阅读说明：本技术 一种大展弦比高强度双层机翼太阳能无人机 (High-aspect-ratio and high-strength double-layer wing solar unmanned aerial vehicle ) 是由 张声伟 张超 明亚丽 杨天星 王延风 于 2020-12-24 设计创作，主要内容包括：本发明属于航空飞行器技术领域,公开了一种大展弦比高强度双层机翼太阳能无人机,包括机身、机翼和尾翼；其中,机翼是横向一贯制机翼,机翼有两个,分别为上层机翼和下层机翼,机身设在上层机翼和下层机翼之间,上层机翼和下层机翼通过翼尖支撑墙连接；上层机翼和下层机翼的上表面都设有太阳能板；尾翼设在机身尾部。本发明的双层联翼布局在满足结构强度限制的前提下,使机翼展弦比提高到30以上,这对于设计升力系数大的飞机气动减阻,意义重大；翼展减小29％,结构重量减小22％,可显著提高飞机的任务载荷与连续飞行时间；机翼的抗弯、抗扭能力强；并能还提供满意的三轴稳定性与操纵能力。(The invention belongs to the technical field of aviation aircrafts, and discloses a high-aspect-ratio and high-strength double-layer wing solar unmanned aerial vehicle which comprises a vehicle body, wings and an empennage; the aircraft comprises an aircraft body, wings, a wing tip supporting wall and a wing tip supporting wall, wherein the wings are transversely and integrally formed, and the two wings are an upper layer wing and a lower layer wing respectively; solar panels are arranged on the upper surfaces of the upper layer wing and the lower layer wing; the empennage is arranged at the tail part of the machine body. The double-layer linked wing layout provided by the invention has the advantages that on the premise of meeting the structural strength limitation, the wing aspect ratio is improved to more than 30, and the double-layer linked wing layout is significant for designing the aerodynamic drag reduction of an airplane with a large lift coefficient; the wingspan is reduced by 29%, the structure weight is reduced by 22%, and the mission load and the continuous flight time of the airplane can be obviously improved; the bending resistance and torsion resistance of the wing are strong; and can also provide satisfactory triaxial stability and handling capability.)

1. A high-aspect-ratio and high-strength double-layer wing solar unmanned aerial vehicle is characterized by comprising a vehicle body (3), wings and a tail wing; the airplane wing assembly comprises two wings, namely an upper wing (2) and a lower wing (1), wherein the airplane body is arranged between the upper wing (2) and the lower wing (1), and the upper wing (2) and the lower wing (1) are connected through wing tip supporting walls (17); solar panels are arranged on the upper surfaces of the upper layer wing (2) and the lower layer wing (1); the empennage is arranged at the tail part of the machine body.

2. The high-aspect-ratio and high-strength double-layer wing solar unmanned aerial vehicle as claimed in claim 1, wherein the number of the fuselages (3) is two, the two fuselages are arranged between the upper layer wing (2) and the lower layer wing (1), and the two fuselages (3) are symmetrical relative to the center line of the aircraft.

3. The high-aspect-ratio and high-strength double-layer wing solar unmanned aerial vehicle according to claim 2, wherein the front edge of the lower layer wing (1) is swept backward by 3 degrees, the rear edge of the lower layer wing (1) is a straight line, the root of the wing adopts a high-lift wing profile with the thickness of 20 percent, the wing tip profile is 13 percent, the aerodynamic torsion angle is-3.6 degrees, and the outer wing of the lower layer wing (1) is turned up by 8 degrees from 70 percent; the inner side of the lower wing (1) with 8% span is loaded with batteries.

4. The high-aspect-ratio and high-strength double-layer wing solar unmanned aerial vehicle according to claim 3, wherein the front edge of the upper layer wing (2) is a straight line, and the rear edge is swept forward by 3 degrees; the axial position of the front edge of the upper layer wing (2) is the same as the axial position of the rear edge of the lower layer wing (1); the upper layer wing (2) is a high lift wing type, the thickness of the wing is reduced by 20% compared with the lower layer wing, and the wing tip of the upper layer wing (2) is connected with the tip part of the lower layer wing (1) through a wing tip supporting wall (17) to form a wing box structure with the double bodies.

