CN115408771B - Design method of high-altitude ultra-long endurance large aspect ratio integrated unmanned aerial vehicle platform - Google Patents

Abstract

The present invention proposes a design method for a high-altitude, ultra-long-flight, large-aspect-ratio integrated unmanned aerial platform. The full-platform design of the rectenna is considered from the perspective of overall integration, and the transferability and relevance of various overall, aerodynamic, structural and other parameters are fully considered. When entering the detailed design stage, the sub-component design under the constraints of the integrated platform can be adopted to speed up the iterative process. In the present invention, the constraints and design indicators are decomposed through an integrated approach, and interactive design is performed on the platform and the antenna, and finally modular research and development under integration is realized. The rectenna is combined with the platform in the form of an integrated module, that is, the antenna and the platform can be completely combined into an integrated system through a simple connection relationship.

Description

A design method for a high-altitude, ultra-long-flight, large-aspect-ratio integrated unmanned aerial platform

Technical Field

The present invention relates to the field of unmanned aerial platform design, and in particular to a design method for a high-altitude, ultra-long-flight, large-aspect-ratio integrated unmanned aerial platform.

Background technique

Affected by the energy crisis and environmental protection situation, one of the most important power sources for high-altitude, long-endurance flight platforms is solar energy. Many high-altitude, long-endurance unmanned aerial vehicles using solar energy as power have been born, mainly including stratospheric solar airships and drones.

1) Stratospheric Solar Airship

Solar-powered airships use solar panels laid on the surface of the hull, and use the converted energy to drive the propeller for flight and control. Due to the huge hull, solar panels can only be laid on the upper part of the hull, and material limitations make it difficult to choose the laying position. The actual available area is not large, and the huge hull equipment and power electricity requirements must be met, which is technically very difficult. At present, only the solar-powered airship developed by the Canadian Solar Airship Company and the CA-21R airship developed by China's Datian Airship Company have achieved trial flights. Relatively speaking, the practicality of solar-powered airships is lower than that of drones.

2) High-altitude, long-flight solar-powered drones

UAVs powered by solar energy maintain power and store electrical energy through solar panels on the wings under lighting conditions. They have high flight altitudes and long endurance, and are currently a hot topic of research in various countries. Although solar-powered UAVs have obvious shortcomings in flight performance and payload capacity compared to conventional powered UAVs, they have obvious advantages in long-flight endurance and can theoretically achieve wireless time cruising. At present, the more famous long-flight UAVs include the "Solar" series of UAVs jointly developed by the American AeroVironment Company and NASA; and the "Zephyr" series of UAVs developed by a company under the British Ministry of Defense. In addition, the AtlantikSolar solar-powered UAV developed by the Swiss Federal Institute of Technology in Zurich set a continuous flight record of 81 hours in 2015, and the ApusDuo autonomous solar-powered UAV designed by UAVOS aims to achieve the goal of 365-day cruising. Domestic research institutes and universities have also carried out a lot of research on solar aircraft design. In July 2019, Professor Zhou Zhou of Northwestern Polytechnical University's "Phantom" UAV MY-12 achieved cross-day and night flight, reaching 27 hours and 37 minutes, which is the longest-lasting solar-powered UAV publicly reported in China.

Obviously, ultra-long-flight UAVs powered by solar energy are currently one of the hot spots for development. However, despite the "unlimited flight time" in theory, due to the influence of weather, latitude, and day-night contrast, solar aircraft need to use solar energy to charge during the day, lower the altitude at night to use batteries to maintain powered flight, and improve flight efficiency. The maximum conversion efficiency of solar panels receiving solar energy does not exceed 30%. When the impact on charging is too great, it is impossible to maintain sustained flight. In addition, since batteries are needed to maintain flight when sunlight is weak, the weight of the battery accounts for almost 40% of the total weight of the aircraft, which seriously affects the carrying capacity of other loads. Therefore, the body structure of solar UAVs is very light. In order to lay as many panels as possible, the aspect ratio is usually large, which brings severe aeroelastic and control problems. It can be said that although unlimited solar energy can be used to achieve flight, poor load capacity, high control difficulty, large impact of environmental factors, and low energy conversion efficiency are the key issues restricting the development of this type of aircraft.

Therefore, the design of ultra-long-flight UAVs driven by electricity needs to be comprehensively considered from the aspects of energy source and utilization, load capacity, aerodynamic structure efficiency and control. The best way should be to achieve efficient flight with unlimited flight time without reducing load capacity, that is, to combine the advantages of conventional powered long-flight aircraft with a stable and unlimited energy source to achieve efficient conversion of radiant energy. However, the current solar aircraft solution obviously cannot meet the requirements.

In order to achieve ultra-long flight time and stable endurance of UAVs, microwaves are used as the energy source of the UAV power device, which has become a new way of powering unmanned platforms. There are relatively few unmanned platforms with microwave energy functions now. Except for some early airships and UAVs, there are not many examples. For example, in 1980, Canada conducted a study called SHARP (Stationary High Altitude Relay Program), in which a 4.1kg UAV was designed. The UAV flew continuously for 20 minutes under the drive of a 10kw ground transmitting antenna, verifying the feasibility of this energy driving method. However, the main problem with these platforms is that because the antenna receiving microwave energy requires the receiving surface to be circular and flat, the current platform solutions are to directly install this receiving antenna in the form of a component on the body of the unmanned platform, and the antenna and the UAV are not integrated. Although this method is simple, it brings various problems in aerodynamics, structure and energy reception, and the performance of the flight platform is extremely poor, and the efficiency of the entire system is low.

Pneumatics:

The receiving antenna is usually mounted on the belly of the unmanned platform in a hoisting manner, which causes great damage to the overall aerodynamic characteristics, mainly manifested in a sharp increase in resistance. The handling and stability characteristics of the entire aircraft will also undergo significant changes. This is a great interference to high-altitude and long-endurance UAVs that use electricity to drive propellers, and seriously deteriorates the final platform performance.

Structural aspects:

Due to the presence of the rectifier circuit board, the receiving antenna has a relatively high density, so the total weight is relatively large, especially on large unmanned aerial platforms with high energy requirements. The weight brought by the large-area rectifier structure antenna is likely to offset the benefits of this form of energy supply. In addition, due to the large size of the antenna, the air bomb and vibration effects during flight are obvious.

Energy reception:

After the rectifier circuit is hoisted separately on the unmanned platform, it is loaded during flight and is prone to structural deformation, which causes displacement and phase deviation of the rectifier receiving unit, affecting the beam collection efficiency.

Summary of the invention

In order to solve the problems existing in the prior art, the present invention proposes a design method for a high-altitude, ultra-long-endurance, large-aspect-ratio integrated unmanned aerial platform. The scheme adopts a large-aspect-ratio wing-body fusion layout, and an integrated microwave energy receiving rectifier antenna is designed under the fuselage. It can receive and convert the radio frequency energy radiated from the ground into direct current, and drive the electric propeller through direct current for cruise flight. The design method is driven by the performance of the rectifier antenna load, and the aerodynamic and structural fusion design of the flight platform is carried out around the load. The layer-by-layer design and step-by-step optimization method of "aerodynamic design → aerodynamic optimization → structural/load electrical characteristics design → -structural/load electrical characteristics optimization → multidisciplinary collaborative optimization of aerodynamic/structural/load electrical characteristics → optimal configuration of aerodynamic/structural/load electrical characteristics" is adopted to improve the design efficiency.

