Solar-powered air-breathing electric propulsion drone

Abstract

The application provides a solar energy suction type electric propulsion unmanned aerial vehicle, which comprises a fuselage, wherein a ring truss type photoelectric conversion device is arranged on the fuselage, the photoelectric conversion device comprises a ring truss unfolding mechanism and a foldable thin film type photoelectric conversion array, the thin film type photoelectric conversion array is supported and connected by the ring truss unfolding mechanism, and the unfolding or folding of the thin film type photoelectric conversion array is realized through the ring truss unfolding mechanism; a thrust device is arranged on the fuselage, and the thrust device adopts a suction type electric propulsion device. The application uses the ring truss type photoelectric conversion device, can realize a large deployment and folding ratio design of the solar energy thin film, has a simple structure and can effectively reduce the structural mass. The application uses the suction type electric propulsion as a power device, does not need to carry additional propellants, can take gas from the surrounding thin atmosphere as the propellants, and can generate corresponding thrust after the gas is ionized and accelerated, so that the unmanned aerial vehicle can work on the orbit for a long time.

Description

TECHNICAL FIELD

[0001] The present application mainly relates to the technical field of unmanned aerial vehicle, and particularly relates to a solar energy powered air-breathing electric propulsion unmanned aerial vehicle. BACKGROUND

[0002] At present, unmanned aerial vehicles have been widely applied in various fields. In recent years, with the development of space technology in the world, major space powers in the world have accelerated the pace of exploring the universe. At present, the most commonly used vehicle for exploring Mars is a Mars rover, but the Mars rover is too slow. The use of unmanned aerial vehicles to explore Mars can explore a wider area on the surface of Mars in a short time. The atmospheric density on the surface of Mars is about one percent of that on the surface of the earth, and the gravity is about 3 / 8 of that on the surface of the earth. American scientists are developing a rotor type Mars helicopter. Theoretically, the rotor of the helicopter can rotate at a speed of more than 2400 revolutions per minute to provide sufficient lift, so that the helicopter can hover in the air for 2-3 minutes and fly a distance of half a kilometer, and then the solar panel charges the power supply to prepare for the next flight.

[0003] Among the existing Mars exploration methods, the Mars rover has poor maneuverability and cannot quickly complete a large-scale exploration task. The height of the on-orbit detector is too high, and it is difficult to accurately observe the ground. Therefore, a flying vehicle that can perform a flight exploration task on the surface of Mars has a significant advantage. At present, only NASA of the United States has proposed a rotor type Mars helicopter. The helicopter relies on the high-speed rotation of the blades to generate lift. Since the atmosphere on the surface of Mars is thin, the lift generated is very small. The helicopter can only reduce the weight as much as possible, so the load carried by the helicopter is limited, and the tasks that can be performed are also limited. SUMMARY

[0004] In view of the technical problems existing in the prior art, the present application provides a solar energy powered air-breathing electric propulsion unmanned aerial vehicle.

[0005] To achieve the above-mentioned purpose, the technical scheme adopted by the present application is as follows:

[0006] A solar energy powered air-breathing electric propulsion unmanned aerial vehicle, comprising a fuselage, wherein a ring truss type photoelectric conversion device is mounted on the fuselage, the photoelectric conversion device comprises a ring truss unfolding mechanism and a foldable thin film type photoelectric conversion array, the thin film type photoelectric conversion array is supported and connected by the ring truss unfolding mechanism, and the unfolding or folding of the thin film type photoelectric conversion array is realized through the ring truss unfolding mechanism; a thrust device is arranged on the fuselage, and the thrust device adopts an air-breathing electric propulsion device.

[0007] As a preferred scheme, the fuselage of the present application adopts a flying wing type aerodynamic layout fuselage.

[0008] As a preferred scheme, the annular truss unfolding mechanism comprises a foldable annular outer frame, a coupling shaft is arranged at the center of the annular outer frame, the coupling shaft is drivingly connected with a power mechanism carried on the fuselage, a plurality of longitudinal rods are supported on the coupling shaft and between the coupling shaft and the annular outer frame, a thin-film photoelectric conversion array is arranged between the longitudinal rods, and the thin-film photoelectric conversion array and the annular outer frame are unfolded or folded by driving the coupling shaft and the longitudinal rods by the power mechanism.