5. The high-aspect-ratio and high-strength double-layer wing solar unmanned aerial vehicle is characterized in that the cross section of the fuselage (3) is in an oval shape, a flight control system, an avionics system and a main landing gear cabin are arranged inside the fuselage (3), and the external structure of the fuselage (3) is a connecting wall of an upper layer wing (2) and a lower layer wing (3); the length of the fuselage (3) is 1/5 of wingspan, and the tail part of the fuselage (3) is quickly contracted into a tail wing stay bar.

6. The high-aspect-ratio and high-strength double-layer wing solar unmanned aerial vehicle as claimed in claim 2, wherein the tail wing is a T-shaped tail wing composed of a horizontal tail and a vertical tail, the tail wing is provided with two tail wings which are respectively arranged at the tail parts of the two bodies (3), and the tail capacity of the horizontal tail is 33% of that of the conventional layout.

7. The high-aspect-ratio and high-strength double-layer wing solar unmanned aerial vehicle according to claim 2, further comprising an elevator (7) installed behind the horizontal tail (6), an inner motorized flap (14) and an outer motorized flap (15) at the rear edge of the upper layer wing (2), and an elevon (16) at the outermost rear edge of the upper layer wing (2), wherein longitudinal operation is controlled by controlling the elevator (7), the inner motorized flap (14), the outer motorized flap (15) and the elevon (16);

the electric steering engine also comprises a rudder (5) behind the vertical tail (4) and two electric propellers (18) symmetrically arranged on two sides of the engine body; the course control is controlled by the rudder (5) and the differential rotation speed of the two motors (18);

the aircraft wing assembly further comprises an aileron (13) at the outermost side rear edge of the lower layer wing (1), an inner side hyperplasia flap (10) and an outer side hyperplasia flap (11) at the rear edge of the lower side wing (1), and rolling operation is controlled by controlling the aileron (13), the lifting aileron (16), the inner side hyperplasia flap (10) and the outer side hyperplasia flap (11).

8. The high-aspect-ratio and high-strength double-layer wing solar unmanned aerial vehicle is characterized in that the number of main landing gears (8) is 2, the main landing gears are respectively positioned at the front parts of the two fuselages (3) and are collected in landing gear cabins at the front parts of the fuselages (3) before the gravity centers of the fuselages (3); and 2 light rear landing gears (9) are respectively arranged at the bottoms of the vertical tails of the two airframes and are retracted into the vertical tail (4) after taking off.

9. The high-aspect-ratio and high-strength double-layer wing solar unmanned aerial vehicle as claimed in claim 1, wherein the power of the power is 10 electric propellers (18), the 10 electric propellers (18) are respectively arranged on the upper wing and the lower wing, the lower wing (1) is provided with 5 electric propellers (18), the upper wing (2) is provided with 5 electric propellers (18), the 8 electric propellers (18) are symmetrical according to two sides of a central axis of the aircraft, and the 2 electric propellers (18) are arranged on the central axis.

Technical Field

The invention belongs to the technical field of aviation aircrafts, relates to a solar unmanned aerial vehicle, and particularly relates to a high-aspect-ratio high-strength double-layer wing solar unmanned aerial vehicle.

Background

The task load of the solar unmanned aerial vehicle which is successfully tried out at present is small, the aspect ratio is small, the pneumatic efficiency is not high, and the day and night flight can be achieved more than a few. The large-scale high-aspect-ratio solar unmanned aerial vehicle is disturbed by gust wind due to the problems of wing strength, rigidity and aeroelasticity, wings are broken, and the aircraft crashes, such as the American 'Sunshen' solar unmanned aerial vehicle.