The technical solution of the present invention is:

A design method for a high-altitude, ultra-long-flight, large-aspect-ratio integrated unmanned aerial platform comprises the following steps:

Step 1: According to the design requirements, the basic characteristics of the unmanned aerial platform and the antenna load are determined; the energy source of the unmanned aerial platform is the microwave radio frequency energy radiated by the ground transmitting end, which is transmitted through space to reach the receiving antenna arranged on the unmanned aerial platform, and then further converted into direct current;

Step 2: According to the antenna load characteristics determined in step 1, determine the constraints and requirements of the antenna load on the unmanned aerial platform;

Step 3: Determine the basic layout of the unmanned aerial platform based on the constraints and requirements of the antenna load on the unmanned aerial platform determined in step 2, as well as the design performance requirements;

Step 4: Determine the airfoil of the inner and outer wing sections based on the constraints and requirements of the antenna load on the unmanned aerial platform:

Step 5: Determine the shape and parameters of the unmanned aerial platform by integrating the basic characteristics of the antenna payload and the unmanned aerial platform, the constraints and requirements of the antenna payload on the unmanned aerial platform, and the basic layout of the unmanned aerial platform:

Based on the wing loading parameters of a typical high aspect ratio aircraft As well as the platform constraints and requirements in step 2 and the basic platform layout in step 3, determine the wing area S, and then obtain the span b according to the initially determined aspect ratio;

Determine the engine thrust T and the corresponding power requirement P based on the takeoff weight m and the estimated cruise lift-to-drag ratio (L/D) _c ;

According to the antenna energy density constraint, the antenna diameter _be is obtained;

According to the static stability margin requirements and the empirical data of the tail capacity of typical V-tail aircraft, the chordwise position of the V-tail and the size of the V-tail are determined by design;

Determine the initial position and size of ailerons and V-tail control surfaces based on typical V-tail aircraft handling characteristics data;

Determine the propeller diameter, number of blades, and blade shape based on the estimated cruise lift-to-drag ratio and power requirements;

Based on the above parameters, the aerodynamic shape of a high aspect ratio flight platform integrating an integrated receiving rectenna is constructed;

Step 6: Simulate and calculate the aerodynamic characteristics of the high aspect ratio flight platform constructed in step 5 to determine the initial aerodynamic performance of the flight platform, including the aerodynamic performance analysis of the take-off and landing configuration and the aerodynamic performance analysis of the cruise configuration. The aerodynamic performance analysis of the take-off and landing configuration includes determining the lift coefficient, drag coefficient and pitch moment coefficient curves _CL ~α, _CD ~α, _Cm ~α from 0 degrees to the stall angle of attack of the take-off and landing configuration; the aerodynamic performance analysis of the cruise configuration includes determining the lift coefficient, drag coefficient and pitch moment coefficient curves _CLp ~α, _CDp ~α, _Cmp ~α from 0 degrees to the stall angle of attack of the cruise configuration, the cruise trim lift-to-drag ratio (L/D) _c and the cruise static stability margin;

Step 7: Based on the initial flight platform aerodynamic performance simulation results obtained in step 6, the initial flight platform multi-objective aerodynamic optimization is carried out with the optimization objectives of improving the cruise lift-to-drag ratio (L/D) _c and reducing the longitudinal nose-down moment C _m of the entire aircraft, and a single aerodynamic optimization design is performed. Based on the aerodynamic shape optimization results, the aerodynamic load distribution is determined;

Step 8: Based on the aerodynamic load distribution obtained in step 7, design the internal structure of the rectenna and the structure of the flight platform, and complete the preliminary structural finite element characteristic analysis;

Step 9: Based on the preliminary structural finite element characteristic analysis in step 7, determine the embedding and fusion method of the rectenna distributed antenna structure, and simulate and analyze the electrical characteristics of the design scheme;

Step 10: Based on the electrical performance characteristics obtained in step 9, optimize the structural/electrical integration design of the integrated antenna platform;

Step 11: Based on the platform shape layout plan, structural finite element plan and rectenna electrical characteristics plan obtained in steps 7-10, conduct multidisciplinary collaborative optimization design of aerodynamic/structural/electrical radiation characteristics to complete the design of a high-altitude, ultra-long-flight, large-aspect-ratio integrated unmanned flying platform.

Furthermore, in step 1,

The basic characteristics of the antenna payload include: the geometric shape of the antenna payload, the receiving area, the antenna diameter, the total thickness, the energy conversion efficiency and the energy density;

The basic characteristics of the unmanned aerial platform include: aspect ratio, cruise lift-to-drag ratio, cruise lift coefficient, take-off gross weight, wing area and static stability margin requirements.

Furthermore, in step 2, the constraints and requirements of the antenna payload on the platform include:

1) According to the receiving area of the antenna load, determine the area of the antenna load layout on the unmanned aerial platform, requiring the antenna load to have sufficient layout area on the unmanned aerial platform;

2) The antenna load layout area is a flat area, and the maximum structural deformation is not greater than the set requirements;

3) The surface skin thickness of the antenna load arrangement area meets the total thickness requirements of the antenna load.

Furthermore, in step 3, the basic layout of the unmanned aerial platform includes:

1) The unmanned aerial platform adopts a high aspect ratio wing-body fusion layout and is driven by a tail-mounted electric propeller;

2) According to the static stability margin requirements, the unmanned aerial platform adopts an inverted V-tail form to avoid the influence of propeller slipstream on the efficiency of the V-tail control surface;

3) According to the deformation constraint of antenna structure, antenna fusion design is performed under the fuselage of the flight platform;

4) Determine the area of the outer wing and inner wing-body combination according to the aspect ratio and lift-to-drag ratio requirements;

5) Use winglets;

6) Adopt the conventional three-point take-off and landing method;

7) Ailerons are located on the outer wing sections and control roll; pitch and yaw are controlled by the V-tail control surfaces.

Furthermore, in step 4, the outer wing section adopts a high lift-to-drag ratio airfoil with a straight lower surface, and the lift-to-drag ratio of the airfoil is not less than the set requirement; the airfoil of the inner wing-body fusion section is further optimized on the basis of the high lift-to-drag ratio airfoil to obtain:

Design status: Ma = 0.1356, Reynolds number Re = 0.34E6, angle of attack α = 0°

Design goal: maximum lift-to-drag ratio

Constraints: Moment coefficient Cm ≥ -0.13

The maximum relative thickness T/C is greater than or equal to 15%

The lower surface of the inner wing-body fusion section in the constrained antenna area is a straight section: the area is restricted to y(x)=0, x∈(x1,x2), (x1,x2) are the coordinates of the lower surface of the inner wing-body fusion section airfoil corresponding to the antenna area.

