Long-endurance unmanned aerial vehicle and aerodynamic optimization method thereof

Abstract

The application provides a long-endurance unmanned aerial vehicle and a method for aerodynamic optimization thereof, comprising a fuselage, a left wing middle section, a right wing middle section, a left wing tail section, a right wing tail section and a tail wing; the fuselage is connected with the left wing middle section and the right wing middle section through outwardly extending carbon tubes and buckles on both sides of the fuselage; the fuselage is connected with the tail wing through buckles on both sides of the tail of the fuselage and positioning pins on the end face of the tail of the fuselage, the positioning pins are symmetrically distributed along the connector, and the installation of the tail wing is completed by locking the buckles after the tail wing is installed in place; the left wing middle section is connected with the left wing tail section through a carbon tube and a buckle; the right wing middle section is connected with the right wing tail section through a carbon tube and a buckle; the belly of the fuselage is provided with a containing cavity for installing a load cabin, and the containing cavity is fixed with the load cabin through a quick release structure; and the application of the technical scheme can improve the endurance time of the unmanned aerial vehicle.

Description

Technical Field

[0001] This invention relates to the field of unmanned aerial vehicle (UAV) technology, and in particular to a long-endurance UAV and its aerodynamic optimization method. Background Technology

[0002] Unmanned aerial vehicles (UAVs) are unmanned unmanned aircraft controlled by radio remote control equipment and their own program control devices. With the rapid development of electronic technology and materials, UAVs are widely used in communication relay, aerial photography, resource exploration, military and other fields.

[0003] Drones are mainly composed of structural components such as the fuselage, wings, and tail. The fuselage structure is the torso and load-bearing foundation of the drone. It not only fixes and supports other components, connecting the entire drone into a whole, but also bears the loads transmitted from each connecting component, as well as the loads of the equipment and mission payloads inside the fuselage, and its own weight and inertia. Therefore, drones are usually large and heavy, resulting in short flight times. Furthermore, the mission payload is usually fixed to the drone, leading to limited mission scope. Different drones must be used for different tasks, resulting in high costs. Currently, the actual flight time of existing vertical takeoff and landing (VTOL) drones is only 60-90 minutes, which is insufficient for long-distance power line inspection operations. Summary of the Invention

[0004] In view of this, the purpose of this invention is to provide a long-endurance unmanned aerial vehicle (UAV) and its aerodynamic optimization method to improve the endurance of the UAV.

[0005] To achieve the above objectives, the present invention adopts the following technical solution: a long-endurance unmanned aerial vehicle, comprising: a fuselage, a middle section of a left wing, a middle section of a right wing, a tail section of a left wing, a tail section of a right wing, and a tail fin;

[0006] The fuselage is connected to the left wing midsection and the right wing midsection respectively via an extended carbon tube assembly and a first latch;

[0007] The fuselage is connected to the tail fin via a second latch and a locating pin;

[0008] The middle section of the left wing is connected to the tail section of the left wing via a first carbon fiber tube and a third locking device; the middle section of the right wing is connected to the tail section of the right wing via a second carbon fiber tube and a fourth locking device.

[0009] The fuselage has a cavity in its underside for mounting a payload bay.

[0010] In a preferred embodiment, the receiving cavity is provided with U-shaped grooves at both ends along the length of the fuselage, and the top of the load compartment is provided with U-shaped rails that mate with the U-shaped grooves at both ends; the length direction of the U-shaped grooves is parallel to the length direction of the U-shaped rails, and the U-shaped grooves and U-shaped rails are mutually restrained along the fuselage direction so that the load compartment is detachably and fixedly connected to the fuselage; the openings between the mutually adapted U-shaped grooves and U-shaped rails are arranged opposite to each other.

[0011] In a preferred embodiment, the extended carbon fiber assembly includes first extended carbon fiber tubes symmetrically extended outward on both sides of the fuselage, a first internal carbon fiber tube disposed inside the fuselage, second extended carbon fiber tubes extending outward at the mid-sections of the left and right wings facing the fuselage, and second internal carbon fiber tubes disposed at the mid-sections of the left and right wings facing the fuselage; the first extended carbon fiber tubes are inserted into the second internal carbon fiber tubes, and the second extended carbon fiber tubes are inserted into the first internal carbon fiber tubes.