[0009] As a preferred scheme, the annular outer frame is a double-layer structure, each layer of the annular outer frame is formed by a plurality of transverse rods connected in a head-to-tail manner through self-locking hinges, and the double-layer annular outer frames are rotatably connected through linkage shafts at every same or different interval between the self-locking hinges, so that the double-layer annular outer frames are connected and synchronously linked to be unfolded or folded.

[0010] As a preferred scheme, the fuselage and the annular truss unfolding mechanism are provided with an unfolding mechanical arm for supporting the annular truss unfolding mechanism at a certain height at a certain angle.

[0011] As a preferred scheme, the fuselage is provided with a plurality of air-breathing electric propulsion devices, the unmanned aerial vehicle is driven by the thrust generated by the air-breathing electric propulsion devices, and the long-time flight of the unmanned aerial vehicle is ensured.

[0012] As a preferred scheme, the fuselage is provided with an air inlet unit, the air inlet unit is connected with a gas guide channel in the fuselage, and the captured gas is sent to each air-breathing electric propulsion device through the gas guide channel.

[0013] As a preferred scheme, the air-breathing electric propulsion device comprises a cooling flow channel, an ionization chamber and a nozzle, the ionization chamber is connected with the nozzle, the cooling flow channel is arranged around the ionization chamber, the gas taken by the air inlet unit is divided into two parts, one part of the gas directly enters the ionization chamber for ionization, and the other part of the gas enters the cooling flow channel through the cooling flow channel inlet, the cooling flow channel is connected with the ionization chamber through the cooling flow channel outlet, the gas entering the cooling flow channel through the cooling flow channel inlet cools the ionization chamber and then enters the ionization chamber through the cooling flow channel outlet for ionization, and the two gas streams entering the ionization chamber are accelerated through the nozzle, so that the thruster generates thrust.

[0014] As a preferred scheme, the air-breathing electric propulsion device is a coaxial structure.

[0015] As a preferred scheme, the air inlet channel inlet of the air inlet unit is a honeycomb structure formed by a plurality of air inlet holes arranged in a honeycomb shape.

[0016] As a preferred solution, the air inlet unit of the application is trumpet-shaped and its inner side wall is coated with aluminized reflective material, and the air inlet unit is gradually tapered in a parabolic cross-section along the direction of the air flow.

[0017] As a preferred solution, the ionization chamber and the nozzle of the application are designed in a split type or in an integrated type.

[0018] As a preferred solution, the RF coil on the outer side of the ionization chamber is located in the cooling flow channel, and the electromagnetic coil on the outer side of the rear section of the ionization chamber and the outer side of the front section of the nozzle is located in the cooling flow channel, and the gas flowing through the cooling flow channel can be fully cooled.

[0019] Compared with the prior art, the technical effects of the application are:

[0020] The application uses a ring truss type photoelectric conversion device, which can realize large expansion and contraction ratio design of solar thin film, and has simple structure and can effectively reduce the structural weight.

[0021] The application uses an air-breathing electric propulsion as a power device, does not need to carry additional propellants, can take gas from the surrounding thin atmosphere as a propellant, and can generate corresponding thrust after ionization and acceleration of the gas, so that the unmanned aerial vehicle can work on orbit for a long time.

[0022] Further, the application has a flying wing type aerodynamic shape layout, improves the overall aerodynamic efficiency of the unmanned aerial vehicle, effectively reduces the drag coefficient of the machine body, has a large load installation space, and can perform complex and diverse detection tasks.

[0023] Further, the air inlet unit of the application uses a parabolic honeycomb type air inlet structure design, which has the advantages of high gas capture efficiency and high gas compression ratio.

[0024] Further, the air-breathing electric propulsion device of the application can work in the earth's super-low orbit space, takes gas from the surrounding thin atmosphere environment as a working medium through the air inlet unit, a part of the gas directly enters the ionization chamber for acceleration, another part of the gas passes through the cooling flow channel to cool the thruster, and then the part of the gas enters the ionization chamber for ionization and acceleration, the electric thruster with the self-cooling device can ensure long-term stable on-orbit work, and has the advantages of simple structure and low cost. BRIEF DESCRIPTION OF DRAWINGS

[0025] In order to more clearly illustrate the technical solutions in the embodiments of the application or the prior art, the drawings needed to be used in the embodiments or the prior art description will be briefly introduced below. Obviously, the drawings in the following description are only some embodiments of the application, and other drawings can be obtained by those skilled in the art without creative labor on the basis of the drawings shown.

