the infinite plate interpolation (IPS) method is used to convert the node loads of the aerodynamic surface 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 load, the corresponding maximum displacement cloud map and maximum Von Mises stress map are obtained. According to the above figure, the corresponding maximum value can be obtained, and the comparison can be made according to the corresponding value. Specifically:

[0070] By simulating and analyzing the wing structure of the UAV in the cruising state, the maximum stress of the wing, the maximum deformation of the wingtip and the maximum torsion angle of the wingtip 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 wing structure of the UAV in the cruising state, the maximum stress, maximum deformation of the wing tip and the maximum torsion angle of the wing tip were obtained, and compared with the maximum stress, maximum deformation and maximum torsion angle of the wing tip required by the design input. In the cruising state, the maximum deformation of the wing tip is 10mm and the maximum torsion angle is -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 one side wing is 787.129N), it is found that the maximum stress of the wing is at the root of the wing, which is 50.3MPa, 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 structure strength analysis was carried out under the load conditions of 3G overload and 1.5 times safety factor (single-side wing lift is 3245.296N). It can be seen that under the action of the limit load, the maximum stress of the wing appears on the step mating surface where the lower skin of the wing root contacts the drive part. 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 the material 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, where the maximum σ 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 component of the wing is less than the strength limit value of the material used, then the strength of each component of the wing meets the technical requirements and the static test is carried out;

[0076] The method of static test is:

[0077] Through the strain electrical measurement method, strain gauges are laid on the surface of the driver body 21 that is not covered with the skin and 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. Among them, the position where the loading connector is installed on the wing body is the installation point, and the spacing between the installation point and the center of the wing installation hole is 400 to 500mm, such as 410mm, 420mm, 430mm, 440mm, 450mm, 460mm, 470mm, 480mm, 490mm, etc., preferably 441mm. And the installation point is at a position where the center of the wing installation hole is offset 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 wing manufactured by the above method was subjected to static tests to examine the static strength of the connection between the metal drive parts and the composite materials in the wing structure, indicating that the strength calculation analysis is not much different from the actual situation and meets the strength requirements. At the same time, the weight of the wing is reduced, thereby improving the flight performance and reducing the cost of the wing.

[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: It includes a wing body, a driving component 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 driving component and the wingtip are both connected between the upper skin and the lower skin, and the driving component is used to connect to a servo.

2. The wing according to claim 1, characterized in that: 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 a 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.

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

4. The wing according to claim 2, characterized in that: 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 surface portion that are relatively arranged in the width direction. The rear wing surface portion is located between the two leading edge portions, and the leading edge portions have the same thickness. The rear wing surface portion includes a plurality of rear wing surface units along the length direction of the upper skin or the lower skin, and the thickness of the rear wing surface units decreases successively from the wing root to the wing tip.

5. The wing according to claim 4, characterized in that: In the direction from the wing root to the wing tip, the rear wing units are respectively the first rear wing unit, the second rear wing unit, the third rear wing unit...the nth rear wing unit, and the thickness of the first rear wing unit is H1, the thickness of the second rear wing unit is H2, and from the third rear wing unit to the nth rear wing 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 to 3 mm; The thickness of the leading edge is h, then h = H n .

6. A method for manufacturing a wing, characterized in that: The following steps are involved: Lay 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 up the carbon fiber cloth for the upper skin according to the design direction; Injecting resin into the mold, and after the resin soaks the carbon fiber cloth, heating the mold to cure the resin after heating, and curing the upper skin, lower skin, foam core, drive member and wing tip as a whole into one part to form a wing; Verify the strength of the wing.

7. The method for manufacturing a wing according to claim 6, characterized in that: Verification of the strength of the wing structure includes dynamic strength tests and static tests, including: Dynamic strength test method is: By simulating and analyzing the wing structure of the UAV in the cruising state, the maximum stress, maximum deformation of the wing tip and the maximum torsion angle of the wing tip are obtained, and compared with the maximum stress, maximum deformation and maximum torsion angle of the wing tip required by the design input. The wing structure strength analysis was conducted under overload conditions to obtain the maximum stress position of the wing under 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 component of the wing is less than the strength limit value of the material used, then the strength of each component of the wing meets the technical requirements and the static test is carried out; The method of static test is: Strain gauges are laid on the surface of the driving part and the connection between the skin and the driving part, and 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.

8. The method for manufacturing a wing according to claim 6, characterized in that: The cruising state of the drone is the state where 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.

9. The method for manufacturing a wing according to claim 6, characterized in that: In cruising state, the maximum deformation of the wingtip is 10mm and the maximum torsion angle is -0.5 degrees.

10. The method for manufacturing a wing according to claim 6, characterized in that: The loading test is carried out in seven levels, with loading forces of 0N, 550N, 1100N, 1650N, 2200N, 2750N, and 3300N; 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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