[0012] In a preferred embodiment, the first latch, the second latch, the third latch, and the fourth latch each include: a latch connector and a latch connecting fitting that cooperates with the latch connector. The latch connector is disposed on the middle section of the left wing, the middle section of the right wing, the tail section of the left wing, the tail section of the right wing, and the tail fin. The latch connecting fitting is disposed on the fuselage, the middle section of the left wing, and the middle section of the right wing.

[0013] In a preferred embodiment, the locking connectors provided on the middle section of the left wing and the middle section of the right wing are both located at one end facing the fuselage, and two locking connectors that cooperate with the locking connectors on the middle sections of the right and left wings are located on both sides of the fuselage.

[0014] In a preferred embodiment, the device includes a limiting ring, a fixing arm, and a torsion spring; the fixing arm is connected to the limiting ring via the torsion spring; a pivot is provided in the middle of the torsion spring, and the pivot is fixedly connected to the middle section of the left wing, the middle section of the right wing, the tail section of the left wing, the tail section of the right wing, and the tail fin; the locking connection is a limiting hook; the limiting hook engages with the limiting ring for limiting.

[0015] In a preferred embodiment, a battery compartment is provided at the front of the fuselage, and an avionics compartment is provided at the rear of the fuselage. The battery compartment and the avionics compartment are used to install batteries and avionics systems, respectively. Antenna mounting covers for installing antennas are provided at the front and rear of the fuselage, respectively. A front landing gear and a rear landing gear are provided at the front and rear of the lower part of the fuselage, and an antenna is embedded inside the rear landing gear.

[0016] This invention also provides an aerodynamic optimization method for a long-endurance unmanned aerial vehicle (UAV), which, using the aforementioned long-endurance UAV, includes the following steps:

[0017] Step S1: Select aircraft parameters: Select an aircraft model from the general UAV selection library and set the wing area S1, wing span b, wing mean aerodynamic chord Ca, wing root chord Cr, wingtip chord Ct, and wing aspect ratio A.

[0018] Step S2: Perform hydrodynamic simulation of the turbulence of the above-mentioned long-endurance UAV using the Reynolds-averaged Navier-Stokes equations; wherein, the method for obtaining the Reynolds-averaged Navier-Stokes equations is as follows:

[0019] Each explanatory variable in the instantaneous Navier-Stokes equations is decomposed into its corresponding mean value. and its pulsating component φ':

[0020]

[0021] Where φ represents any physical quantity among velocity component, pressure, energy, and substance concentration;

[0022] The decomposed average value The mean equation is obtained by inserting its pulsating component φ' into the instantaneous Navier-Stokes equations; the mean equation includes the average mass equation and the momentum transfer equation; among which...

[0023] The average mass equation is:

[0024]

[0025] The momentum transfer equation is:

[0026]

[0027] in, Let ρ represent the outer product, and ρ be the density. and These represent the average velocity and average pressure, respectively; I is the identity tensor; T is the viscous stress tensor; and f is the average velocity and average pressure, respectively. b The resultant force of various volume forces acting on a unit volume of the continuum; the additional term is a Reynolds stress tensor T. RANS Its definition is as follows:

[0028]

[0029] Among them, V′ x V′ y V′ z These represent the velocity fluctuations of the drone along the x, y, and z axes, respectively.

[0030] In a preferred embodiment, the Reynolds stress tensor T is expressed by the turbulent eddy viscosity coefficient.RANS Modeled as a function of average flow rate:

[0031]

[0032]

[0033] Where S is the average strain rate tensor;

[0034] Turbulent eddy viscosity coefficient μ t The calculation formula is as follows:

[0035]

[0036] Where μ is the hydrodynamic viscosity, k is the turbulent pulsation kinetic energy, and C μ =0.09 is a constant, f μ Let T be the damping function, and T be the turbulence time scale.

[0037] The transport equations for the turbulent pulsating kinetic energy k and the turbulent dissipation rate ε are as follows:

[0038]

[0039]

[0040] Where, σ k σ ε C ε1 C ε2 f1 is a constant, f2 is the damping function, and P is a constant. k P ε To generate the conditions, S k S ε For specified conditions, ε0 is the environmental turbulence value that counteracts turbulence attenuation;

[0041] S3: Different Reynolds stress tensors T RANS Corresponding to different aerodynamic formulas, based on the standard library consulted according to this design, approximate aerodynamic coefficient formulas were obtained. The simulated aircraft parameters were then substituted into the aerodynamic coefficient formulas:

[0042]

[0043] In the formula: L GB For lift, V GB The free-flow velocity is given by S, which is the reference area, either the wing area or the maximum cross-sectional area of ​​the fuselage.

