CN117622552A - Long-endurance unmanned aerial vehicle and pneumatic optimization method thereof - Google Patents

Long-endurance unmanned aerial vehicle and pneumatic optimization method thereof Download PDF

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Publication number
CN117622552A
CN117622552A CN202311838994.XA CN202311838994A CN117622552A CN 117622552 A CN117622552 A CN 117622552A CN 202311838994 A CN202311838994 A CN 202311838994A CN 117622552 A CN117622552 A CN 117622552A
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China
Prior art keywords
wing
middle section
tail
unmanned aerial
aerial vehicle
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CN202311838994.XA
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Inventor
陈伯建
韩腾飞
张伟豪
陈金玉
吴晓杰
吴文斌
梁曼舒
王仁书
李哲舟
陈卓磊
程海涛
王淼
王泽昭
徐鹏鹏
孔祥玉
董晓峰
蔡焕青
谈家英
付晶
林永琎
赵广炎
张占斌
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Beijing Yuandu Internet Technology Co ltd
State Grid Power Space Technology Co ltd
State Grid Corp of China SGCC
China Electric Power Research Institute Co Ltd CEPRI
Electric Power Research Institute of State Grid Fujian Electric Power Co Ltd
State Grid Fujian Electric Power Co Ltd
State Grid Qinghai Electric Power Co Ltd
Original Assignee
Beijing Yuandu Internet Technology Co ltd
State Grid Power Space Technology Co ltd
State Grid Corp of China SGCC
China Electric Power Research Institute Co Ltd CEPRI
Electric Power Research Institute of State Grid Fujian Electric Power Co Ltd
State Grid Fujian Electric Power Co Ltd
State Grid Qinghai Electric Power Co Ltd
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Application filed by Beijing Yuandu Internet Technology Co ltd, State Grid Power Space Technology Co ltd, State Grid Corp of China SGCC, China Electric Power Research Institute Co Ltd CEPRI, Electric Power Research Institute of State Grid Fujian Electric Power Co Ltd, State Grid Fujian Electric Power Co Ltd, State Grid Qinghai Electric Power Co Ltd filed Critical Beijing Yuandu Internet Technology Co ltd
Priority to CN202311838994.XA priority Critical patent/CN117622552A/en
Publication of CN117622552A publication Critical patent/CN117622552A/en
Pending legal-status Critical Current

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    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64U—UNMANNED AERIAL VEHICLES [UAV]; EQUIPMENT THEREFOR
    • B64U20/00—Constructional aspects of UAVs
    • B64U20/60—UAVs characterised by the material
    • B64U20/65—Composite materials
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C1/00—Fuselages; Constructional features common to fuselages, wings, stabilising surfaces or the like
    • B64C1/0009—Aerodynamic aspects
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C1/00—Fuselages; Constructional features common to fuselages, wings, stabilising surfaces or the like
    • B64C1/26—Attaching the wing or tail units or stabilising surfaces
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C1/00—Fuselages; Constructional features common to fuselages, wings, stabilising surfaces or the like
    • B64C1/36—Fuselages; Constructional features common to fuselages, wings, stabilising surfaces or the like adapted to receive antennas or radomes
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64F—GROUND OR AIRCRAFT-CARRIER-DECK INSTALLATIONS SPECIALLY ADAPTED FOR USE IN CONNECTION WITH AIRCRAFT; DESIGNING, MANUFACTURING, ASSEMBLING, CLEANING, MAINTAINING OR REPAIRING AIRCRAFT, NOT OTHERWISE PROVIDED FOR; HANDLING, TRANSPORTING, TESTING OR INSPECTING AIRCRAFT COMPONENTS, NOT OTHERWISE PROVIDED FOR
    • B64F5/00—Designing, manufacturing, assembling, cleaning, maintaining or repairing aircraft, not otherwise provided for; Handling, transporting, testing or inspecting aircraft components, not otherwise provided for
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64U—UNMANNED AERIAL VEHICLES [UAV]; EQUIPMENT THEREFOR
    • B64U20/00—Constructional aspects of UAVs
    • B64U20/70—Constructional aspects of the UAV body
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64U—UNMANNED AERIAL VEHICLES [UAV]; EQUIPMENT THEREFOR
    • B64U20/00—Constructional aspects of UAVs
    • B64U20/80—Arrangement of on-board electronics, e.g. avionics systems or wiring
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64U—UNMANNED AERIAL VEHICLES [UAV]; EQUIPMENT THEREFOR
    • B64U30/00—Means for producing lift; Empennages; Arrangements thereof
    • B64U30/10—Wings
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64U—UNMANNED AERIAL VEHICLES [UAV]; EQUIPMENT THEREFOR
    • B64U30/00—Means for producing lift; Empennages; Arrangements thereof
    • B64U30/40—Empennages, e.g. V-tails
    • G—PHYSICS
    • G06—COMPUTING; CALCULATING OR COUNTING
    • G06F—ELECTRIC DIGITAL DATA PROCESSING
    • G06F30/00—Computer-aided design [CAD]
    • G06F30/10—Geometric CAD
    • G06F30/15—Vehicle, aircraft or watercraft design
    • G—PHYSICS
    • G06—COMPUTING; CALCULATING OR COUNTING
    • G06F—ELECTRIC DIGITAL DATA PROCESSING
    • G06F30/00—Computer-aided design [CAD]
    • G06F30/20—Design optimisation, verification or simulation
    • G06F30/28—Design optimisation, verification or simulation using fluid dynamics, e.g. using Navier-Stokes equations or computational fluid dynamics [CFD]