Furthermore, in step 7, the design goal is:

The cruise lift-to-drag ratio is the largest (L/D) _c , and the longitudinal nose-down moment of the entire aircraft is the smallest

min 1/(L/D) _c

min|C _m |

The design constraints are:

The cruise lift coefficient is not less than the initial value C _Lcinitial :

C _Lc ≥C _Lcinitial

The antenna area does not change:

Area _e = Area _einitial

The aspect ratio of the whole machine remains unchanged:

Furthermore, in step 8, the internal structure of the rectenna and the structure of the unmanned aerial platform are designed, and a preliminary structural finite element characteristic analysis is completed, including:

According to the wing aerodynamic load and aeroelastic requirements, it is determined to adopt a double-beam structure that takes into account both load-bearing and deformation efficiency, and the preliminary front and rear beam positions, relative thickness and height distribution are given;

According to the aerodynamic load constraints of the rectenna area, the antenna structure is determined to be a multi-layer composite material, the total thickness δ _e is consistent with the wing skin thickness, and the multi-layer structure includes a packaging surface composite material and a honeycomb layer;

According to the connection method between the antenna area and the platform, the periphery of the antenna area of the platform is structurally reinforced, and the antenna structure and the platform are connected by bonding;

According to the bending moment bearing requirements of the whole aircraft, box beams are arranged transversely inside the fuselage;

According to the appearance design, the V-tail tail strut is directly inserted into the fuselage with a carbon tube;

According to the load-bearing requirements, multiple frames are arranged longitudinally in the wings and fuselage to connect the tail strut, landing gear, engine and skin;

According to the overall stress condition of the unmanned aerial platform, a finite element analysis is conducted on the structural characteristics of the entire aircraft to determine the maximum deformation of the entire aircraft and the stress concentration points.

Furthermore, in step 9, the embedding and integration method of the rectenna distributed antenna structure is determined, and the electrical characteristics of the design scheme are simulated and analyzed, including:

According to the structural design and stress conditions of the unmanned aerial platform, the integrated rectenna is embedded in the multi-layer composite material structure of the antenna area to form a foam sandwich. The sandwich structure of the antenna part is: packaging surface composite material, upper honeycomb, antenna layer, lower honeycomb, composite material.

According to the determined total thickness constraint δ _e of the platform skin, the electric radiation performance parameters of the integrated receiving rectenna are determined, including the directivity pattern, peak gain, and coupling degree; the electric performance is simulated to obtain the electric radiation performance characteristics.

Furthermore, in step 10, the structural/electrical integration design optimization process of the integrated antenna platform is as follows:

The objective function is the radiation performance f(x) and the overall structural characteristics g(x), that is, the optimal mechanical performance and the optimal electrical radiation performance are taken as the optimization targets; a is taken as the skin thickness, b is the foam thickness, δ _wf is the front beam thickness, δ _wb is the rear beam thickness, δ _r1 ,…δ _rn is the thickness of the longitudinally arranged wing ribs, and the mechanical and electrical integration design mathematical model of the integrated array is established as follows:

The design goals are: the lightest overall structural weight g(x) and the best electrical radiation performance f(x)

min 1/f(x)

min g(x)

The design constraints are: Skin thickness limit range a _min ≤ a ≤ _{a max}

The foam thickness is limited to b _min ≤ b ≤ b _max

Front beam thickness limit range δ _wfmin ≤δ _wf ≤δ _wfmax

Rear beam thickness limit range δ _wbmin ≤δ _wb ≤δ _wbmax

Rib thickness limit

Limit range of wing tip deformation: e _wt ≤ e _wtmax

The optimization algorithm is used to optimize the structural/electrical integration design of the fusion integrated antenna platform to obtain the layout plan with the best structural performance and electrical radiation performance.

Furthermore, in step 11, the process of multidisciplinary collaborative optimization design of aerodynamic/structural/electric radiation characteristics is as follows: given design variables, including:

Aerodynamics: wingspan, leading edge sweep angle, trailing edge sweep angle, dihedral installation angle, characteristic section thickness, characteristic section curvature, characteristic section maximum thickness position, root-to-wing ratio, winglet camber angle;

Structure: front beam thickness, rear beam thickness, fuselage reinforcement beam position, reinforcement beam thickness, characteristic section wing rib thickness, antenna area foam thickness, packaging composite material thickness, wing tip deformation;

Electrical characteristics: antenna element size, number of antenna elements, and antenna element arrangement;

The design goals are:

Aerodynamics: The cruise lift-to-drag ratio of the whole aircraft is not less than the optimized value in step 7 (L/C) _c ≥ (L/C) _c-firstopt

Structural aspect: The weight of the whole machine structure does not exceed the optimization value in step 10: g(x)≤g(x) _firstopt

Electrical performance: The radiation performance of the whole machine is not less than the optimized value in step 10: f(x)≥f(x) _firstopt

The design constraints are:

Aerodynamic shape constraints:

The cruise lift coefficient is not less than the optimized value in step 7: C _Lc ≥ _{C Lcfirstopt}

The antenna area does not change: Area _e = Area _einitial

The aspect ratio of the whole machine remains unchanged:

Structural constraints:

Skin thickness limit range a _min ≤ a ≤ a _max

The foam thickness is limited to b _min ≤ b ≤ b _max

Front beam thickness limit range δ _wfmin ≤δ _wf ≤δ _wfmax

Rear beam thickness limit range δ _wbmin ≤δ _wb ≤δ _wbmax

Rib thickness limit

Wing tip deformation limit range e _wt ≤ e _wtmax

Electrical characteristics constraints:

Array element effective area limit range: S _es ≤ S _esmax

Array element quantity limit range: _Nes ≤ _Nesmax

An optimization algorithm is used to conduct multidisciplinary collaborative optimization design of aerodynamic/structural/electric radiation characteristics, and a solution for a high-altitude, ultra-long-flight, large-aspect-ratio integrated unmanned flying platform is obtained.

Beneficial Effects

The design method of the high-altitude, ultra-long-flight, large-aspect-ratio integrated unmanned aerial platform proposed in the present invention considers the full-platform design of the rectenna from the perspective of overall integration, fully considering the transferability and relevance of various overall, aerodynamic, structural and other parameters, and entering the detailed design stage, that is, the sub-component design under the constraints of the integrated platform can be adopted to speed up the iterative process. In the present invention, the constraints and design indicators are decomposed through an integrated approach, and interactive design is performed on the platform and the antenna, and finally modular research and development under integration is realized. The rectenna is combined with the platform in the form of an integrated module, that is, the antenna and the platform can be completely combined into an integrated system through a simple connection relationship.

The high aspect ratio integrated unmanned aerial platform finally designed adopts a high aspect ratio wing-body fusion layout, and adds winglets at the wingtips to further reduce induced drag. As the source of the whole aircraft power unit, the rectifier antenna is integrated with the aerodynamic and structural design of the whole aircraft, which ensures the aerodynamic characteristics of the whole aircraft from the appearance. Since the rectifier antenna adopts a composite design form such as skin and foam similar to the surface structure of the fuselage, the two can achieve a high degree of structural unity and integration, and have similar mechanical properties. Therefore, its load-bearing characteristics can be considered as a whole, effectively avoiding the disadvantages of poor structural weight characteristics after the external rectifier antenna is installed.