[0044]

[0045] In the formula, D GB As resistance;

[0046]

[0047] In the formula, M GB c is the pitching moment; A The mean aerodynamic chord of the wing;

[0048] Based on formulas (3-1)-(3-3), the angle of attack is set to -5° to 15°, and the lift coefficient C is obtained. L Drag coefficient C D Pitch moment coefficient C m scope.

[0049] Calculate the lift-to-drag ratio According to equations (3-1)-(3-3), the maximum lift-to-drag ratio is obtained, which corresponds to the optimal angle of attack; a lift-to-drag ratio ≥20 indicates that the UAV has good aerodynamic characteristics and is suitable for long-endurance flight operations.

[0050] Compared with the prior art, the present invention has the following beneficial effects:

[0051] 1. By using the above aerodynamic optimization methods, the optimal angle of attack of the UAV can be accurately calculated, thereby optimizing and improving the UAV's endurance.

[0052] 2. The drone's skin is mainly made of high-strength carbon fiber composite material, while glass fiber composite material is used in special locations where the antenna is installed. Both materials have the characteristics of high strength and low density, which makes the aircraft lighter while meeting the requirements of flight.

[0053] 3. The drone truss uses carbon fiber composite panels and balsa wood sandwich composite panels, making the whole machine lighter while meeting the strength requirements.

[0054] 4. The aircraft adopts a highly integrated avionics system, which is small in size and light in weight, occupies little space, and is applicable to aircraft of all sizes and types.

[0055] 5. The mid-section of the fuselage is designed with an n-shaped structure for mounting the payload bay, which can accommodate different payloads. The payload bay can be changed by swapping it out, thus meeting the different mission requirements of the UAV. Attached Figure Description

[0056] Figure 1 This is an isometric view of the whole machine according to a preferred embodiment of the present invention;

[0057] Figure 2 This is a bottom view of the entire machine according to a preferred embodiment of the present invention;

[0058] Figure 3 This is an isometric view of the fuselage of a preferred embodiment of the present invention;

[0059] Figure 4 This is a schematic diagram of the junction of the fuselage and tail fin in a preferred embodiment of the present invention;

[0060] Figure 5 This is a schematic diagram of the tail fin structure of a preferred embodiment of the present invention;

[0061] Figure 6 This is a schematic diagram of the left wing midsection structure according to a preferred embodiment of the present invention;

[0062] Figure 7 This is a partially enlarged structural diagram of the middle section of the left wing according to a preferred embodiment of the present invention;

[0063] Figure 8 This is a schematic diagram of the housing cavity structure of the fuselage according to a preferred embodiment of the present invention;

[0064] Figure 9 This is a schematic diagram of the load chamber structure according to a preferred embodiment of the present invention;

[0065] Figure 10 This is a partially enlarged schematic diagram of the quick-release structure of the load compartment according to a preferred embodiment of the present invention;

[0066] Figure 11 This is a top view of the tail section of the left wing according to a preferred embodiment of the present invention;

[0067] Figure 12 This is a bottom-view schematic diagram of the tail section of the left wing according to a preferred embodiment of the present invention;

[0068] Figure 13 A graph showing the relevant aerodynamic forces and moment coefficients of a preferred embodiment of the present invention (I);

[0069] Figure 14 The aerodynamic force and torque coefficient curves of the preferred embodiment of the present invention are shown in Figure (II).

[0070] Figure 15 A graph (III) showing the relevant aerodynamic forces and moment coefficients of a preferred embodiment of the present invention;

[0071] Figure 16 The relevant aerodynamic force and torque coefficient curves (IV) are for a preferred embodiment of the present invention.

[0072] Reference numerals: 1-Fuselage; 11-Receiving cavity; 111-U-shaped slide rail; 12-First extended carbon fiber tube; 13-First internal carbon fiber tube; 14-Limiting hook; 15-Positioning pin; 2-Left wing midsection; 21-Lock connector; 211-Limiting ring; 212-Fixing arm; 22-Second extended carbon fiber tube; 3-Right wing midsection; 4-Left wing tail section; 41-First carbon fiber tube; 42-Third lock; 5-Right wing tail section; 6-Tail; 7-Payment bay; 71-U-shaped slide rail; 8-Avionics system; 9-Battery bay. Detailed Implementation

[0073] The present invention will be further described below with reference to the accompanying drawings and embodiments.