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C1/00—Fuselages; Constructional features common to fuselages, wings, stabilising surfaces or the like
    • B64C2001/0054—Fuselage structures substantially made from particular materials
    • B64C2001/0072—Fuselage structures substantially made from particular materials from composite materials
    • G—PHYSICS
    • G06—COMPUTING; CALCULATING OR COUNTING
    • G06F—ELECTRIC DIGITAL DATA PROCESSING
    • G06F2113/00—Details relating to the application field
    • G06F2113/08—Fluids
    • G—PHYSICS
    • G06—COMPUTING; CALCULATING OR COUNTING
    • G06F—ELECTRIC DIGITAL DATA PROCESSING
    • G06F2119/00—Details relating to the type or aim of the analysis or the optimisation
    • G06F2119/14—Force analysis or force optimisation, e.g. static or dynamic forces

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  • Engineering & Computer Science (AREA)
  • Aviation & Aerospace Engineering (AREA)
  • Physics & Mathematics (AREA)
  • Mechanical Engineering (AREA)
  • Remote Sensing (AREA)
  • Geometry (AREA)
  • General Physics & Mathematics (AREA)
  • Theoretical Computer Science (AREA)
  • General Engineering & Computer Science (AREA)
  • Fluid Mechanics (AREA)
  • Computer Hardware Design (AREA)
  • Evolutionary Computation (AREA)
  • Mathematical Optimization (AREA)
  • Mathematical Analysis (AREA)
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  • Mathematical Physics (AREA)
  • Computing Systems (AREA)
  • Algebra (AREA)
  • Manufacturing & Machinery (AREA)
  • Transportation (AREA)
  • Computational Mathematics (AREA)
  • Microelectronics & Electronic Packaging (AREA)
  • Automation & Control Theory (AREA)
  • Chemical & Material Sciences (AREA)
  • Composite Materials (AREA)
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  • Aiming, Guidance, Guns With A Light Source, Armor, Camouflage, And Targets (AREA)

Abstract

The invention provides a long-endurance unmanned plane and a pneumatic optimization method thereof, comprising the following steps: the device comprises 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; two sides of the fuselage are respectively connected with the middle section of the left wing and the middle section of the right wing through the overhanging carbon tubes and the lock catches; the machine body is connected with the tail wing through lock catches and positioning pins, the lock catches are locked on two sides of the tail part of the machine body, the positioning pins are arranged on the end face of the tail part of the machine body, the positioning pins are symmetrically distributed along the connector, and the lock catches are locked after the tail wing is installed in place, so that the installation of the tail wing is completed; the middle section of the left wing is connected with the tail section of the left wing through a carbon tube and a lock catch; the middle section of the right wing is connected with the tail section of the right wing through a carbon tube and a lock catch; the belly of the machine body is provided with a containing cavity, the containing cavity is used for installing a load cabin, and the containing cavity is fixed with the load cabin through a quick-release structure; by the aid of the technical scheme, unmanned aerial vehicle endurance time can be improved.

Description

Long-endurance unmanned aerial vehicle and pneumatic optimization method thereof
Technical Field
The invention relates to the technical field of unmanned aerial vehicles, in particular to a long-endurance unmanned aerial vehicle and a pneumatic optimization method thereof.
Background
The unmanned aerial vehicle is an unmanned aerial vehicle, is abbreviated as an UAV, is an unmanned aerial vehicle operated by using radio remote control equipment and a self-provided program control device, and is widely applied to the fields of communication relay, aviation shooting, resource exploration, military and the like along with the rapid development of electronic technology and materials.
The unmanned aerial vehicle mainly comprises structural components such as fuselage, wing, fin, and unmanned aerial vehicle body structure is unmanned aerial vehicle's truck and atress basis, not only needs other parts of fixed and support unmanned aerial vehicle, connects whole unmanned aerial vehicle into a whole, still bears the load that each connecting element transmitted, bears the equipment, task load and own gravity and inertia of load in the fuselage inside. Therefore, the unmanned aerial vehicle is large and heavy, and the unmanned aerial vehicle has short endurance time. In addition, the task load on the unmanned aerial vehicle is usually with unmanned aerial vehicle fixed connection, leads to unmanned aerial vehicle task singleness, and different unmanned aerial vehicles must be changed to accomplish different tasks, and is with high costs. The actual endurance time of the existing vertical take-off and landing unmanned aerial vehicle is 60-90 minutes at present, and the requirement of long-distance power line inspection operation is difficult to meet.
Disclosure of Invention
Therefore, the invention aims to provide the unmanned aerial vehicle with long endurance and the pneumatic optimization method thereof, so as to realize the improvement of the endurance time of the unmanned aerial vehicle.