[0026] Figure 1 is a structural schematic diagram of an embodiment of the present application;

[0027] Figure 2 is a structural schematic diagram of an embodiment of the present application;

[0028] Figure 3 is a structural schematic diagram of a fuselage of an embodiment of the present application;

[0029] Figure 4 is a structural schematic diagram of an air intake unit of an embodiment of the present application;

[0030] Figure 5 is a schematic diagram of an air-breathing electric propulsion device of an embodiment of the present application;

[0031] Figure 6 is a structural schematic diagram of a photoelectric conversion device of an embodiment of the present application;

[0032] Figure 7 is a schematic diagram of an unfolding process of a photoelectric conversion device of an embodiment of the present application;

[0033] Figure 8 is a diagram of the variation of Knudsen number at different heights on the surface of Mars in an embodiment of the present application;

[0034] Figure 9 is a diagram of the variation of drag of a vehicle at different heights in an embodiment of the present application;

[0035] Figure 10 is a structural schematic diagram of an air-breathing electric propulsion device of an embodiment of the present application;

[0036] Figure 11 is Figure 10 a diagram of the flow of air in the embodiment shown;

[0037] The diagram is annotated as follows:

[0038] 100, fuselage;

[0039] 200, photoelectric conversion device; 201, ring truss unfolding mechanism; 202, thin-film photoelectric conversion array; 203, ring outer frame; 204, coupling; 205, vertical rod; 206, horizontal rod; 207, self-locking hinge; 208, linkage shaft; 209, unfolding mechanical arm;

[0040] 300, thrust device; 301, air intake unit; 302, air guide; 303, cooling flow channel; 304, ionization chamber; 305, nozzle; 306, cooling flow channel outlet; 307, air intake hole; 308, radio frequency coil; 309, electromagnetic coil; 310, cooling flow guide air channel; 311, main air guide; 312, cooling flow guide air channel shell; 313, thruster shell. DETAILED DESCRIPTION

[0041] The technical solutions in the embodiments of the present application will be clearly and completely described below with reference to the drawings in the embodiments of the present application. Obviously, the described embodiments are only part of the embodiments of the present application, rather than all the embodiments of the present application. Based on the embodiments in the present application, all other embodiments obtained by a person of ordinary skill in the art without creative effort are within the protection scope of the present application.

[0042] It should be noted that all directional indications, such as up, down, left, right, front, back, etc., in the embodiments of the present application are only used to explain the relative position relationship, movement condition, etc. between components in a certain posture (as shown in the drawings), and if the certain posture changes, the directional indications also change accordingly.

[0043] In addition, the description such as "first", "second" and the like in the present application is only for the purpose of description, and cannot be understood as indicating or implying the relative importance of the indicated technical features or implicitly indicating the number of the indicated technical features. Therefore, the features defined as "first", "second" can explicitly or implicitly include at least one of the features. In the description of the present application, the meaning of "plurality" is at least two, for example, two, three, etc., unless otherwise explicitly specified.

[0044] In the present application, unless otherwise explicitly specified and limited, the terms "connection", "fixing" and the like should be understood in a broad sense, for example, "fixing" can be fixed connection, or detachable connection, or integral; can be mechanical connection, or electrical connection, or physical connection, or wireless communication connection; can be direct connection, or indirect connection through an intermediate medium, or the internal communication of two elements or the interaction relationship between two elements, unless otherwise explicitly limited. For a person of ordinary skill in the art, the specific meaning of the above terms in the present application can be understood according to the specific circumstances.

[0045] In addition, the technical solutions of each embodiment of the present application can be combined with each other, but it must be based on the fact that a person of ordinary skill in the art can realize it, and when the combination of technical solutions appears contradictory or unachievable, it should be considered that the combination of technical solutions does not exist, and is not within the protection scope of the present application.

[0046] Reference Figure 1 , Figure 2In an embodiment of the present application, a solar-powered air-breathing electric propulsion unmanned aerial vehicle is provided, comprising a fuselage 100, wherein a ring truss type photoelectric conversion device 200 is mounted on the fuselage 100, the photoelectric conversion device 200 comprises a ring truss unfolding mechanism 201 and a foldable thin film type photoelectric conversion array 202, the thin film type photoelectric conversion array 202 is supported and connected by the ring truss unfolding mechanism 201, and the unfolding or folding of the thin film type photoelectric conversion array 202 is realized through the ring truss unfolding mechanism 201; the fuselage 100 is provided with a thrust device 300, and the thrust device 300 adopts an air-breathing electric propulsion device. The fuselage 100 is mounted with a plurality of air-breathing electric propulsion devices, and the air-breathing electric propulsion device can generate thrust to drive the unmanned aerial vehicle without carrying additional fuel, thereby ensuring long-time flight of the unmanned aerial vehicle to perform tasks.