Additional aspects and advantages of the present invention will be given in part in the following description and in part will be obvious from the following description, or will be learned through practice of the present invention.

BRIEF DESCRIPTION OF THE DRAWINGS

The above and/or additional aspects and advantages of the present invention will become apparent and easily understood from the description of the embodiments in conjunction with the following drawings, in which:

Figure 1: Global view of the large aspect ratio integrated unmanned aerial platform;

Figure 2: Three views of the large aspect ratio integrated unmanned aerial platform;

Figure 3: Location view and antenna view of the receiving antenna of the large aspect ratio integrated unmanned aerial platform;

Figure 4: Schematic diagram of the structure of the large aspect ratio integrated unmanned aerial platform rectifying receiving antenna.

Among them: label 1 is the winglet, label 2 is the outer wing section, label 3 is the front landing gear, label 4 is the electric propeller propulsion system, label 5 is the tail strut, label 6 is the inverted V-tail, label 7 is the inner wing section (wing-body fusion section), label 8 is the aileron, label 9 is the V-tail rudder, label 10 is the integrated receiving rectifier antenna, and label 11 is the fusion connection structure of the antenna and the platform.

Detailed ways

The design of an ultra-long-flight UAV driven by electricity needs to be comprehensively considered from the aspects of energy source and utilization, load capacity, aerodynamic structure efficiency and control. The best way should be to achieve efficient flight with unlimited flight time without reducing the load capacity, that is, to combine the advantages of conventional powered long-flight aircraft with a stable and unlimited energy source to achieve efficient conversion of radiant energy. However, current solar aircraft cannot meet the requirements. Therefore, in order to achieve ultra-long flight and stable endurance of UAVs, using microwaves as the energy source of the UAV's power device has become a new way to supply energy to unmanned platforms. The antenna that receives microwave energy requires the receiving surface to be circular and flat, while the existing microwave energy function unmanned platforms all install this receiving antenna directly on the unmanned platform body in the form of a component without any fusion design, resulting in many problems in aerodynamics, structure and energy reception.

To this end, the present invention proposes a design method for a high-altitude, ultra-long-flight, large-aspect-ratio integrated unmanned aerial platform under the premise of ensuring energy receiving efficiency, which can highly integrate the rectifying receiving antenna and the unmanned platform, eliminate the aerodynamic, structural and energy receiving problems caused by the outdoor antenna, and improve the efficiency of the entire platform.

The present invention is described below in conjunction with specific embodiments:

The main function of the high-altitude, ultra-long-flight, high-aspect-ratio UAV required to be realized in this embodiment is to carry out fixed-point cruising at an altitude of more than 18km, to achieve continuous and uninterrupted monitoring and reconnaissance of a certain area, and to perform ground information monitoring and signal relay tasks during emergency rescue. To this end, the unmanned aerial platform adopts a large aspect ratio wing-body fusion layout, and an integrated microwave energy receiving rectifier antenna is designed under the fuselage, which can receive and convert the radio frequency energy radiated from the ground into direct current, and drive the electric propeller through direct current for cruising flight. The design features of the unmanned aerial platform are mainly reflected in: 1) The large aspect ratio wing-body fusion layout plus the winglet can ensure a good lift-to-drag ratio characteristic in a high-altitude low Reynolds number environment, thereby reducing flight power consumption and improving power efficiency; 2) The integrated receiving rectifier antenna layout form integrates the rectifier antenna with the body shape and structure into an integrated design, which can reduce or eliminate the influence of the antenna load on the body shape and structural characteristics.

The basic idea of the unmanned aerial platform design method is: first, according to the design requirements, the basic characteristics of the platform and antenna loads are decomposed to clarify the mutual interference relationship; according to the load requirements, the design boundary of the platform is determined to construct a preliminary layout form; from the two-dimensional wing profile to the three-dimensional configuration, the shape design of the flight platform is completed, and the first aerodynamic optimization is carried out first to obtain a better shape scheme from the aerodynamic level; on the basis of the limitation of the aerodynamic shape scheme, the internal structural characteristics and the electromechanical characteristics of the antenna are designed, and the comprehensive optimization design of the whole machine structure/electrical performance under the limitation of the aerodynamic shape is carried out to obtain the first round of electromechanical optimized shape; finally, on the basis of the existing optimized aerodynamic shape, structural scheme and antenna scheme, the gradient method is used to carry out multidisciplinary collaborative optimization design of the aerodynamic/structural/electrical characteristics of the whole platform to obtain the optimal scheme configuration.

Specific steps are as follows:

Step 1: Determine the basic characteristics of the platform and antenna payload based on the design requirements:

The energy source of the platform is the microwave radio frequency energy radiated by the ground transmitter, which reaches the receiving antenna after being transmitted through space and is further converted into direct current. Because the microwave forms a circular coverage area after transmission, the receiving end is also required to be a flat circular surface, so the basic characteristics of the receiving end antenna load include:

Geometric shape: circular; receiving area: _Se , related to the final power requirement of the entire platform; antenna diameter: _De ; total thickness: _he , related to the structural design scheme; energy conversion efficiency: η; energy density: _ρa , related to the electrical characteristics of the antenna.

The basic features of the platform include:

Aspect Ratio: Cruise lift-to-drag ratio: (L/D) _c ; Cruise lift coefficient: C _Lc ; Take-off gross weight: m; Wing area: S _w ; Static stability margin requirement, which is 10% in this embodiment.

Step 2: Determine the constraints and requirements of the antenna payload on the platform based on the antenna payload characteristics:

According to the antenna load characteristics obtained in step 1, the constraints on platform design are:

1) The antenna has sufficient layout area;

2) The antenna layout area is less affected by structural deformation and is a flat area, that is, the lower surface of the antenna has flatness requirements;

3) The surface skin thickness _hw of the antenna load arrangement area meets the requirements, that is, _hw = _he .

Step 3: Determine the basic layout of the flight platform based on antenna load constraints and requirements, as well as design performance requirements:

According to the constraints and requirements of the antenna payload on the platform in step 2 and the requirements of the basic characteristics of the platform in step 1, the basic layout of the platform is preliminarily determined, mainly including:

1) According to the requirements of high altitude and low Reynolds number, the platform adopts a large aspect ratio wing-body fusion layout and is driven by a tail-mounted electric propeller;

2) According to the static stability margin requirements, the flight platform adopts an inverted V-tail form to avoid the influence of propeller slipstream on the efficiency of the V-tail control surface;

3) Based on constraints such as antenna structure deformation, antenna fusion design is performed under the fuselage of the flight platform;

4) Under the influence of aspect ratio and lift-to-drag ratio, determine the area of the outer wing and inner wing-body combination;

5) To reduce induced drag and improve lift-to-drag ratio, winglets are used;

6) Use conventional take-off and landing methods;

7) Ailerons are located on the outer wing sections and control roll; pitch and yaw are controlled by the V-tail control surfaces.