In order to achieve the above purpose, the invention adopts the following technical scheme: a long-endurance unmanned aerial vehicle, comprising: the device comprises 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;
two sides of the fuselage are respectively connected with the left wing middle section and the right wing middle section through an overhanging carbon tube assembly and a first lock catch;
the machine body is connected with the tail wing through a second lock catch and a locating pin;
the middle section of the left wing is connected with the tail section of the left wing through a first carbon tube and a third lock catch; the middle section of the right wing is connected with the tail section of the right wing through a second carbon tube and a fourth lock catch;
the belly of fuselage is provided with holds the chamber, hold the chamber and be used for installing the load cabin.
In a preferred embodiment, two ends of the accommodating cavity along the length direction of the machine body are respectively provided with a U-shaped sliding groove, and two ends of the top of the load cabin are respectively provided with a U-shaped sliding rail matched with the U-shaped sliding grooves; the length direction of the U-shaped sliding groove is parallel to the length direction of the U-shaped sliding rail, and the U-shaped sliding groove and the U-shaped sliding rail are mutually limited along the direction of the machine body so that the load cabin is detachably and fixedly connected with the machine body; the opening between the mutually adapted U-shaped sliding groove and the U-shaped sliding rail is arranged oppositely.
In a preferred embodiment, the overhanging carbon tube assembly comprises a first overhanging carbon tube, a first internal carbon tube, a second overhanging carbon tube and a second internal carbon tube, wherein the first overhanging carbon tube is arranged at two sides of the fuselage and symmetrically extends outwards; the first overhanging carbon tube is inserted into the second inner carbon tube, and the second overhanging carbon tube is inserted into the first inner carbon tube.
In a preferred embodiment, the first lock, the second lock, the third lock, and the fourth lock each include: the device comprises a lock catch connecting piece and a lock catch connecting matching piece matched with the lock catch connecting piece, wherein the lock catch connecting piece is arranged on a left wing middle section, a right wing middle section, a left wing tail section, a right wing tail section and a tail wing, and the lock catch connecting matching piece is arranged on the fuselage, the left wing middle section and the right wing middle section.
In a preferred embodiment, the latch connector disposed on the middle section of the left wing and the latch connector disposed on the middle section of the right wing are disposed at one end facing the fuselage, and two latch connector fittings disposed on the fuselage and respectively engaged with the latch connectors of the middle section of the right wing and the middle section of the left wing are disposed at two sides of the fuselage.
In a preferred embodiment, the stop collar, the fixed arm, and the torsion spring; the fixed arm is connected with the limit ring through a torsion spring; the middle part of the torsion spring is provided with a rotating shaft which is fixedly connected with the left wing middle section, the right wing middle section, the left wing tail section, the right wing tail section and the tail wing respectively; the lock catch connecting fitting is a limiting hook; the limit hook is in limit fit with the limit ring.
In a preferred embodiment, the front section of the fuselage is provided with a battery compartment, the rear section of the fuselage is provided with an avionic compartment, and the battery compartment and the avionic compartment are respectively used for installing a battery and an avionic system; the front end and the rear end of the machine body are respectively provided with an antenna mounting cover for mounting an antenna; front foot frames and rear foot frames are arranged in front of and behind the lower portion of the machine body, and antennas are buried in the rear foot frames.
The invention also provides a pneumatic optimization method of the long-endurance unmanned aerial vehicle, which adopts the long-endurance unmanned aerial vehicle and comprises the following steps:
step S1: selecting model parameters: selecting one model from a general unmanned aerial vehicle selection library, and setting wing area S 1 Wing extension b,Average aerodynamic chord length Ca, root chord length Cr, tip chord length Ct, and wing aspect ratio A;
step S2: carrying out fluid mechanics simulation on the turbulence of the long-endurance unmanned aerial vehicle by utilizing a Reynolds average N-S equation; the method for acquiring the Reynolds average N-S equation comprises the following steps:
decomposing each interpretation variable in the instantaneous N-S equation into a corresponding average valueA pulsation component phi':
wherein phi represents any one of a velocity component, pressure, energy, and substance concentration;
mean value after decompositionAnd the pulsation component phi' is inserted into an instantaneous N-S equation to obtain a mean equation; the mean value equation comprises an average mass equation and a momentum transfer equation; wherein,
the average mass equation is:
the momentum transfer equation is:
wherein,representing the outer product, ρ is the density, +.>And->Average velocity and average pressure, respectively, I is the identity tensor, T is the viscous stress tensor, f b Is the resultant of various volume forces acting on the unit volume of the continuum; the additional term is a Reynolds stress tensor T RANS It is defined as follows:
wherein V 'is' x 、V′ y 、V′ z The speed pulsations of the drone along the x, y, z axes, respectively.
In a preferred embodiment, the Reynolds stress tensor T is determined by the turbulent eddy current viscosity coefficient RANS Modeling as a function of average flow:
wherein S is the average strain rate tensor;
turbulent vortex viscosity coefficient mu t The calculation formula of (2) is as follows:
wherein μ is hydrodynamic viscosity, k is turbulent pulsating kinetic energy, C μ =0.09 as a constant, f μ T is the turbulence time scale as a damping function;
the transport equation for turbulent pulsation kinetic energy k and turbulent dissipation rate ε is:
wherein sigma k 、σ ε 、C ε1 、C ε2 Is constant, f 2 As a damping function, P k 、P ε To generate conditions S k 、S ε To specify conditions ε 0 An ambient turbulence value to counteract turbulence attenuation;
s3: different reynolds stress tensors T RANS According to different aerodynamic formulas, a standard library is queried according to the design to obtain an approximate aerodynamic coefficient formula, and model parameters after simulation are substituted into the aerodynamic coefficient formula:
wherein: l (L) GB Is the lifting force, V GB S is a reference area, and the wing area or the maximum cross-sectional area of the fuselage is taken as the free flow speed;
wherein D is GB Is resistance;
wherein M is GB Is pitching moment; c A The average aerodynamic chord length of the wing;
setting the attack angle to be-5 degrees to 15 degrees according to formulas (3-1) - (3-3) to obtain a lift coefficient C L Coefficient of resistance C D Pitch moment coefficient C m Range.