[0074] It should be noted that the following detailed descriptions are illustrative and intended to provide further explanation of this application. Unless otherwise specified, all technical and scientific terms used herein have the same meaning as commonly understood by one of ordinary skill in the art to which this application pertains.

[0075] It should be noted that the terminology used herein is for the purpose of describing particular implementations only and is not intended to limit the exemplary implementations according to this application; as used herein, the singular form is intended to include the plural form as well, unless the context clearly indicates otherwise; furthermore, it should be understood that when the terms “comprising” and / or “including” are used in this specification, they indicate the presence of features, steps, operations, devices, components and / or combinations thereof.

[0076] A long-endurance unmanned aerial vehicle, reference Figure 1-12 It includes the fuselage 1, the middle section of the left wing 2, the middle section of the right wing 3, the tail section of the left wing 4, the tail section of the right wing 5, and the tail 6;

[0077] The fuselage 1 is connected to the left wing mid-section 2 and the right wing mid-section 3 on both sides via an extended carbon tube assembly and a first latch, respectively.

[0078] The fuselage 1 is connected to the tail wing 6 via the second latch and the positioning pin 15. The latch is located on both sides of the tail of the fuselage 1, and the positioning pin 15 is located on the tail end face of the fuselage 1. The positioning pin 15 is symmetrically distributed along the connector. After the tail wing 6 is installed in place, the latch is locked to complete the installation of the tail wing 6.

[0079] The middle section 2 of the left wing is connected to the tail section 4 of the left wing via the first carbon tube 41 and the third latch 42; the middle section 3 of the right wing is connected to the tail section 5 of the right wing via the second carbon tube and the fourth latch; the connection between the middle section 2 of the left wing and the tail section 4 of the left wing, and the connection between the middle section 3 of the right wing and the tail section 5 of the right wing, are the same as the connection between the two sides of the fuselage 1 and the middle sections 2 and 3 of the left and right wings.

[0080] The underside of the fuselage 1 has a receiving cavity 11 for mounting the payload bay 7. The receiving cavity 11 is fixed to the payload bay 7 via a quick-release structure. The middle section of the underside of the fuselage 1 is designed with an n-shaped structure for mounting the payload bay 7, which can hold different payloads. The payload bay 7 can be replaced by changing it, thus meeting the different mission requirements of the UAV.

[0081] The quick-release structure includes: a U-shaped slide 111 and a U-shaped slide rail 71; the receiving cavity 11 is provided with U-shaped slide 111 at both ends along the length of the fuselage 1, and the top of the load compartment 7 is provided with U-shaped slide rail 71 that mates with the U-shaped slide 111 at both ends; the length of the U-shaped slide 111 is parallel to the length of the U-shaped slide rail 71, and the U-shaped slide 111 and the U-shaped slide rail 71 mutually limit each other along the direction of the fuselage 1 so that the load compartment 7 is detachably connected to the fuselage 1.

[0082] The openings 711 between the mutually adaptable U-shaped slide groove 111 and the U-shaped slide rail 71 are arranged opposite to each other.

[0083] The U-shaped slide 111 and the U-shaped slide rail 71 are respectively fixed to the fuselage 1 and the load compartment 7 with screws.

[0084] The top of the payload bay 7 slides into the U-shaped groove 111 of the fuselage 1 via two U-shaped rails 71, and automatically locks in place after installation. The payload bay 7's body and bottom are used to install payloads. Depending on different mission requirements, the internal mounting bracket structure of the payload bay 7 can be adjusted to accommodate different payloads (pods, radar, etc.). The payload bay 7 is installed in the middle of the fuselage 1, so even changing different payloads has minimal impact on the aircraft's overall center of gravity and moment of inertia. Therefore, by adjusting the structure of the payload bay 7 and replacing it with different payload bays, different payloads can be mounted without altering the fuselage 1 or other structures.