The great task system of weight can't be installed to current well small-size solar energy unmanned aerial vehicle, and its energy system also can't provide long-time task system power supply, because structure weight is great, its power-to-weight ratio is less, hardly realizes crossing long-time flight round the clock, therefore engineering use value is not big.

The difficult point of large-scale solar energy unmanned aerial vehicle design lies in: the structure weight is reduced, the task load is increased, and the strength, the rigidity and the aerodynamic efficiency of the wing are improved. As the span and wing area increase, the structural weight increases significantly. The conventional layout is difficult to reduce the structural weight, the wing aspect ratio cannot be too large due to the limitation of the strength and the rigidity of the wing, and therefore the aerodynamic efficiency is not high.

Disclosure of Invention

In order to solve the problems, the invention provides a high-aspect-ratio high-strength double-layer wing solar unmanned aerial vehicle which has a large task load and the ability of flying at high altitude for a long time in a day and night mode, and can perform tasks such as reconnaissance, early warning and communication relay in a subcritical space.

The technical scheme of the invention is as follows:

a high-aspect-ratio and high-strength double-layer wing solar unmanned aerial vehicle comprises a vehicle body, wings and an empennage; the aircraft comprises an aircraft body, wings, a wing tip supporting wall and a wing tip supporting wall, wherein the wings are transversely and integrally formed, and the two wings are an upper layer wing and a lower layer wing respectively; solar panels are arranged on the upper surfaces of the upper layer wing and the lower layer wing; the empennage is arranged at the tail part of the machine body.

Furthermore, the number of the two fuselages is two, the two fuselages are arranged between the upper layer wing and the lower layer wing, and the two fuselages are symmetrical relative to the center line of the airplane.

Furthermore, the front edge of the lower layer wing is swept back by 3 degrees, the rear edge of the lower layer wing is a straight line, the root of the wing adopts a high-lift wing type with the thickness of 20 percent, the wing section at the tip of the wing has the thickness of 13 percent, the pneumatic torsion angle is-3.6 degrees, and the outer wing of the lower layer wing is turned up by 8 degrees from 70 percent; the storage battery is loaded on the inner side of the 8% wingspan of the lower-layer wing, so that the horizontal direction moment of inertia can be reduced.

Further, the front edge of the upper layer wing is a straight line, and the rear edge of the upper layer wing is swept forward by 3 degrees; the axial position of the front edge of the upper layer wing is the same as that of the rear edge of the lower layer wing; the sunlight incident angle can be ensured to be 90 degrees, the lower-layer wing can not be shielded, and the aerodynamic efficiency reduction caused by the downwash airflow of the front wing can be avoided. The upper layer wing is a high lift wing, the thickness of the wing is reduced by 20% compared with the lower layer wing, and the wing tip of the upper layer wing is connected with the tip part of the lower layer wing by a wing tip support wall to form a wing box structure with the double bodies. The wing tip support wall can reduce the transverse flow of the upper wing and the lower wing and can also increase the torsional rigidity of the wings.

Under the same wing area and aspect ratio, compared with a single-wing aircraft, the wing span of the double-layer linked wing type wing can be reduced by 29%, and the structural mass can be reduced by 22%. The reduction in structural weight may increase the power-to-weight ratio and mission capacity of the aircraft. The double-fuselage, the connecting wall of the wing tip and the upper and lower wings form a bearing wing box, so that the strength and the rigidity of the wing with the large aspect ratio are effectively improved.

Furthermore, the cross section of the fuselage is in a vertical oval shape, so that the distance between the upper wing and the lower wing can be increased, the pneumatic interference is reduced, a flight control system, an avionic system and a main landing gear cabin are arranged in the fuselage, and the external structure of the fuselage is a connecting wall of the upper layer of wings and the lower layer of wings to form a bearing wing box structure; the length of the fuselage is only 1/5 with wingspan, and the tail of the fuselage is quickly contracted into an empennage brace rod, so that the structural weight is effectively reduced.