Calculation of lift-to-drag ratioObtaining a maximum lift-drag ratio corresponding to an optimal attack angle according to formulas (3-1) - (3-3); the lift-drag ratio is more than or equal to 20, namely the unmanned aerial vehicle is considered to have better aerodynamic characteristics, and is suitable for long-endurance flight operation.
Compared with the prior art, the invention has the following beneficial effects:
1. by the pneumatic optimization method, the optimal attack angle of the unmanned aerial vehicle can be accurately calculated, and the duration of the unmanned aerial vehicle is optimized and improved.
2. The unmanned aerial vehicle skin mainly adopts high strength carbon fiber composite material, adopts glass fiber composite material in the special position of antenna installation, and two kinds of materials possess characteristics such as intensity is high, and density is little, satisfies weight lighter under the aircraft flight prerequisite.
3. The unmanned aerial vehicle truss adopts the carbon fiber composite material board and the balsawood sandwich composite material board, makes the complete machine weight lighter under the condition that intensity satisfies the requirement.
4. The whole machine adopts a highly integrated avionics system, and the system has small volume, light weight and small occupied space and can be suitable for airplanes of various sizes.
5. The middle section of the belly is designed into an n-shaped structure and is used for installing load cabins, and the load cabins can be used for installing different loads. The load cabin can be replaced by replacing the load cabin, so that different task demands of the unmanned aerial vehicle are met.
Drawings
FIG. 1 is an isometric view of a complete machine according to a preferred embodiment of the present invention;
FIG. 2 is a bottom view of the complete machine of the preferred embodiment of the present invention;
FIG. 3 is an isometric view of a fuselage of a preferred embodiment of the present invention;
FIG. 4 is a schematic view of the attachment of the fuselage to the tail wing in accordance with the preferred embodiment of the present invention;
fig. 5 is a schematic view of a tail structure according to a preferred embodiment of the invention;
FIG. 6 is a schematic view of the left wing midsection configuration of a preferred embodiment of the invention;
FIG. 7 is a schematic view of a partial enlarged configuration of the left wing midsection of a preferred embodiment of the invention;
FIG. 8 is a schematic view of the structure of the accommodation chamber of the body according to the preferred embodiment of the present invention;
FIG. 9 is a schematic view of the structure of a load compartment according to a preferred embodiment of the invention;
FIG. 10 is an enlarged partial schematic view of a quick release structure of a load compartment according to a preferred embodiment of the invention;
FIG. 11 is a schematic top view of a left wing tail section of a preferred embodiment of the present invention;
FIG. 12 is a schematic view of the left wing tail section from below in accordance with the preferred embodiment of the present invention;
FIG. 13 is a graph (I) of the aerodynamic force and moment coefficients associated with a preferred embodiment of the present invention;
FIG. 14 is a graph (II) of the aerodynamic force and moment coefficients associated with a preferred embodiment of the present invention;
FIG. 15 is a graph (III) of the aerodynamic force and moment coefficients associated with a preferred embodiment of the present invention;
FIG. 16 is a graph (IV) of aerodynamic forces and moment coefficients associated with a preferred embodiment of the present invention.
Reference numerals: 1-a fuselage; 11-a receiving cavity; 111-U-shaped sliding grooves; 12-a first overhanging carbon tube; 13-a first inner carbon tube; 14-limiting hooks; 15-locating pins; 2-the middle section of the left wing; 21-a latch connection; 211-limit rings; 212-a fixed arm; 22-a second overhanging carbon tube; 3-the middle section of the right wing; 4-left wing tail section; 41-a first carbon tube; 42-a third lock catch; 5-right wing tail section; 6-tail fin; 7-loading bay; 71-U-shaped slide rails; 8-avionics system; 9-battery compartment.
Detailed Description
The invention will be further described with reference to the accompanying drawings and examples.
It should be noted that the following detailed description is illustrative and is intended to provide further explanation of the present application. Unless defined otherwise, 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 belongs.
It is noted that the terminology used herein is for the purpose of describing particular embodiments only and is not intended to be limiting of example embodiments in accordance with the present application; as used herein, the singular is also intended to include the plural unless the context clearly indicates otherwise, and furthermore, it is to be understood that the terms "comprises" and/or "comprising" when used in this specification are taken to specify the presence of stated features, steps, operations, devices, components, and/or combinations thereof.
1-12, a long-endurance unmanned aerial vehicle comprises a fuselage 1, a left wing middle section 2, a right wing middle section 3, a left wing tail section 4, a right wing tail section 5 and a tail wing 6;
the two sides of the fuselage 1 are respectively connected with the left wing middle section 2 and the right wing middle section 3 through an overhanging carbon tube component and a first lock catch.