Step 4: Determine the inner and outer wing sections’ airfoils based on the antenna load constraints and platform requirements:

In order to ensure the lift-to-drag characteristics of the whole aircraft and take into account the flatness constraint of the lower surface of the antenna, the outer wing section adopts a high lift-to-drag ratio airfoil with a straight lower surface, and the inner wing-body fusion section adopts an optimized airfoil with flatness constraints:

The inner wing-body fusion section airfoil is further optimized based on the high lift-to-drag ratio airfoil, with the main goal of improving the lift-to-drag ratio at a cruising altitude of 18km:

Design status: Ma = 0.1356, Reynolds number Re = 0.34E6, angle of attack α = 0°

Design goal: maximum lift-to-drag ratio

Constraints: Moment coefficient Cm ≥ -0.13

The maximum relative thickness T/C is greater than or equal to 15%

The lower surface of the inner wing-body fusion section in the constrained antenna area is a straight section, that is, the area is restricted to y(x) = 0, x∈(x1,x2), (x1,x2) are the coordinates of the lower surface of the inner wing-body fusion section airfoil corresponding to the antenna area;

The Hicks-Henne analytical function method is used to parameterize the airfoil, and the MQPSO optimization algorithm is used to optimize the maximum lift-to-drag ratio of the airfoil. The population size is 20 and the iteration number is 30. The optimized inner wing-to-body fusion section airfoil is obtained.

Step 5: Consider the characteristics of the load and platform to determine the basic platform shape and parameters:

Based on the wing loading parameters of a typical high aspect ratio aircraft As well as the platform constraints and requirements in step 2 and the basic platform layout in step 3, determine the wing area S, and then obtain the span b according to the initially determined aspect ratio;

Determine the engine thrust T and the corresponding power requirement P based on the takeoff weight m and the estimated cruise lift-to-drag ratio (L/D) _c ;

According to the antenna energy density constraint, the antenna diameter _be is obtained;

According to the static stability margin requirements and the empirical data of the tail capacity of typical V-tail aircraft, the chordwise position of the V-tail and the size of the V-tail are determined by design;

Determine the initial position and size of ailerons and V-tail control surfaces based on typical V-tail aircraft handling characteristics data;

Determine the propeller diameter, number of blades, and blade shape based on the estimated cruise lift-to-drag ratio and power requirements;

Based on the above parameters, the aerodynamic shape of a high aspect ratio flight platform integrating an integrated receiving rectenna is constructed.

Step 6: Simulate and calculate the aerodynamic characteristics of the high aspect ratio flight platform constructed in step 5 to determine the initial aerodynamic performance of the flight platform, including the aerodynamic performance analysis of the take-off and landing configuration and the aerodynamic performance analysis of the cruise configuration. The aerodynamic performance analysis of the take-off and landing configuration includes determining the lift coefficient, drag coefficient and pitch moment coefficient curves _CL ~α, _CD ~α, _Cm ~α from 0 degrees to the stall angle of attack of the take-off and landing configuration; the aerodynamic performance analysis of the cruise configuration includes determining the lift coefficient, drag coefficient and pitch moment coefficient curves _CLp ~α, _CDp ~α, _Cmp ~α from 0 degrees to the stall angle of attack of the cruise configuration, the cruise trim lift-to-drag ratio (L/D) _c and the cruise static stability margin;

Step 7: Based on the initial flight platform aerodynamic performance simulation results obtained in step 6, perform single aerodynamic optimization design:

According to the difference between the initial flight platform aerodynamic data and the design target, the multi-objective aerodynamic optimization of the initial flight platform was carried out with the optimization targets of improving the cruise lift-to-drag ratio (L/D) _c and reducing the longitudinal nose-down moment C _m of the whole aircraft.

The design goals are:

The cruise lift-to-drag ratio is the largest (L/D) _c , and the longitudinal nose-down moment of the entire aircraft is the smallest

min 1/(L/D) _c

min|C _m |

The design constraints are:

The cruise lift coefficient is not less than the initial value C _Lcinitial :

C _Lc ≥C _Lcinitial

The antenna area does not change:

Area _e = Area _einitial

The aspect ratio of the whole machine remains unchanged:

The parameterization of the scheme in the optimization design uses the FFD method, the mesh deformation uses the IDW method, the improved PSO particle swarm optimization algorithm is used for optimization, and the proxy model is used for optimization acceleration. The optimization design first uses Latin hypercube to generate no less than 500 initial sample points, and uses the mixed point addition method to add points no less than 300 times. The fixed design space is used in the early optimization, and the adaptive design space is used in the later stage to expand the optimization results.

According to the aerodynamic shape optimization results, the aerodynamic load distribution is determined.

Step 8: Based on the aerodynamic load distribution obtained in step 7, design the internal structure of the rectenna and the structure of the flight platform, and complete the preliminary structural finite element characteristic analysis, including:

According to the wing aerodynamic load and aeroelastic requirements, determine the double-beam structure that takes into account both load-bearing and deformation efficiency, and give the preliminary front and rear beam positions, relative thickness and height distribution;

According to the aerodynamic load constraints of the rectenna area, the antenna structure is determined to be a multi-layer composite material. The total thickness _δe is consistent with the wing skin thickness. The multi-layer structural forms are: package surface composite material, honeycomb layer, etc.

According to the connection method between the antenna area and the platform, the periphery of the antenna area of the platform is structurally reinforced, and the antenna structure and the platform are connected by bonding;

According to the bending moment bearing requirements of the whole aircraft, box beams are arranged transversely inside the fuselage;

According to the appearance design, the V-tail tail strut is directly inserted into the fuselage with a carbon tube;

According to the load-bearing requirements, multiple frames are arranged longitudinally inside the wings and fuselage to connect the tail strut, landing gear, engine and skin, etc.

According to the overall stress conditions, the FEM method is used to conduct finite element analysis on the structural characteristics of the entire machine to determine the maximum deformation of the entire machine and the stress concentration point.

Step 9: Based on the preliminary structural finite element characteristic analysis, determine the embedding and integration method of the rectenna distributed antenna structure, and simulate and analyze the electrical characteristics of the design scheme:

According to the structural design and stress conditions of the entire platform, the integrated rectenna is embedded in the multi-layer composite material structure of the antenna area to form a foam sandwich, that is, the sandwich structure of the antenna part is: packaging surface composite material, upper honeycomb, antenna layer, lower honeycomb, composite material;

According to the determined total thickness constraint δ _e of the platform skin, the electric radiation performance parameters of the integrated receiving rectenna are determined, including the directivity pattern, peak gain, and coupling degree; the electric performance is simulated to obtain the electric radiation performance characteristics;

Step 10: Based on the initial electrical performance characteristics, optimize the structural/electrical integration design of the integrated antenna platform:

The objective function is the radiation performance f(x) and the overall structural characteristics g(x), that is, the optimal mechanical performance (including rigidity) and the optimal electrical radiation performance (including directivity, peak gain, coupling, etc.) are optimized. Let a be the skin thickness, b be the foam thickness, δ _wf be the front beam thickness, δ _wb be the rear beam thickness, δ _r1 ,…δ _rn be the longitudinally arranged rib thickness, and the mechanical and electrical integration design mathematical model of the integrated array is established as follows:

The design goals are: the lightest overall structural weight g(x) and the best electrical radiation performance f(x)

min 1/f(x)

min g(x)

The design constraints are: Skin thickness limit range a _min ≤ a ≤ _{a max}

The foam thickness is limited to b _min ≤ b ≤ b _max

Front beam thickness limit range δ _wfmin ≤δ _wf ≤δ _wfmax

Rear beam thickness limit range δ _wbmin ≤δ _wb ≤δ _wbmax

Rib thickness limit

Limit range of wing tip deformation: e _wt ≤ e _wtmax ;

The optimization algorithm is used to optimize the structural/electrical integration design of the fusion integrated antenna platform to obtain the layout plan with the best structural performance and electrical radiation performance.