Furthermore, the tail wing is a T-shaped tail wing composed of a horizontal tail and a vertical tail, the T-shaped tail wing is provided with two tail wings which are respectively arranged at the tail parts of the two airframes, the longitudinal net stability of the double-layer linked wing layout in a tailless state can reach more than-0.035, so that the area of the horizontal tail is reduced, the tail capacity of the horizontal tail is only 33 percent of that of the conventional layout, and the requirement on the longitudinal stability can be met. Reduction of vertical tail area and side force derivative C_YβThe anti-crosswind disturbance and track control capability of the airplane can be improved.

The elevator is arranged behind the horizontal tail, the inner side motor flap and the outer side motor flap of the rear edge of the upper layer wing, and the elevon of the outermost side rear edge of the upper layer wing, and the longitudinal operation is controlled by controlling the elevator, the inner side motor flap, the outer side motor flap and the elevon;

the steering engine also comprises a rudder behind the vertical tail and two motors symmetrically arranged on two sides of the engine body; the course control is controlled by the rudder and the differential rotation speed of the two motors;

the wing-type aircraft further comprises an aileron at the outermost side trailing edge of the lower layer of wing, an inner side hyperplasia flap and an outer side hyperplasia flap at the trailing edge of the lower side wing, and the rolling operation is controlled by controlling the aileron, the lifting aileron, the inner side hyperplasia flap and the outer side hyperplasia flap. The maximum lift coefficient of the low-speed configuration of the airplane can be improved, so that the maximum takeoff speed from the ground is reduced, and the takeoff is assisted.

Furthermore, 2 main undercarriages are respectively positioned at the front parts of the two fuselages and are collected in an undercarriages at the front parts of the fuselages before the gravity centers of the fuselages; and 2 light rear landing gears are arranged at the bottoms of the vertical tails of the two airframes respectively and are retracted into the vertical tails after taking off. The take-off and landing modes of the airplane are the same as those of a rear three-point landing gear airplane.

Furthermore, the power of the airplane is 10 electric propellers, the 10 electric propellers are respectively arranged on the upper wing and the lower wing, the lower wing is provided with 5 electric propellers, the upper wing is provided with 5 electric propellers, 8 electric propellers are symmetrically arranged on two sides of the central axis of the airplane respectively, 2 electric propellers are arranged on the central axis, the distributed power system can utilize propeller slipstream to increase the maximum lift coefficient and the stall angle of attack, and the course control can also be realized by utilizing the differential rotating speed of motors on two sides.

The invention has the advantages that:

(1) the double-layer linked wing type wing design can reduce the wing span by 29% and the structural mass by 22% compared with a single wing aircraft under the same wing area and aspect ratio.

(2) The span-chord ratio of the double-layer linked wing type wing is doubled compared with that of a single wing machine under the condition of the same wing area and the same span length, so that the double-layer linked wing type wing is very favorable for reducing induced resistance and improving aerodynamic performance.

(3) The upper wing moved backwards increases the longitudinal static stability of the airplane, and the longitudinal control capability provided by the maneuvering flap and the elevon can reduce the area of the horizontal tail.

(4) The double-fuselage layout design can be used as a flight control system, equipment and a main landing gear cabin, and can also be used as a connecting wall of an upper wing and a lower wing to form a bearing wing box structure, so that the torsional rigidity of the wings is improved, and the aeroelastic deformation of the wings is reduced.

(5) The 4-wheel type undercarriage is simple in structure, light in weight, small in occupied space and beneficial to reducing drag of the appearance of the airplane.

(6) The multifunctional wing surface control system with the assistance of differential power is designed, so that the area of the tail wing is reduced by 60 percent, and the structural weight of the airplane is effectively reduced.

(7) The distributed power system design can utilize the slipstream of the propeller to increase the maximum lift coefficient and the stall attack angle, and can also utilize the differential rotation speed of the motors at two sides to realize course control.