The machine body 1 is connected with the tail wing 6 through a second lock catch and a positioning pin 15, the lock catches are locked at two sides of the tail of the machine body 1, the positioning pin 15 is arranged on the end face of the tail of the machine body 1, the positioning pins 15 are symmetrically distributed along the connector, and the tail wing 6 is locked after being installed in place, so that the installation of the tail wing 6 is completed;
the left wing middle section 2 is connected with the left wing tail section 4 through a first carbon tube 41 and a third lock catch 42; the right wing middle section 3 is connected with the right wing tail section 5 through a second carbon tube and a fourth lock catch; the connection mode between the left wing middle section 2 and the left wing tail section 4 and the connection mode between the right wing middle section 3 and the right wing tail section 5 are the same as the connection mode between the two sides of the fuselage 1 and the left wing middle section 2 and the right wing middle section 3.
The belly of fuselage 1 is provided with holds chamber 11, holds chamber 11 and is used for installing load cabin 7, holds chamber 11 and load cabin 7 through quick detach structure and fix. The belly middle section of the fuselage 1 is designed into an n-shaped structure for the installation of a load compartment 7, and the load compartment 7 can be installed with different loads. The load cabin 7 can be replaced by replacing the load cabin 7, so that different task demands of the unmanned aerial vehicle are met.
The quick detach structure includes: a U-shaped chute 111 and a U-shaped slide rail 71; the two ends of the accommodating cavity 11 along the length direction of the machine body 1 are respectively provided with a U-shaped sliding groove 111, and the two ends of the top of the load cabin 7 are respectively provided with a U-shaped sliding rail 71 matched with the U-shaped sliding groove 111; the length direction of the U-shaped sliding groove 111 is parallel to the length direction of the U-shaped sliding rail 71, and the U-shaped sliding groove 111 and the U-shaped sliding rail 71 are mutually limited along the direction of the machine body 1 so that the load compartment 7 is detachably connected with the machine body 1.
The opening 711 between the mutually matched U-shaped sliding groove 111 and the U-shaped sliding rail 71 is arranged opposite to each other.
The U-shaped chute 111 and the U-shaped slide rail 71 are respectively fixed to the machine body 1 and the load compartment 7 by screws.
The top of the load cabin 7 is scratched into the U-shaped chute 111 of the machine body 1 through two U-shaped sliding rails 71, and the load cabin can be automatically locked after being installed in place. The body and the bottom of the load compartment 7 are used for mounting loads, and the mounting bracket structure inside the body of the load compartment 7 can be adjusted according to different task requirements, so that the load compartment 7 is used for mounting different loads (a nacelle, a radar and the like). The load cabin 7 is arranged in the middle of the fuselage 1, even if the influence of changing different loads on the mass center and the rotational inertia of the whole aircraft is small, the different load cabins 7 can be changed by adjusting the structure of the load cabin 7, so that the mounting of different loads can be realized, and the fuselage 1 and other structures are not changed.
The overhanging carbon tube assembly comprises a first overhanging carbon tube 12, a first inner carbon tube 13 and a second overhanging carbon tube 22, wherein the first overhanging carbon tube 12 is arranged at two sides of the fuselage 1 and symmetrically extends outwards, the first inner carbon tube 13 is arranged in the fuselage 1, the second overhanging carbon tube 22 is arranged at one end of the left wing middle section 2 and the right wing middle section 3, which faces the fuselage 1, and the second inner carbon tube is respectively arranged at one end of the left wing middle section 2 and the right wing middle section 3, which faces the fuselage 1; the first carbon tubes 12 are inserted into the second inner carbon tubes, and the second carbon tubes 22 are inserted into the first inner carbon tubes 13. The number of the carbon tubes and the lock catches is not limited to that shown in the figure, and can be properly adjusted according to actual needs.
The first lock catch, the second lock catch, the third lock catch and the fourth lock catch all comprise: the lock catch connecting piece 21 and the lock catch connecting matching piece matched with the lock catch connecting piece 21 are arranged on the left wing middle section 2, the right wing middle section 3, the left wing tail section 4, the right wing tail section 5 and the tail wing 6, and the lock catch connecting matching piece is arranged on the fuselage 1, the left wing middle section 2 and the right wing middle section 3.
The left wing middle section 2 is provided with a lock catch connecting piece 21 and the right wing middle section 3 is provided with a lock catch connecting piece 21, the lock catch connecting piece 21 is arranged at one end facing the fuselage 1, and two lock catch connecting matching pieces which are respectively matched with the lock catch connecting pieces 21 of the right wing middle section 3 and the left wing middle section 2 and are arranged on the fuselage 1 are arranged at two sides of the fuselage 1; the latch connector 21 includes: a stop collar 211, a fixed arm 212, and a torsion spring; the fixed arm 212 is connected with the limit ring 211 through a torsion spring; the middle part of the torsion spring is provided with a rotating shaft which is fixedly connected with the left wing middle section, the right wing middle section, the left wing tail section 4, the right wing tail section 5 and the tail wing 6 respectively; specifically, two ends of the rotating shaft are respectively fixed on 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 wing 6; the lock catch connecting fitting piece is a limiting hook 14; the limit hook 14 is in limit fit with the limit ring 211; the fixed arm 212 rotates along the first direction to drive the limit ring 211 to move along the direction away from the limit hook 14, so that the limit ring 211 and the limit hook 14 are more firmly limited, and the torsion spring is in a normal state at the moment; the fixed arm 212 rotates along the second direction to drive the limit ring 211 to move along the direction approaching the limit hook 14, so that the limit ring 211 and the limit hook 14 are not limited.