[0085] The extended carbon fiber assembly includes first extended carbon fiber tubes 12 symmetrically extended outward on both sides of the fuselage 1, a first internal carbon fiber tube 13 disposed inside the fuselage 1, and second extended carbon fiber tubes 22 extending outward at the ends of the left wing midsection 2 and right wing midsection 3 facing the fuselage 1, respectively. The second internal carbon fiber tubes are respectively disposed at the ends of the left wing midsection 2 and right wing midsection 3 facing the fuselage 1. The first extended carbon fiber tubes 12 are inserted into the second internal carbon fiber tubes, and the second extended carbon fiber tubes 22 are inserted into the first internal carbon fiber tubes 13. The number of carbon fiber tubes and locking devices is not limited to that shown in the figure and can be adjusted appropriately according to actual needs.

[0086] The first, second, third, and fourth latches each include a latch connector 21 and a latch connecting fitting that mates with the latch connector 21. The latch connector 21 is located on the left wing midsection 2, the right wing midsection 3, the left wing tail section 4, the right wing tail section 5, and the tail 6. The latch connecting fitting is located on the fuselage 1, the left wing midsection 2, and the right wing midsection 3.

[0087] The locking connector 21 on the left wing mid-section 2 and the locking connector 21 on the right wing mid-section 3 are both located at the end facing the fuselage 1. Two locking connectors on the fuselage 1, which mate with the locking connectors 21 on the right wing mid-section 3 and the left wing mid-section 2 respectively, are located on both sides of the fuselage 1. The locking connector 21 includes a limiting ring 211, a fixing arm 212, and a torsion spring. The fixing arm 212 is connected to the limiting ring 211 by a torsion spring. A pivot is provided in the middle of the torsion spring, and the pivot is connected to the left wing mid-section, the right wing mid-section, the left wing tail section 4, the right wing tail section 5, and the tail fin respectively. 6. Fixed connection; Specifically, the two ends of the rotating shaft are fixed to the middle section of the left wing, the middle section of the right wing, the tail section 4 of the left wing, the tail section 5 of the right wing, and the tail fin 6, respectively; the locking connection is a limiting hook 14; the limiting hook 14 and the limiting ring 211 are in a limiting engagement; the fixed arm 212 rotates in the first direction, causing the limiting ring 211 to move away from the limiting hook 14, so that the limiting ring 211 and the limiting hook 14 are more firmly limited, at which time the torsion spring is in the normal state; the fixed arm 212 rotates in the second direction, causing the limiting ring 211 to move towards the limiting hook 14, so that the limiting ring 211 and the limiting hook 14 are no longer mutually limiting.

[0088] When the fuselage 1 is connected to the middle section 2 of the left wing and the middle section 3 of the right wing, the limiting hook 14 and the limiting ring 211 mutually limit and cooperate in a direction perpendicular to the fuselage 1; when the fuselage 1 is connected to the tail 6, the limiting hook 14 and the limiting ring 211 mutually limit and cooperate in a direction parallel to the fuselage 1.

[0089] Specifically, the tail of the fuselage 1 is provided with a positioning pin 15, and the tail wing 6 is provided with a positioning hole. After the positioning pin 15 is inserted into the positioning hole, the fuselage 1 and the tail wing 6 are connected by a latch.

[0090] Installation process: The left wing mid-section 2 and the right wing mid-section 3 are connected to the two sides of the fuselage 1 through the extended carbon tube assembly and the first locking buckle, respectively. When installing the left wing mid-section 2 and the right wing mid-section 3, the extended carbon tube of the fuselage 1 is inserted into the interior of the left wing mid-section 2 and the right wing mid-section 3. At the same time, the extended carbon tube of the left wing mid-section 2 and the right wing mid-section 3 is inserted into the rear carbon tube flush with the fuselage 1. After the left wing mid-section 2 and the right wing mid-section 3 are installed in place, the locking buckle is locked, thus completing the installation of the left wing mid-section 2 and the right wing mid-section 3.

[0091] The drone's overall support structure is made of high-strength carbon fiber plates and balsa wood; the fuselage skin is primarily made of high-strength carbon fiber composite materials, with some canopies using fiberglass composite materials. Both materials possess characteristics such as high strength and low density, allowing for lighter weight while meeting flight requirements. The drone's truss utilizes carbon fiber composite panels and balsa wood sandwich composite panels, further enhancing the overall weight while maintaining sufficient strength.

[0092] The forward section of fuselage 1 houses the battery compartment 9, and the aft section houses the avionics compartment, used for the installation of batteries and avionics systems 8, respectively. The avionics system 8, which generates significant heat, is fixed to the bottom of fuselage 1 within the avionics compartment. Antenna mounting covers are located at both the front and rear of fuselage 1 for antenna installation. Two landing gears, a front landing gear and a rear landing gear, are designed at the lower part of fuselage 1, with the antenna embedded within the rear landing gear. Avionics system 8 features high integration, light weight, and small size. Heat sinks are also designed on the bottom mounting surface to aid in heat dissipation for avionics system 8.