(8) The small tail capacity empennage structure design can lighten the structural mass, improve the anti-crosswind disturbance capability of the winglet-mounted airplane and accurately control the flight path. With the support of the multifunctional airfoil control system, good three-axis control capability can be provided

(9) The wing tip connecting wall increases the torsional rigidity of the wing, reduces the transverse flow of the upper wing and the lower wing, is favorable for reducing the induced resistance and improves the pneumatic efficiency.

Drawings

FIG. 1 is a top view of an embodiment of the present invention;

FIG. 2 is a front view of an embodiment of the present invention;

FIG. 3 is a side view of an embodiment of the present invention;

wherein 1 — lower wing; 2-upper wing; 3-a fuselage; 4, hanging the tail; 5-rudder; 6, flattening the tail; 7-elevator; 8 — main landing gear; 9 — rear landing gear; 10-inboard high-lift flaps; 11-outboard high-lift flaps; 12-lower layer wing outside flaperon; 13-ailerons; 14-inboard motorized flaps; 15-outboard motorized flaps; 16-elevon; 17-wing tip support wall; 18-electric propeller.

Detailed Description

This section is an example of the present invention and is provided to explain and illustrate the technical solutions of the present invention.

A high-aspect-ratio high-strength double-layer wing solar unmanned aerial vehicle, as shown in fig. 1, 2 and 3, comprises a vehicle body 3, wings and a tail wing; the airplane wing structure comprises two wings, namely an upper wing 2 and a lower wing 1, wherein the airplane body is arranged between the upper wing 2 and the lower wing 1, and the upper wing 2 and the lower wing 1 are connected through a wing tip supporting wall 17; solar panels are arranged on the upper surfaces of the upper layer wing 2 and the lower layer wing 1; the empennage is arranged at the tail part of the machine body.

The two fuselages 3 are arranged between the upper layer wing 2 and the lower layer wing 1, and the two fuselages 3 are symmetrical relative to the center line of the airplane.

The front edge of the lower layer wing 1 is swept backward by 3 degrees, the rear edge is a straight line, the root part of the wing adopts a high-lift wing profile with the thickness of 20 percent, the wing tip profile is 13 percent in thickness, the aerodynamic twist angle is-3.6 degrees, and the outer wing of the lower layer wing 1 is turned up by 8 degrees from 70 percent; the storage battery is loaded on the inner side of the 8% wingspan of the lower-layer wing 1, so that the horizontal direction inertia moment can be reduced.

The front edge of the upper layer wing 2 is a straight line, and the rear edge is swept forward by 3 degrees; the axial position of the front edge of the upper layer wing 2 is the same as the axial position of the rear edge of the lower layer wing 1; the sunlight incident angle can be ensured to be 90 degrees, the lower-layer wing can not be shielded, and the aerodynamic efficiency reduction caused by the downwash airflow of the front wing can be avoided. The upper layer wing 2 is a high lift wing type, the thickness of the wing is reduced by 20% compared with the lower layer wing, the wing tip of the upper layer wing 2 is connected with the tip part of the lower layer wing 1 through a wing tip supporting wall 17, and a wing box structure is formed by the upper layer wing 2 and the lower layer wing. The wing tip support wall 17 can reduce the transverse flow of the upper and lower wings and can also increase the torsional stiffness of the wings.

Under the same wing area and aspect ratio, compared with a single-wing aircraft, the wing span of the double-layer linked wing type wing can be reduced by 29%, and the structural mass can be reduced by 22%. The reduction in structural weight may increase the power-to-weight ratio and mission capacity of the aircraft. The double-fuselage, the connecting wall of the wing tip and the upper and lower wings form a bearing wing box, so that the strength and the rigidity of the wing with the large aspect ratio are effectively improved.