When the fuselage 1 is connected with the left wing middle section 2 and the right wing middle section 3, the limiting hooks 14 and the limiting rings 211 are in limiting fit with each other along the direction perpendicular to the fuselage 1; when the machine body 1 is connected with the tail wing 6, the limit hooks 14 and the limit rings 211 are in limit fit with each other along the direction parallel to the machine body 1.
Specifically, the tail of the machine body 1 is provided with a locating pin 15, the tail wing 6 is provided with a locating hole, and after the locating pin 15 is inserted into the locating hole, the machine body 1 is connected with the tail wing 6 through a lock catch.
The installation process comprises the following steps: the two sides of the fuselage 1 are respectively connected with the left wing middle section 2 and the right wing middle section 3 through an overhanging carbon tube component and a first lock catch, when the left wing middle section 2 and the right wing middle section 3 are installed, overhanging carbon tubes of the fuselage 1 penetrate into the left wing middle section 2 and the right wing middle section 3, and meanwhile, overhanging carbon tubes of the left wing middle section 2 and the right wing middle section 3 are inserted into rear carbon tubes which are flush with the fuselage 1; after the left wing middle section 2 and the right wing middle section 3 are installed in place, the lock catch is locked, and then the installation of the left wing middle section 2 and the right wing middle section 3 is completed;
the unmanned aerial vehicle whole machine supporting material is made of a high-strength carbon plate and a balsawood structure; the engine body skin is mainly made of high-strength carbon fiber composite materials, and part of the cabin cover is made of glass fiber composite materials. The two materials have the characteristics of high strength, low density and the like, and the weight of the aircraft is lighter on the premise of meeting the flight requirement of the aircraft. The unmanned aerial vehicle truss adopts the carbon fiber composite material board and the balsawood sandwich composite material board, makes the complete machine weight lighter under the condition that intensity satisfies the requirement.
The front section of the machine body 1 is provided with a battery compartment 9, and the rear section of the machine body 1 is provided with an avionic compartment which is respectively used for installing a battery and an avionic system 8; the avionics system 8 with large heating value in the avionics cabin is fixed at the bottom of the machine body 1; the front end and the rear end of the machine body 1 are respectively provided with an antenna mounting cover for mounting an antenna; two foot rests are designed in front of and behind the lower part of the machine body 1, and an antenna is embedded in the rear foot rest. The avionics system 8 has the characteristics of high integration level, light weight, small volume and the like, and the bottom mounting surface is simultaneously designed with the radiating fins, so that the avionics system 8 can radiate heat.
Specifically, the wing tip shape is a bending-up type and the tail wing type is a V type.
The invention also provides a pneumatic optimization method of the long-endurance unmanned aerial vehicle, which adopts the long-endurance unmanned aerial vehicle and comprises the following steps:
step S1: selecting model parameters: selecting one model from a general unmanned aerial vehicle selection library, and setting wing area S 1 Wing span b, wing average aerodynamic chord length Ca, wing root chord length Cr, wing tip chord length Ct, wing aspect ratio A;
step S2: carrying out fluid mechanics simulation on the turbulence of the long-endurance unmanned aerial vehicle by utilizing a Reynolds average N-S equation; the method for acquiring the Reynolds average N-S equation comprises the following steps:
decomposing each interpretation variable in the instantaneous N-S equation into a corresponding average valueA pulsation component phi':
wherein phi represents any one of a velocity component, pressure, energy, and substance concentration;
will be decomposed into flatMean value ofAnd the pulsation component phi' is inserted into an instantaneous N-S equation to obtain a mean equation; the mean value equation comprises an average mass equation and a momentum transfer equation; wherein,
the average mass equation is:
the momentum transfer equation is:
wherein,representing the outer product, ρ is the density, +.>And->Average velocity and average pressure, respectively, I is the identity tensor, T is the viscous stress tensor, f b Is the resultant of various volume forces acting on the unit volume of the continuum; the additional term is a Reynolds stress tensor T RANS It is defined as follows:
wherein V 'is' x 、V′ y 、V′ z Speed pulsation of the unmanned aerial vehicle along x, y and z axes is respectively carried out;
aircraft designs require high precision to simulate the viscous effects of flow fields to properly calculate aerodynamic characteristics of aircraft having different profiles. The N-S equation of the Reynolds average is calculated by adding the turbulence model, which is a main stream scheme of modern numerical simulation of complex viscous flow fields, and the calculation accuracy of the RANS equation after adding the turbulence model in the aspect of aerodynamic force or moment applied to an aircraft is improved.