Step 11: Based on the platform shape layout scheme, structural finite element scheme and rectenna electrical characteristic scheme obtained in steps 7-10, conduct multidisciplinary collaborative optimization design of aerodynamic/structural/electrical radiation characteristics:

In step 7, the aerodynamic characteristics of the flight platform were optimized by the optimization algorithm, and the relatively optimal configuration from the aerodynamic aspect was obtained. Then, based on the aerodynamic shape, in step 10, the structure and electrical radiation performance characteristics of the fusion rectenna scheme were further optimized, and the optimal layout scheme for structural performance and electrical radiation performance was obtained. The optimization of the two steps respectively obtained the optimal schemes for the shape and internal electromechanical characteristics, and the next step is to carry out the multidisciplinary collaborative optimization design of the aerodynamic/structural/electrical characteristics of the integrated rectenna of the full flight platform based on the optimal internal and external configurations. Because the optimization schemes of steps 7 and 10 are already the relatively optimal initial values of the schemes, relatively reasonable initial values can be given in the multidisciplinary collaborative optimization design of this step.

According to the design requirements of the entire platform and the optimization design process of steps 7 and 10, the design variables are given, including:

Aerodynamics: wingspan, leading edge sweep angle, trailing edge sweep angle, dihedral installation angle, characteristic section thickness, characteristic section curvature, characteristic section maximum thickness position, root-to-wing ratio, winglet camber angle;

Structure: front beam thickness, rear beam thickness, fuselage reinforcement beam position, reinforcement beam thickness, characteristic section wing rib thickness, antenna area foam thickness, packaging composite material thickness, wing tip deformation;

Electrical characteristics: antenna element size, number of antenna elements, and antenna element arrangement

The design goals are:

Aerodynamics: The cruise lift-to-drag ratio of the whole aircraft is not less than the optimized value in step 7 (L/C) _c ≥ (L/C) _c-firstopt

Structural aspect: The weight of the whole machine structure does not exceed the optimization value in step 10: g(x)≤g(x) _firstopt

Electrical performance: The radiation performance of the whole machine is not less than the optimized value in step 10: f(x)≥f(x) _firstopt

The design constraints are:

Aerodynamic shape constraints:

The cruise lift coefficient is not less than the optimized value in step 7: C _Lc ≥ _{C Lcfirstopt}

The antenna area does not change: Area _e = Area _einitial

The aspect ratio of the whole machine remains unchanged:

Structural constraints:

Skin thickness limit range a _min ≤ a ≤ a _max

The foam thickness is limited to b _min ≤ b ≤ b _max

Front beam thickness limit range δ _wfmin ≤δ _wf ≤δ _wfmax

Rear beam thickness limit range δ _wbmin ≤δ _wb ≤δ _wbmax

Rib thickness limit

Wing tip deformation limit range e _wt ≤ e _wtmax

Electrical characteristics constraints:

Array element effective area limit range: S _es ≤ S _esmax

Array element quantity limit range: _Nes ≤ _Nesmax

So far, the multidisciplinary collaborative optimization design of the aerodynamic/structural/electrical radiation characteristics of the entire platform has been completed, and the final platform geometric characteristics are:

1) The wingspan of the aircraft is 30m and the aspect ratio is 20. The large aspect ratio can ensure the aerodynamic efficiency of low-speed and high-altitude flight. Of course, the aspect ratio is further optimized and modified according to the performance requirements. This parameter is related to the power supply of the integrated receiving antenna, so it is optimized and set according to the corresponding requirements.

2) The wing area is 45m ² , the rectenna area is 15.918m ² , and the area ratio is determined according to the energy demand;

3) The airfoil of the wing-body fusion section is a high lift-to-drag ratio airfoil whose lift-to-drag ratio meets the set requirements, with a maximum relative thickness of 12.0044%, and the maximum thickness position is located at 30% of the chord length. The lower surface is adapted to the antenna requirements and has a long straight section; the winglet airfoil is a conventional NACA0012 airfoil, which is mainly used to reduce induced drag and improve the aerodynamic efficiency of the whole aircraft; the whole aircraft has no torsion along the span direction. Of course, the selection of the airfoil can be further optimized and adapted according to the performance requirements;

4) The integrated rectenna is a circular plane with a diameter of 4.502m. The center of the circle is 3.015m away from the nose. The integrated rectenna is composed of skin, upper and lower foam layers, and rectenna packaging. The energy receiving capacity is about 1km/ ^m2 , and the converted cruise power supply and demand requirement is 8KW;

5) The tail strut is 6m long, and the area of each V-tail is ^1.8375m2 . 40% of the area of the trailing edge is the rudder surface, with a span of 2.5m and a chord length of 0.3m, which is mainly used for longitudinal and lateral attitude control;

6) The aileron is located on the outside of the wing, close to the wing tip, with a span of 3.5m and a chord length of 0.15m. It participates in lateral control together with the V-tail;

7) A two-blade electric tail propeller is used as the power device. The propeller blade adopts the S1223 blade element airfoil, the torsion angle from the root to the tip is 60 degrees, and the blade diameter is 2.5m. Of course, the blade can be further optimized and improved according to the requirements of the propulsion device;

8) The front three-point landing gear is adopted. The front lift is retracted upward into the upper compartment of the fuselage due to the influence of the lower surface antenna, and the main lift is retracted forward into the landing gear compartment on the lower surface.

The high-altitude, ultra-long-duration, high-aspect-ratio integrated unmanned aerial platform finally obtained by the present invention adopts a high-aspect-ratio wing-body fusion layout, and adds winglets at the wingtips to further reduce induced drag. As the source of the whole aircraft power device, the rectifier receiving antenna is integrated with the aerodynamic and structural design of the whole aircraft, which ensures the aerodynamic characteristics of the whole aircraft from the appearance. Since the rectifier receiving antenna adopts a composite design form such as skin and foam similar to the surface structure of the fuselage, the two can achieve a high degree of structural unity and integration, and have similar mechanical properties. Therefore, its load-bearing characteristics can be considered as a whole, effectively avoiding the disadvantage of poor structural weight characteristics after the external rectifier antenna is installed.