The section of the fuselage 3 is in a vertical oval shape, so that the distance between an upper wing and a lower wing can be increased, the aerodynamic interference is reduced, a flight control system, an avionic system and a main landing gear cabin are arranged in the fuselage 3, and the external structure of the fuselage 3 is a connecting wall of an upper layer of wings 2 and a lower layer of wings 3 to form a bearing wing box structure; the length of the fuselage 3 is only 1/5 of wingspan, and the tail of the fuselage 3 is quickly contracted into an empennage brace, so that the structural weight is effectively reduced.

The tail wing is a T-shaped tail wing composed of a horizontal tail and a vertical tail, the T-shaped tail wing is provided with two tail wings which are respectively arranged at the tail parts of the two machine bodies 3, the longitudinal net stability of the double-layer linked wing layout in a tailless state can reach more than-0.035, the area of the horizontal tail is reduced, the tail capacity of the horizontal tail is only 33 percent of that of the conventional layout, and the requirement of the longitudinal stability can be met, and the horizontal tail elevator [7 ] is provided with an upper-layer wing elevator aileron 16]The longitudinal operation requirement can be met. Reduction of vertical tail area and side force derivative C_YβThe anti-crosswind disturbance and track control capability of the airplane can be improved.

The aircraft further comprises an elevator 7 arranged behind the horizontal tail 6, an inner maneuvering flap 14 and an outer maneuvering flap 15 at the rear edge of the upper layer wing 2, and an elevating aileron 16 at the outermost rear edge of the upper layer wing 2, wherein the longitudinal operation is controlled by controlling the elevator 7, the inner maneuvering flap 14, the outer maneuvering flap 15 and the elevating aileron 16;

the electric steering engine also comprises a rudder 5 behind the vertical tail 4 and two electric propellers 18 symmetrically arranged at two sides of the engine body; the course control is controlled by the rudder 5 and the differential rotation speed of the two motors 18;

the aircraft also comprises an aileron 13 at the outermost rear edge of the lower layer wing 1, an inner hyperplasia flap 10 and an outer hyperplasia flap 11 at the rear edge of the lower layer wing 1, and rolling operation is controlled by controlling the aileron 13, the elevon 16, the inner hyperplasia flap 10 and the outer hyperplasia flap 11. The maximum lift coefficient of the low-speed configuration of the airplane can be improved, so that the maximum takeoff speed from the ground is reduced, and the takeoff is assisted.

Two main undercarriages 8 are respectively positioned at the front parts of the two fuselages 3 and are collected in the undercarriage cabins at the front parts of the fuselages 3 before the gravity centers of the fuselages 3; and two light rear landing gears 9 are arranged at the bottoms of the vertical tails of the two airframes respectively and are retracted into the vertical tail 4 after taking off. The take-off and landing modes of the airplane are the same as those of a rear three-point landing gear airplane.

The power of the airplane is 10 electric propellers, the 10 electric propellers are respectively arranged on an upper wing and a lower wing, the lower wing 1 is provided with 5 electric propellers, the upper wing 2 is provided with 5 electric propellers, 8 electric propellers are respectively symmetrical according to two sides of a central axis of the airplane, the central axis is provided with 2 electric propellers, a distributed power system can utilize propeller slipstream to increase the maximum lift coefficient and the stall attack angle, and the course control can also be realized by utilizing the differential rotating speed of motors at two sides.

Another embodiment of the present invention is described below.

An unmanned aerial vehicle according to the inventive concept was designed in the following 9 steps.

Step 1. performance requirement determination: the task load of the example airplane is not less than 80kg, the maximum flying weight is not more than 1200kg, the lift-drag ratio for cruising is not less than 30, the cruising height is 22000m, and the continuous flying time is not less than 55 days;

and 2, optimizing the wing load and the wing aspect ratio by taking the flight time as an optimization target, wherein the wing load of the example airplane is 4.6kg/m2, the total wing area is 256m2, the single wing area is 128m2, the aspect ratio is 30, and the wingspan is 62 m.