For T RANS The solution difficulty of the formula is that T is calculated according to the average flow RANS Modeling is performed so that the control equation is closed. The Reynolds stress tensor T can be determined by turbulent eddy-current viscosity coefficient RANS Modeling was performed as a function of average flow, with the most common model being called the Boussinesq approximation (cloth Xin Niesi g approximation, kinetic term):
wherein S is the average strain rate tensor;
turbulent vortex viscosity coefficient mu t The calculation formula of (2) is as follows:
μ t ＝ρC μ f μ kT
wherein μ is hydrodynamic viscosity, k is turbulent pulsating kinetic energy, C μ =0.09 as a constant, f μ T is the turbulence time scale as a damping function;
the transport equation for turbulent pulsation kinetic energy k and turbulent dissipation rate ε is:
wherein sigma k 、σ ε 、C ε1 、C ε2 Is constant, f 2 As a damping function, P k 、P ε To generate conditions S k 、S ε To specify conditions ε 0 An ambient turbulence value to counteract turbulence attenuation;
s3: different reynolds stress tensors T RANS According to different aerodynamic formulas, a standard library is queried according to the design to obtain an approximate aerodynamic coefficient formula, and model parameters after simulation are substituted into the aerodynamic coefficient formula:
wherein: l (L) GB Is the lifting force, V GB S is a reference area, and the wing area or the maximum cross-sectional area of the fuselage is taken as the free flow speed;
wherein D is GB Is resistance;
wherein M is GB Is pitching moment; c A The average aerodynamic chord length of the wing;
setting the attack angle to be-5 degrees to 15 degrees according to formulas (3-1) - (3-3) to obtain a lift coefficient C L Coefficient of resistance C D Pitch moment coefficient C m Range.
Calculation of lift-to-drag ratioObtaining a maximum lift-drag ratio corresponding to an optimal attack angle according to formulas (3-1) - (3-3); the lift-drag ratio is more than or equal to 20, namely the unmanned aerial vehicle is considered to have better aerodynamic characteristics, and is suitable for long-endurance flight operation.
Since the acquired wind tunnel test data is aerodynamic coefficients, refer to GBT16638.4-2008 aerodynamic concepts, quantities and symbols-part 4: to avoid confusion with the symbols in the text, the symbols defined by national standards are marked with footmarks and written. Further, a aerodynamic coefficient formula was obtained as described in Table 1.
TABLE 1 aerodynamic coefficient formula table
The relevant aerodynamic force and moment coefficient graphs are calculated according to the formulas in table 1 and are shown in fig. 13-16.
Based on the full turbulence CFD method, the pneumatic characteristics of the whole unmanned aerial vehicle are obtained as follows:
aoa in the above table represents the angle of attack, C D Representing the drag coefficient, C L Representing the lift coefficient, C M Representing the pitch moment coefficient.
Based on the full turbulence CFD method, the aerodynamic characteristics of the unmanned aerial vehicle provided by the application are summarized as follows:
1) When the cruising lift coefficient is 0.7, the attack angle is about 2.5 degrees, and the resistance is small;
2) At an angle of attack of about 3 deg., the lift-drag ratio reaches 21.9.
The aerodynamic characteristics of the unmanned aerial vehicle obtained through CFD simulation can be known, and the aerodynamic characteristics of the unmanned aerial vehicle meet the design requirements.

Claims (8)

1. A long-endurance unmanned aerial vehicle, comprising: the device comprises 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;
two sides of the fuselage are respectively connected with the left wing middle section and the right wing middle section through an overhanging carbon tube assembly and a first lock catch;
the machine body is connected with the tail wing through a second lock catch and a locating pin;
the middle section of the left wing is connected with the tail section of the left wing through a first carbon tube and a third lock catch; the middle section of the right wing is connected with the tail section of the right wing through a second carbon tube and a fourth lock catch;
the belly of fuselage is provided with holds the chamber, hold the chamber and be used for installing the load cabin.
2. The long-endurance unmanned aerial vehicle of claim 1, wherein the two ends of the accommodating cavity along the length direction of the fuselage are respectively provided with a U-shaped chute, and the two ends of the top of the load cabin are respectively provided with a U-shaped slide rail matched with the U-shaped chute; the length direction of the U-shaped sliding groove is parallel to the length direction of the U-shaped sliding rail, and the U-shaped sliding groove and the U-shaped sliding rail are mutually limited along the direction of the machine body so that the load cabin is detachably connected with the machine body; the opening between the mutually adapted U-shaped sliding groove and the U-shaped sliding rail is arranged oppositely.
3. The long-endurance unmanned aerial vehicle of claim 1, wherein the overhanging carbon tube assembly comprises a first overhanging carbon tube, a first internal carbon tube, a second overhanging carbon tube and a second internal carbon tube, wherein the first overhanging carbon tube is arranged on two sides of the fuselage and symmetrically extends outwards; the first overhanging carbon tube is inserted into the second inner carbon tube, and the second overhanging carbon tube is inserted into the first inner carbon tube.
4. The long-endurance unmanned aerial vehicle of claim 1, wherein the first latch, the second latch, the third latch, and the fourth latch each comprise: the device comprises a lock catch connecting piece and a lock catch connecting matching piece matched with the lock catch connecting piece, wherein the lock catch connecting piece is arranged on a left wing middle section, a right wing middle section, a left wing tail section, a right wing tail section and a tail wing, and the lock catch connecting matching piece is arranged on the fuselage, the left wing middle section and the right wing middle section.