During the actual design and processing, the full platform design of the rectenna is considered from the perspective of overall integration, and the transferability and relevance of various overall, aerodynamic, structural and other parameters are fully considered. When entering the detailed design stage, the sub-component design under the constraints of the integrated platform can be adopted to speed up the iteration process. In the present invention, the constraints and design indicators are decomposed through an integrated approach, and interactive design is performed on the platform and the antenna, and finally modular research and development under integration is realized. The rectenna is combined with the platform in the form of an integrated module, that is, the antenna and the platform can be completely combined into an integrated system through a simple connection relationship.

Although the embodiments of the present invention have been shown and described above, it is to be understood that the above embodiments are illustrative and are not to be construed as limitations on the present invention. A person skilled in the art may change, modify, substitute and modify the above embodiments within the scope of the present invention without departing from the principles and purpose of the present invention.

Claims (10)

1. A design method for a high-altitude, ultra-long-flight, large-aspect-ratio integrated unmanned aerial platform, characterized in that it comprises the following steps: Step 1: According to the design requirements, the basic characteristics of the unmanned aerial platform and the antenna load are determined; the energy source of the unmanned aerial platform is the microwave radio frequency energy radiated by the ground transmitting end, which is transmitted through space to reach the receiving antenna arranged on the unmanned aerial platform, and then further converted into direct current; Step 2: According to the antenna load characteristics determined in step 1, determine the constraints and requirements of the antenna load on the unmanned aerial platform; Step 3: Determine the basic layout of the unmanned aerial platform based on the constraints and requirements of the antenna load on the unmanned aerial platform determined in step 2, as well as the design performance requirements; Step 4: Determine the airfoil of the inner and outer wing sections based on the constraints and requirements of the antenna load on the unmanned aerial platform: Step 5: Determine the shape and parameters of the unmanned aerial platform by integrating the basic characteristics of the antenna payload and the unmanned aerial platform, the constraints and requirements of the antenna payload on the unmanned aerial platform, and the basic layout of the unmanned aerial platform: Based on the wing loading parameters of a typical high aspect ratio aircraft As well as the platform constraints and requirements in step 2 and the basic platform layout in step 3, determine the wing area S, and then obtain the span b _l according to the initially determined aspect ratio; Determine the engine thrust T and the corresponding power requirement P based on the takeoff weight m and the estimated cruise lift-to-drag ratio (L/D) _c ; According to the antenna energy density constraint, the antenna diameter _be is obtained; According to the static stability margin requirements and the empirical data of the tail capacity of typical V-tail aircraft, the chordwise position of the V-tail and the size of the V-tail are determined by design; Determine the initial position and size of ailerons and V-tail control surfaces based on typical V-tail aircraft handling characteristics data; Determine the propeller diameter, number of blades, and blade shape based on the estimated cruise lift-to-drag ratio and power requirements; Based on the above parameters, the aerodynamic shape of a high aspect ratio flight platform integrating an integrated receiving rectenna is constructed; Step 6: Simulate and calculate the aerodynamic characteristics of the high aspect ratio flight platform constructed in step 5 to determine the initial aerodynamic performance of the flight platform, including the aerodynamic performance analysis of the take-off and landing configuration and the aerodynamic performance analysis of the cruise configuration. The aerodynamic performance analysis of the take-off and landing configuration includes determining the lift coefficient, drag coefficient and pitch moment coefficient curves _CL ~α, _CD ~α, _Cm ~α from 0 degrees to the stall angle of attack of the take-off and landing configuration; the aerodynamic performance analysis of the cruise configuration includes determining the lift coefficient, drag coefficient and pitch moment coefficient curves _CLp ~α, _CDp ~α, _Cmp ~α from 0 degrees to the stall angle of attack of the cruise configuration, the cruise trim lift-to-drag ratio (L/D) _c and the cruise static stability margin; Step 7: Based on the initial flight platform aerodynamic performance simulation results obtained in step 6, the initial flight platform multi-objective aerodynamic optimization is carried out with the optimization objectives of improving the cruise lift-to-drag ratio (L/D) _c and reducing the longitudinal nose-down moment C _m of the entire aircraft, and a single aerodynamic optimization design is performed. Based on the aerodynamic shape optimization results, the aerodynamic load distribution is determined; Step 8: Based on the aerodynamic load distribution obtained in step 7, design the internal structure of the rectenna and the structure of the flight platform, and complete the preliminary structural finite element characteristic analysis; Step 9: Based on the preliminary structural finite element characteristic analysis in step 7, determine the embedding and fusion method of the rectenna distributed antenna structure, and simulate and analyze the electrical characteristics of the design scheme; Step 10: Based on the electrical performance characteristics obtained in step 9, optimize the structural/electrical integration design of the integrated antenna platform; Step 11: Based on the platform shape layout plan, structural finite element plan and rectenna electrical characteristics plan obtained in steps 7-10, conduct multidisciplinary collaborative optimization design of aerodynamic/structural/electrical radiation characteristics to complete the design of a high-altitude, ultra-long-flight, large-aspect-ratio integrated unmanned flying platform. 2. According to the design method of a high-altitude, ultra-long-flight, large-aspect-ratio integrated unmanned aerial platform according to claim 1, it is characterized in that: in step 1, The basic characteristics of the antenna payload include: the geometric shape of the antenna payload, the receiving area, the antenna diameter, the total thickness, the energy conversion efficiency and the energy density; The basic characteristics of the unmanned aerial platform include: aspect ratio, cruise lift-to-drag ratio, cruise lift coefficient, take-off gross weight, wing area and static stability margin requirements. 3. The design method of a high-altitude, ultra-long-flight, large-aspect-ratio integrated unmanned aerial platform according to claim 2 is characterized in that: in step 2, the constraints and requirements of the antenna load on the platform include: 1) According to the receiving area of the antenna load, determine the area of the antenna load layout on the unmanned aerial platform, requiring the antenna load to have sufficient layout area on the unmanned aerial platform; 2) The antenna load layout area is a flat area, and the maximum structural deformation is not greater than the set requirements; 3) The surface skin thickness of the antenna load arrangement area meets the total thickness requirements of the antenna load. 4. According to the design method of a high-altitude, ultra-long-flight, large-aspect-ratio integrated unmanned aerial platform as described in claim 3, it is characterized in that: in step 3, the basic layout form of the unmanned aerial platform includes: 1) The unmanned aerial platform adopts a high aspect ratio wing-body fusion layout and is driven by a tail-mounted electric propeller; 2) According to the static stability margin requirements, the unmanned aerial platform adopts an inverted V-tail form to avoid the influence of propeller slipstream on the efficiency of the V-tail control surface; 3) According to the deformation constraint of antenna structure, antenna fusion design is performed under the fuselage of the flight platform; 4) Determine the area of the outer wing and inner wing-body combination according to the aspect ratio and lift-to-drag ratio requirements; 5) Use winglets; 6) Adopt the conventional three-point take-off and landing method; 7) Ailerons are located on the outer wing sections and control roll; pitch and yaw are controlled by the V-tail control surfaces. 