[0093] Specifically, the wingtip shape is upward-curved, and the tail fin type is V-shaped.

[0094] This invention also provides an aerodynamic optimization method for a long-endurance unmanned aerial vehicle (UAV), which employs a long-endurance UAV and includes the following steps:

[0095] Step S1: Select aircraft parameters: Select an aircraft model from the general UAV selection library and set the wing area S1, wing span b, wing mean aerodynamic chord Ca, wing root chord Cr, wingtip chord Ct, and wing aspect ratio A.

[0096] Step S2: Perform hydrodynamic simulation of the turbulence of the above-mentioned long-endurance UAV using the Reynolds-averaged Navier-Stokes equations; wherein, the method for obtaining the Reynolds-averaged Navier-Stokes equations is as follows:

[0097] Each explanatory variable in the instantaneous Navier-Stokes equations is decomposed into its corresponding mean value. and its pulsating component φ':

[0098]

[0099] Where φ represents any physical quantity among velocity component, pressure, energy, and substance concentration;

[0100] The decomposed average value The mean equation is obtained by inserting its pulsating component φ' into the instantaneous Navier-Stokes equations; the mean equation includes the average mass equation and the momentum transfer equation; among which...

[0101] The average mass equation is:

[0102]

[0103] The momentum transfer equation is:

[0104]

[0105] in, Let ρ represent the outer product, and ρ be the density. and These represent the average velocity and average pressure, respectively; I is the identity tensor; T is the viscous stress tensor; and f is the average velocity and average pressure, respectively.b The resultant force of various volume forces acting on a unit volume of the continuum; the additional term is a Reynolds stress tensor T. RANS Its definition is as follows:

[0106]

[0107] Among them, V′ x V′ y V′ z These represent the velocity fluctuations of the UAV along the x, y, and z axes, respectively.

[0108] Aircraft design requires high-precision simulation of the viscous effects of flow fields to accurately calculate the aerodynamic characteristics of aircraft with different shapes. Incorporating a turbulence model into the Reynolds-averaged RANS equations is the mainstream approach for modern numerical simulation of complex viscous flow fields. The RANS equations with the turbulence model improved accuracy in calculating the aerodynamic forces and moments acting on the aircraft.

[0109] For T RANS The difficulty in solving the formula lies in understanding the relationship between the average flow rate and T. RANS Modeling is performed to close the governing equations. The Reynolds stress tensor T can be expressed using the turbulent eddy viscosity coefficient. RANS Modeling it as a function of average flow rate, the most common model is called the Boussinesq approximation (a kinetic term):

[0110]

[0111] Where S is the average strain rate tensor;

[0112] Turbulent eddy viscosity coefficient μ t The calculation formula is as follows:

[0113] μ t =ρC μ f μ kT

[0114] Where μ is the hydrodynamic viscosity, k is the turbulent pulsation kinetic energy, and C μ =0.09 is a constant, f μ Let T be the damping function, and T be the turbulence time scale.

[0115] The transport equations for the turbulent pulsating kinetic energy k and the turbulent dissipation rate ε are as follows:

[0116]

[0117]

[0118] Where, σ k σ ε C ε1 C ε2 f1 is a constant, f2 is the damping function, and P is a constant. k P ε To generate the conditions, S k S ε For specified conditions, ε0 is the environmental turbulence value that counteracts turbulence attenuation;

[0119] S3: Different Reynolds stress tensors T RANS Corresponding to different aerodynamic formulas, based on the standard library consulted according to this design, approximate aerodynamic coefficient formulas were obtained. The simulated aircraft parameters were then substituted into the aerodynamic coefficient formulas:

[0120]

[0121] In the formula: L GB For lift, V GB The free-flow velocity is given by S, which is the reference area, either the wing area or the maximum cross-sectional area of ​​the fuselage.

[0122]

[0123] In the formula, D GB As resistance;

[0124]

[0125] In the formula, M GB c is the pitching moment; A The mean aerodynamic chord of the wing;

[0126] Based on formulas (3-1)-(3-3), the angle of attack is set to -5° to 15°, and the lift coefficient C is obtained. L Drag coefficient C D Pitch moment coefficient C m scope.