[0047] As a preferred embodiment, refer to Figure 3 In this embodiment, the fuselage 100 adopts a flying wing type aerodynamic layout fuselage, and four thrust devices 300 are arranged at the rear end of the fuselage 100. The air inlet unit 301 is connected to the air guide 302 in the fuselage 100, and the captured gas is sent to each air-breathing electric propulsion device through the air guide 302.

[0048] An air inlet unit (not shown in the figure) is designed at the front end of the flying wing type fuselage, and the specific position of the air inlet unit can be flexibly arranged and adjusted by those skilled in the art according to the prior art, and the structure of the air inlet unit is not limited. The gas captured by the air inlet unit can be transported to the four thrust devices 300 through the air guide in the fuselage 100, and various payloads can be installed in the middle part of the fuselage 100. The flying wing type aerodynamic layout improves the aerodynamic efficiency of the aircraft body, reduces the aerodynamic drag coefficient of the aircraft body, and at the same time provides sufficient space for placing the payload, thereby facilitating the performance of various detection tasks.

[0049] As a preferred embodiment, the air inlet unit 301 is a parabolic honeycomb structure, which can efficiently capture and compress the incoming flow. Refer to Figure 4 In this embodiment, the air inlet of the air inlet unit 301 is a honeycomb structure, which is formed by a plurality of air inlet holes 307 arranged in a honeycomb shape. The cross section of the air inlet hole 307 is hexagonal. For an air inlet with a wind area of 1m 2 , the size of the air inlet hole 307 is 10-30mm. Further, the air inlet unit 301 is trumpet-shaped and the inner side wall thereof is coated with an aluminized reflective material, the air inlet unit 301 is gradually tapered in a parabolic shape along the direction of the airflow, and the scale in the axial direction is close to the scale in the radial direction. In a preferred embodiment, the ratio of the scale in the axial direction to the scale in the radial direction of the air inlet unit 301 is 1.2.

[0050] Refer to Figure 5In the embodiment shown, the air-breathing electric propulsion device is coaxial structure, which includes a cooling flow channel 303, an ionization chamber 304 and a nozzle 305, the ionization chamber 304 is communicated with the nozzle 305, the cooling flow channel 303 is arranged on the periphery of the ionization chamber 304, the radio frequency coil 308 outside the ionization chamber 304 is located in the cooling flow channel 303, the electromagnetic coil 309 outside the rear section of the ionization chamber 304 and the front section of the nozzle 305 is located in the cooling flow channel 303, and the gas flowing through the cooling flow channel 303 can be sufficiently cooled. In this way, a part of the gas is cooled in the cooling flow channel 303 before being ionized and accelerated, so as to ensure that the propulsion device can work stably for a long time. The air-breathing electric propulsion device ionizes the gas through the radio frequency coil 308, and then accelerates the gas through the magnetic nozzle, and the ionization chamber 304 is made of quartz glass. In this way, each air-breathing electric propulsion device ionizes the gas through radio frequency, and then the ionized gas is accelerated and sprayed out to obtain thrust under the action of the magnetic nozzle. The radio frequency ionization method has the advantages of low power and high ionization rate, and can ensure high performance of the thruster.

[0051] Specifically, the gas captured by the air inlet unit 301 is sent to each air-breathing electric propulsion device through the gas guide channel 302. The air inlet structure of the air-breathing electric propulsion device divides the delivered gas into two parts, and the air inlet structure of the air-breathing electric propulsion device includes a main gas guide channel 311 and a cooling flow guide channel 310. One part of the gas directly enters the ionization chamber 304 for ionization through the main gas guide channel 311, and the other part of the gas is guided to the cooling flow channel inlet 305 through the cooling flow guide channel 310 to enter the cooling flow channel 303. The cooling flow channel 303 is communicated with the ionization chamber 304 through the cooling flow channel outlet 306. The gas entering the cooling flow channel 303 cools the ionization chamber 304 and then enters the ionization chamber 304 for ionization through the cooling flow channel outlet 306. The two gas streams entering the ionization chamber 304 are ionized and then accelerated through the nozzle 305, so as to make the thruster generate thrust.