And 3, designing the lower-layer wing of the double-layer linked wing type wing: the front edge of the wing is swept by 3 degrees, the rear edge is a straight line, and the root ratio is 0.5. The root part of the wing adopts a high-lift wing profile with the thickness of 20 percent, the wing tip profile with the thickness of 13 percent and the aerodynamic twist angle of-3.6 degrees, and the outer wing is turned up by 8 degrees from the span-wise position to 70 percent. The trailing edge of the outer wing is an aileron, the spanwise position is 0.7-0.98, the length of the aileron chord is 30% of the local chord length, the maximum deflection angle is plus or minus 20 degrees, the trailing edge of the wing between the two bodies is a high lift flap, and the chord length of the flap is 0.35% of the local chord length.

Step 4, designing the upper-layer wing of the double-layer linked wing type wing: slightly more than the lower layer wing, the front edge of the lower layer wing is a straight line, the rear edge of the lower layer wing is swept forward by 3 degrees, and the axial position of the lower layer wing is positioned at the rear edge of the lower layer wing. The wing adopts a high-lift wing type, the thickness of the wing is reduced by 20% compared with that of the lower-layer wing, and the wing tip is connected with the tip part of the lower-layer wing by an end plate to form a wing box structure. The trailing edge of the outer wing is a lifting aileron, the spanwise position is 0.7-0.98, the length of the aileron chord is 30% of the local chord length, the maximum deflection angle is 20 degrees, the trailing edge of the wing between the two bodies is a maneuvering flap, and the chord length of the flap is 0.33% of the local chord length and is used for longitudinal pitching operation.

And 5: double-body layout design: the cross section of the middle part of the fuselage is a vertical ellipse, so that the aerodynamic interference of the upper wing and the lower wing is reduced, and the space of the fuselage meets the requirements of the arrangement of a flight control system, an avionic system and a main landing gear cabin; the length of the fuselage is 13.6m, and the rear fuselage is quickly contracted into an empennage stay bar.

Step 6: designing a tail wing with small tail capacity: the plane of the empennage is rectangular, the airfoil adopts NACA64 laminar flow airfoil, and the camber of the airfoil is 0. The volume of the horizontal tail is 0.22, the thickness is 0.11, the installation angle is-2, and the horizontal tail is arranged at the top of the vertical tail. The vertical fin is arranged at the tail part of the rear fuselage, the capacity of the vertical fin is 0.012, the thickness of the vertical fin is 0.15, and the bottom of the vertical fin is 500mm away from the ground for installing the rear landing gear.

And 7: light 4-wheel landing gear design: the main landing gear is positioned at the front part of the fuselage, 0.6MAC before the center of gravity, and is received in a landing gear cabin at the front part of the fuselage. The total length of the main landing gear is 1200mm, and the diameter of a tire is 400 mm; the length of the rear landing gear is 500mm, the diameter of the airplane wheel is 300mm, the rear landing gear is installed at the bottom of the vertical tail, and the rear landing gear can be retracted into the vertical tail after taking off. The take-off and landing modes of the airplane are the same as those of a rear three-point landing gear airplane.

And 8: designing a distributed power system: the power of the airplane comes from 10 electric propellers, the upper wing and the lower wing are respectively 5, the power of each motor is 1.33KW, the efficiency of each propeller is not less than 0.82, and the power control system can enable the propellers on two sides of the airplane to rotate in a differential mode so as to achieve course control.

And step 9: and calculating the weight, aerodynamic force, lift limit and flight time of the airplane according to the design data, and evaluating the performance.

The weight data are as follows: the total weight of the aircraft is 1178kg, the structural weight is 716kg, the system weight is 52kg, the weight of the storage battery is 203kg, the weight of the battery panel is 125kg, the mission load is 82kg, the cruise lift-drag ratio is 31.2, the lift limit is 23800m, the continuous flight time is 62 days, and the design requirements are met.