[0127] Calculate the lift-to-drag ratio According to equations (3-1)-(3-3), the maximum lift-to-drag ratio is obtained, which corresponds to the optimal angle of attack; a lift-to-drag ratio ≥20 indicates that the UAV has good aerodynamic characteristics and is suitable for long-endurance flight operations.

[0128] Since the wind tunnel test data obtained are aerodynamic coefficients, referring to GB / T 16638.4-2008 Aerodynamics: Concepts, Quantities and Symbols - Part 4: Aerodynamics, Moments and Their Coefficients and Derivatives of Aircraft, to avoid confusion with the symbols in this text, the symbols defined in the national standard are annotated with subscripts and accompanied by textual explanations. The aerodynamic coefficient formulas are then obtained, as shown in Table 1.

[0129] Table 1 Formulas for Aerodynamic Coefficients

[0130]

[0131] Based on the formulas in Table 1, the relevant aerodynamic force and moment coefficient curves are calculated as follows: Figure 13-16 As shown.

[0132] Based on the full-turbulence CFD method, the overall aerodynamic characteristics of the UAV are obtained as follows:

[0133]

[0134]

[0135] In the table above, Aoa represents the angle of attack, and C... D C represents the drag coefficient. L C represents the lift coefficient. M This represents the pitching moment coefficient.

[0136] Based on the fully turbulent CFD method, the aerodynamic characteristics of the UAV provided in this application are summarized as follows:

[0137] 1) With a cruise lift coefficient of 0.7, the angle of attack is approximately 2.5°, resulting in relatively low drag;

[0138] 2) At an angle of attack of approximately 3°, the lift-to-drag ratio reaches 21.9.

[0139] The aerodynamic characteristics of the UAV obtained through CFD simulation show that the aerodynamic characteristics of the UAV in this application meet the design requirements.

Claims

1. An aerodynamic optimization method for a long-endurance unmanned aerial vehicle (UAV), characterized in that, Includes the following steps: Step S1: Select aircraft parameters: Select an aircraft model from the general UAV selection library and set the wing area S1, wing span b, wing mean aerodynamic chord Ca, wing root chord Cr, wingtip chord Ct, and wing aspect ratio A. Step S2: Perform hydrodynamic simulation of the turbulence of the above-mentioned long-endurance UAV using the Reynolds-averaged Navier-Stokes equations. The method for obtaining the Reynolds-averaged Navier-Stokes equations is as follows: Each explanatory variable in the instantaneous Navier-Stokes equations is decomposed into its corresponding mean value. and its pulsating components : ,in, It represents any physical quantity among velocity component, pressure, energy, and substance concentration; The decomposed average value and its pulsating components Inserting the instantaneous Navier-Stokes equations yields the mean equation; this mean equation includes the average mass equation and the momentum transfer equation; among which, The average mass equation is: The momentum transfer equation is: ,in, Indicates the outer product. For density, and Let I be the average velocity and average pressure, respectively; let I be the identity tensor and T be the viscous stress tensor. The resultant force of all volume forces acting on a unit volume of a continuum; Additional terms It is a Reynolds stress tensor, defined as follows: ,in, , , These represent the velocity fluctuations of the UAV along the x, y, and z axes, respectively. The Reynolds stress tensor is derived from the turbulent eddy viscosity coefficient. Modeled as a function of average flow rate: , , where S is the average strain rate tensor; Turbulent eddy viscosity coefficient The calculation formula is as follows: ,in, For fluid dynamic viscosity, For turbulent pulsating kinetic energy, =0.09 is a constant. Let T be the damping function, and T be the turbulence time scale. The transport equations for the turbulent pulsating kinetic energy k and the turbulent dissipation rate ε are as follows: , ,in, , , , It is a constant. Let be the damping function. , To generate the conditions, , To specify conditions, To compensate for the attenuation of environmental turbulence values; S3: Different Reynolds stress tensors Corresponding to different aerodynamic formulas, based on the standard library consulted according to this design, approximate aerodynamic coefficient formulas were obtained. The simulated aircraft parameters were then substituted into the aerodynamic coefficient formulas: (3-1) In the formula: For lift, The free-flow velocity is given by S, which is the reference area, either the wing area or the maximum cross-sectional area of ​​the fuselage. (3-2) In the formula, As resistance; (3-3) In the formula, For pitching moment; The mean aerodynamic chord of the wing; Based on formulas (3-1)-(3-3), the angle of attack is set to -5° to 15°, and the lift coefficient C is obtained. L Drag coefficient C D Pitch moment coefficient C m scope; Calculate the lift-to-drag ratio According to equations (3-1)-(3-3), the maximum lift-to-drag ratio is obtained, which corresponds to the optimal angle of attack; a lift-to-drag ratio ≥20 indicates that the UAV has good aerodynamic characteristics and is suitable for long-endurance flight operations.