[0052] Referring to Figure 6 and Figure 7In this embodiment of the present application, the photoelectric conversion device uses a ring truss type, the structural mass of the photoelectric conversion module is concentrated on the outer ring frame, and the total structural mass does not increase proportionally with the expansion area of the thin film photoelectric conversion array, so that the photoelectric conversion device has the advantages of light weight and large expansion / retraction ratio. Specifically, the ring truss expansion mechanism 201 includes a foldable ring outer frame 203, the ring outer frame 203 is provided with a shaft coupling 204 at the center, the shaft coupling 204 is drivingly connected with a power mechanism (such as a motor carried in the fuselage) carried on the fuselage 100, a plurality of longitudinal rods 205 are supported on the shaft coupling 204 and between the shaft coupling 204 and the ring outer frame 203, and a thin film photoelectric conversion array 202 is mounted between the longitudinal rods 205. The foldable thin film photoelectric conversion array 202 is connected with the ring outer frame 203 through a hinge, the shaft coupling 204 is driven by the power mechanism to drive the longitudinal rods 205, the ring outer frame 203 can be expanded under the rotary drive of the power mechanism, and the thin film photoelectric conversion array 202 is expanded according to the crease, and the area can be adjusted as needed. When folding is needed, the motor mechanism is controlled to reverse the action. In this way, the rotary expansion or folding of the thin film photoelectric conversion array 202 and the ring outer frame 203 is realized.

[0053] The ring outer frame 203 is a double-layer structure, each layer of the ring outer frame 203 is formed by a plurality of transverse rods 206 connected end to end through self-locking hinges 207, and the double-layer ring outer frame 203 is rotatably connected through linkage shafts 208 at every same or different interval between the self-locking hinges 207, so as to realize the connection and synchronous linkage rotation expansion or folding of the double-layer ring outer frame 203.

[0054] The fuselage 100 and the ring truss expansion mechanism 201 are provided with an expansion mechanical arm 209 for supporting the ring truss expansion mechanism 201 at a certain angle and height.

[0055] Figure 8 is a diagram of the variation of Knudsen number at different heights on the surface of Mars in an embodiment of the present application; Figure 9 is a diagram of the variation of resistance of the aircraft at different heights in an embodiment of the present application; the flight height of the unmanned aerial vehicle is between 80-90km, from Figure 8 It can be seen that the atmosphere below 90km height on the surface of Mars is still a continuous flow, and aerodynamic force can be generated, from Figure 8 It can be seen that at 90km height, the aerodynamic resistance of the aircraft with a windward area of 1m2 is about 2.8mN.

[0056] Figure 10 is a structural principle diagram of the air-breathing electric propulsion device in an embodiment of the present application, Figure 11 is Figure 10The air flow direction chart of the embodiment shows that the air-breathing electric propulsion device comprises a cooling flow channel 303, an ionization chamber 304 and a nozzle 305, the ionization chamber 304 is communicated with the nozzle 305, and the cooling flow channel 303 is arranged at the periphery of the ionization chamber 304. The air inlet structure of the air-breathing electric propulsion device divides the delivered gas into two parts, and the air inlet structure of the air-breathing electric propulsion device comprises a main guide channel 311 and a cooling flow guide channel 310, one part of the gas directly enters the ionization chamber 304 for ionization through the main guide channel 311, and the other part of the gas is guided to the cooling flow channel inlet 305 into the cooling flow channel 303 for cooling the high-temperature part of the propulsion device, and then the part of the gas is combined with the gas taken by the main guide channel 311 to enter the ionization chamber 500 for ionization and acceleration, and the cooling design has the advantages of simple structure and high energy utilization rate.

[0057] Specifically, the cooling flow channel 303 is communicated with the ionization chamber 304 through a cooling flow channel outlet 306, the gas entering the cooling flow channel 303 through the cooling flow channel inlet 305 cools the ionization chamber 304, and then enters the ionization chamber 304 for ionization through the cooling flow channel outlet 306, and the two gas streams entering the ionization chamber 304 are ionized and then accelerated through the nozzle 305, so that the thruster generates thrust.