5. According to the design method of a high-altitude, ultra-long-flight, large-aspect-ratio integrated unmanned aerial platform as described in claim 3, it is characterized in that: in step 4, the outer wing section adopts a high lift-to-drag ratio airfoil with a straight lower surface, and the airfoil lift-to-drag ratio is not less than the set requirement; the airfoil of the inner wing-body fusion section is further optimized on the basis of the high lift-to-drag ratio airfoil to obtain: Design status: Ma = 0.1356, Reynolds number Re = 0.34E6, angle of attack α = 0° Design goal: maximum lift-to-drag ratio Constraints: Moment coefficient Cm ≥ -0.13 The maximum relative thickness T/C is greater than or equal to 15% The lower surface of the inner wing-body fusion section in the constrained antenna area is a straight section: the area is restricted to y(x)=0, x∈(x1,x2), (x1,x2) are the coordinates of the lower surface of the inner wing-body fusion section airfoil corresponding to the antenna area. 6. According to the design method of a high-altitude, ultra-long-flight, large-aspect-ratio integrated unmanned aerial platform as described in claim 1, it is characterized in that: in step 7, the design goal is: The cruise lift-to-drag ratio is the largest (L/D) _c , and the longitudinal nose-down moment of the entire aircraft is the smallest min 1/(L/D) _c min |C _m | The design constraints are: The cruise lift coefficient is not less than the initial value C _Lcinitial : C _Lc ≥C _Lcinitial The antenna area does not change: Area _e = Area _einitial The aspect ratio of the whole machine remains unchanged: 7. According to claim 6, a high-altitude, ultra-long-flight, large-aspect-ratio integrated unmanned aerial platform design method is characterized in that: in step 8, the internal structure of the rectenna and the structure of the unmanned aerial platform are designed, and the preliminary structural finite element characteristic analysis is completed, including: According to the wing aerodynamic load and aeroelastic requirements, it is determined to adopt a double-beam structure that takes into account both load-bearing and deformation efficiency, and the preliminary front and rear beam positions, relative thickness and height distribution are given; According to the aerodynamic load constraints of the rectenna area, the antenna structure is determined to be a multi-layer composite material, the total thickness δ _e is consistent with the wing skin thickness, and the multi-layer structure includes a package surface composite material and a honeycomb layer; According to the connection method between the antenna area and the platform, the periphery of the antenna area of the platform is structurally reinforced, and the antenna structure and the platform are connected by bonding; According to the bending moment bearing requirements of the whole aircraft, box beams are arranged transversely inside the fuselage; According to the appearance design, the V-tail tail strut is directly inserted into the fuselage with a carbon tube; According to the load-bearing requirements, multiple frames are arranged longitudinally in the wings and fuselage to connect the tail strut, landing gear, engine and skin; According to the overall stress condition of the unmanned aerial platform, a finite element analysis is conducted on the structural characteristics of the entire aircraft to determine the maximum deformation of the entire aircraft and the stress concentration points. 8. According to claim 7, a high-altitude, ultra-long-duration, large-aspect-ratio integrated unmanned aerial platform design method is characterized in that: in step 9, determining the embedding and fusion mode of the rectenna distributed antenna structure, and simulating and analyzing the electrical characteristics of the design scheme includes: According to the structural design and stress conditions of the unmanned aerial platform, the integrated rectenna is embedded in the multi-layer composite material structure of the antenna area to form a foam sandwich. The sandwich structure of the antenna part is: packaging surface composite material, upper honeycomb, antenna layer, lower honeycomb, composite material. According to the determined total thickness constraint δ _e of the platform skin, the electric radiation performance parameters of the integrated receiving rectenna are determined, including the directivity pattern, peak gain and coupling degree; the electric performance is simulated to obtain the electric radiation performance characteristics. 9. According to the design method of a high-altitude, ultra-long-flight, large-aspect-ratio integrated unmanned aerial platform as described in claim 8, it is characterized in that: in step 10, the structural/electrical integration design optimization process of the integrated antenna platform is: The objective function is the radiation performance f(x) and the overall structural characteristics g(x), that is, the optimal mechanical performance and the optimal electrical radiation performance are taken as the optimization targets; a is taken as the skin thickness, b is the foam thickness, δ _wf is the front beam thickness, δ _wb is the rear beam thickness, δ _r1 ,…δ _rn is the thickness of the longitudinally arranged wing ribs, and the mechanical and electrical integration design mathematical model of the integrated array is established as follows: The design goals are: the lightest overall structural weight g(x) and the best electrical radiation performance f(x) min 1/f(x) min g(x) The design constraints are: Skin thickness limit range a _min ≤ a ≤ _{a max} The foam thickness is limited to b _min ≤ b ≤ b _max Front beam thickness limit range δ _wfmin ≤δ _wf ≤δ _wfmax Rear beam thickness limit range δ _wbmin ≤δ _wb ≤δ _wbmax Rib thickness limit Limit range of wing tip deformation: e _wt ≤ e _wtmax The optimization algorithm is used to optimize the structural/electrical integration design of the fusion integrated antenna platform to obtain the layout plan with the best structural performance and electrical radiation performance. 10. The design method of a high-altitude, ultra-long-flight, large-aspect-ratio integrated unmanned aerial platform according to claim 9, characterized in that: in step 11, the process of multidisciplinary collaborative optimization design of aerodynamic/structural/electric radiation characteristics is: given the design variables, including: Aerodynamics: wingspan, leading edge sweep angle, trailing edge sweep angle, dihedral installation angle, characteristic section thickness, characteristic section curvature, characteristic section maximum thickness position, root-to-wing ratio and winglet camber angle; Structural aspects: front beam thickness, rear beam thickness, fuselage position reinforcement beam position, reinforcement beam thickness, characteristic section wing rib thickness, antenna area foam thickness, packaging composite material thickness and wing tip deformation; Electrical characteristics: antenna element size, number of antenna elements and antenna element arrangement; The design goals are: Aerodynamics: The cruise lift-to-drag ratio of the whole aircraft is not less than the optimized value in step 7 (L/C) _c ≥ (L/C) _c-firstopt Structural aspect: The weight of the whole machine structure does not exceed the optimization value in step 10: g(x)≤g(x) _firstopt Electrical performance: The radiation performance of the whole machine is not less than the optimized value in step 10: f(x)≥f(x) _firstopt The design constraints are: Aerodynamic shape constraints: The cruise lift coefficient is not less than the optimized value in step 7: C _Lc ≥ _{C Lcfirstopt} The antenna area does not change: Area _e = Area _einitial The aspect ratio of the whole machine remains unchanged: Structural constraints: Skin thickness limit range a _min ≤ a ≤ a _max The foam thickness is limited to b _min ≤ b ≤ b _max Front beam thickness limit range δ _wfmin ≤δ _wf ≤δ _wfmax Rear beam thickness limit range δ _wbmin ≤δ _wb ≤δ _wbmax Rib thickness limit Wing tip deformation limit range e _wt ≤ e _wtmax Electrical characteristics constraints: Array element effective area limit range: S _es ≤ S _esmax Array element quantity limit range: _Nes ≤ _Nesmax An optimization algorithm is used to conduct multidisciplinary collaborative optimization design of aerodynamic/structural/electric radiation characteristics, and a solution for a high-altitude, ultra-long-flight, large-aspect-ratio integrated unmanned flying platform is obtained.