2. A long-endurance unmanned aerial vehicle (UAV) manufactured using the aerodynamic optimization method described in claim 1, characterized in that, include: The fuselage, the middle section of the left wing, the middle section of the right wing, the tail section of the left wing, the tail section of the right wing, and the tail fin; The fuselage is connected to the left wing midsection and the right wing midsection respectively via an extended carbon tube assembly and a first latch; The fuselage is connected to the tail fin via a second latch and a locating pin; The middle section of the left wing is connected to the tail section of the left wing via a first carbon fiber tube and a third locking device; the middle section of the right wing is connected to the tail section of the right wing via a second carbon fiber tube and a fourth locking device. The fuselage has a cavity in its underside for mounting a payload bay. The receiving cavity is provided with U-shaped grooves at both ends along the length of the fuselage, and the top of the load compartment is provided with U-shaped rails at both ends that mate with the U-shaped grooves; the length of the U-shaped grooves is parallel to the length of the U-shaped rails, and the U-shaped grooves and U-shaped rails are mutually restrained along the length of the fuselage so that the load compartment is detachably connected to the fuselage; the openings between the mutually adaptable U-shaped grooves and U-shaped rails are arranged opposite to each other.

3. The long-endurance UAV manufactured using the aerodynamic optimization method for long-endurance UAVs according to claim 2, characterized in that, The extended carbon fiber assembly includes a first extended carbon fiber symmetrically extending outward on both sides of the fuselage, a first internal carbon fiber symmetrically disposed inside the fuselage, a second extended carbon fiber symmetrically extending outward at the mid-section of the left wing and the mid-section of the right wing facing the fuselage, and a second internal carbon fiber symmetrically disposed at the mid-section of the left wing and the mid-section of the right wing facing the fuselage; the first extended carbon fiber symmetrically inserted into the second internal carbon fiber symmetrically inserted into the first internal carbon fiber symmetrically inserted into the first internal carbon fiber symmetrically inserted into the second ... second internal carbon fiber symmetrically inserted into the 4. The long-endurance UAV manufactured using the aerodynamic optimization method for long-endurance UAVs according to claim 2, characterized in that, The first, second, third, and fourth latches each include a latch connector and a latch connecting fitting that cooperates with the latch connector. The latch connector is disposed on the middle section of the left wing, the middle section of the right wing, the tail section of the left wing, the tail section of the right wing, and the tail fin. The latch connecting fitting is disposed on the fuselage, the middle section of the left wing, and the middle section of the right wing.

5. The long-endurance UAV manufactured using the aerodynamic optimization method for long-endurance UAVs according to claim 4, characterized in that, The locking connectors on the middle section of the left wing and the middle section of the right wing are both located at the end facing the fuselage. Two locking connectors that mate with the locking connectors on the middle sections of the right and left wings are located on both sides of the fuselage.

6. The long-endurance UAV manufactured using the aerodynamic optimization method for long-endurance UAVs according to claim 5, characterized in that, The locking connector includes: a limiting ring, a fixing arm, and a torsion spring; the fixing arm and the limiting ring are connected by the torsion spring; a rotating shaft is provided in the middle of the torsion spring, and the rotating shaft is fixedly connected to the middle section of the left wing, the middle section of the right wing, the tail section of the left wing, the tail section of the right wing, and the tail fin; the locking connector is a limiting hook; the limiting hook is limited and engaged with the limiting ring.

7. The long-endurance UAV manufactured using the aerodynamic optimization method for long-endurance UAVs according to claim 2, characterized in that, The forward section of the fuselage houses the battery compartment, while the aft section houses the avionics compartment. The battery compartment and avionics compartment are used to install batteries and avionics systems, respectively. Antenna mounting covers for installing antennas are located at the front and rear of the fuselage. The lower part of the fuselage has front and rear landing gears, with the antenna embedded inside the rear landing gear.

Images