5. The long-endurance unmanned aerial vehicle of claim 4, wherein the latch connector provided on the middle section of the left wing and the latch connector provided on the middle section of the right wing are provided at one end facing the fuselage, and two latch connector fittings provided on the fuselage and respectively engaged with the latch connectors of the middle section of the right wing and the middle section of the left wing are provided at both sides of the fuselage.
6. The long-endurance unmanned aerial vehicle of claim 4, wherein the latching connection comprises: limit ring, fixed arm and torsion spring; the fixed arm is connected with the limit ring through a torsion spring; the middle part of the torsion spring is provided with a rotating shaft which is fixedly connected with the left wing middle section, the right wing middle section, the left wing tail section, the right wing tail section and the tail wing respectively; the lock catch connecting fitting is a limiting hook; the limit hook is in limit fit with the limit ring.
7. The long-endurance unmanned aerial vehicle of claim 1, wherein the front section of the fuselage is provided with a battery compartment, the rear section of the fuselage is provided with an avionics compartment, and the battery compartment and the avionics compartment are respectively used for installing a battery and an avionics system; the front end and the rear end of the machine body are respectively provided with an antenna mounting cover for mounting an antenna; front foot frames and rear foot frames are arranged in front of and behind the lower portion of the machine body, and antennas are buried in the rear foot frames.
8. A method for the aerodynamic optimization of a long-endurance unmanned aerial vehicle, characterized in that a long-endurance unmanned aerial vehicle as claimed in any one of the preceding claims 1 to 7 is used, comprising the following steps:
step S1: selecting model parameters: selecting one model from a general unmanned aerial vehicle selection library, and setting wing area S 1 Wing span b, wing average aerodynamic chord length Ca, wing root chord length Cr, wing tip chord length Ct, wing aspect ratio A;
step S2: and carrying out fluid mechanics simulation on the turbulence of the long-endurance unmanned aerial vehicle by utilizing a Reynolds average N-S equation, wherein the acquisition method of the Reynolds average N-S equation comprises the following steps:
decomposing each interpretation variable in the instantaneous N-S equation into a corresponding average valueA pulsation component phi':
wherein phi represents any one of a velocity component, pressure, energy, and substance concentration;
mean value after decompositionAnd the pulsation component phi' is inserted into an instantaneous N-S equation to obtain a mean equation; the mean value equation comprises an average mass equation and a momentum transfer equation; wherein,
the average mass equation is:
the momentum transfer equation is:
wherein,representing the outer product, ρ is the density, +.>And->Average velocity and average pressure, respectively, I is the identity tensor, T is the viscous stress tensor, f b Is the resultant of various volume forces acting on the unit volume of the continuum; additional item T RANS Is a reynolds stress tensor defined as follows:
wherein V 'is' x 、V′ y 、V′ z Speed pulsation of the unmanned aerial vehicle along x, y and z axes is respectively carried out;
reynolds stress tensor T by turbulent vortex viscosity coefficient RANS Modeling as a function of average flow:
wherein S is the average strain rate tensor;
turbulent vortex viscosity coefficient mu t The calculation formula of (2) is as follows:
μ t ＝ρC μ f μ kT
wherein μ is hydrodynamic viscosity, k is turbulent pulsating kinetic energy, C μ =0.09 as a constant, f μ T is the turbulence time scale as a damping function;
the transport equation for turbulent pulsation kinetic energy k and turbulent dissipation rate ε is:
wherein sigma k 、σ ε 、C ε1 、C ε2 Is constant, f 2 As a damping function, P k 、P ε To generate conditions S k 、S ε To specify conditions ε 0 An ambient turbulence value to counteract turbulence attenuation;
s3: different reynolds stress tensors T RANS According to different aerodynamic formulas, a standard library is queried according to the design to obtain an approximate aerodynamic coefficient formula, and model parameters after simulation are substituted into the aerodynamic coefficient formula:
wherein: l (L) GB Is the lifting force, V GB S is a reference area, and the wing area or the maximum cross-sectional area of the fuselage is taken as the free flow speed;
wherein D is GB Is resistance;
wherein M is GB Is pitching moment; c A The average aerodynamic chord length of the wing;
setting the attack angle to be-5 degrees to 15 degrees according to formulas (3-1) - (3-3) to obtain a lift coefficient C L Coefficient of resistance C D Pitch moment coefficient C m A range;
calculation of lift-to-drag ratioObtaining a maximum lift-drag ratio corresponding to an optimal attack angle according to formulas (3-1) - (3-3); the lift-drag ratio is more than or equal to 20, namely the unmanned aerial vehicle is considered to have better aerodynamic characteristics, and is suitable for long-endurance flight operation.
CN202311838994.XA 2023-12-28 2023-12-28 Long-endurance unmanned aerial vehicle and pneumatic optimization method thereof Pending CN117622552A (en)

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Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN117885923A (en) * 2024-03-14 2024-04-16 航大汉来（天津）航空技术有限公司 Long-endurance patrol unmanned aerial vehicle

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN117885923A (en) * 2024-03-14 2024-04-16 航大汉来（天津）航空技术有限公司 Long-endurance patrol unmanned aerial vehicle
CN117885923B (en) * 2024-03-14 2024-06-14 航大汉来（天津）航空技术有限公司 Long-endurance patrol unmanned aerial vehicle

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