[0058] The ionization chamber 304 and the nozzle 305 of the present application adopt an assembled split design or an integrated design. The ionization chamber 304 and the nozzle 305 of the present application can adopt an assembled split design, that is, the ionization chamber 304 and the nozzle 305 are separately manufactured and then assembled together. The ionization chamber 304 and the nozzle 305 of the present application can also adopt an integrated design. The material of the ionization chamber 304 and the nozzle 305 is preferably quartz glass, and the integrated design of the ionization chamber 304 and the nozzle 305 does not need to design a connecting structure between the two, has good integrity, and stable and reliable performance. The nozzle 305 of the present application is a horn-shaped nozzle, and the nozzle flow channel area continuously expands in the direction of the gas flow. In an embodiment, the expansion ratio of the nozzle 305 is 2.

[0059] The cooling flow duct 310 is an annular channel between the cooling flow duct shell 312 and the thruster shell 313, the cooling flow duct shell 312 encloses the main flow duct 311, the ionization chamber 304 and the nozzle 305 inside, and the annular space between the cooling flow duct shell 312 and the inner side of the main flow duct outer wall, the ionization chamber outer wall and the nozzle outer wall forms the cooling flow channel 303. A plurality of annularly distributed air guide holes are opened on the cooling flow duct shell 312 near the end of the nozzle 305, which serve as the cooling flow channel inlet 305. A plurality of annularly distributed air guide holes are opened on the main flow duct outer wall in the cooling flow channel 303 or / and on the ionization chamber outer wall near the end of the main flow duct, which serve as the cooling flow channel outlet 306. The cooling flow duct shell 312 is preferably made of low-density composite material.

[0060] The particularity of the low-orbit atmospheric environment is an important consideration for the design of air-breathing electric propulsion systems. This project selects The Mars Global Ionosphere-Thermosphere Model, which combines the GITM framework on Earth with basic physical parameters on Mars, ion-neutral chemistry, and key radiation processes to obtain basic observational characteristics of the thermal, compositional, and dynamical structure of the Martian atmosphere from the surface to the outer atmosphere (0-250 km). Includes low, middle, and high atmospheric processes, partly based on formulas used in previous GCMs of the low and high Martian atmosphere. This allows the M-GITM code to run under various seasons, solar cycles, and dust conditions.

[0061] The model provides information on atmospheric composition, density, temperature, and other parameters as a function of altitude, time, latitude and longitude, and solar and geomagnetic activity. Among them, solar and geomagnetic activity is defined by two indicators: F10.7 solar flux (solar radio flux at a wavelength of 10.7 cm) and geomagnetic index Ap. Formula 1 shows the relationship between the number density of the composition of the Martian atmosphere and the altitude function under the average solar activity F10.7=F10.7avg=140 and geomagnetic conditions Ap=15.

[0062] The orbital working height of the system is 70-90 km, and typical rarefied air data can be obtained by formula 2, and by calculating the Knudsen number, the atmospheric flow state can be judged:

[0063]

[0064] Where λ is the mean free path of gas molecules, d is the average diameter of particles, n is the particle density, and L is the characteristic size, which is set to 10 m, 15 m, and 20 m in this study.

[0065] When Kn<0.1, the flow regime can be considered as continuous flow; when Kn>10, the flow regime can be considered as free molecular flow; the flow regime between the two is considered as transitional flow.

[0066] The inlet generates shock wave effect in continuous flow gas environment, compressing the incoming gas flow; in free molecular flow gas environment, it uses parabolic reflection focusing effect to compress the incoming gas particles; in the above two effects of transition flow state. The preliminarily compressed gas enters the working fluid tank or thruster to ensure that the system can obtain sufficient propellant in any atmospheric environment. As shown in Figure 8

[0067] The drag model of air-breathing electric propulsion system is generally simplified as the general orbiting vehicle drag calculation model, and the drag formula is calculated as:

[0068]

[0069] Where, ρ is the density of the thin atmosphere at the orbit height, CD is the drag coefficient, A is the cross-sectional area of the vehicle, and v is the vehicle speed. Research shows that below 450 km orbit, the drag coefficient is in the range of 2.2±5%, and the inlet structure has little effect on the drag coefficient. The higher the orbit height, the greater the drag coefficient due to the change of atmospheric composition and the change of vehicle speed and particle thermal speed. In this study, the drag coefficient is 2.2.

[0070] The above formula is based on the continuous flow assumption of fluid mechanics, and there may be errors for free molecular flow in super low orbit application. At present, there is no mature super low orbit satellite drag calculation model applied to free molecular flow atmospheric environment. At present, another model that may be more suitable for air-breathing electric propulsion drag calculation is based on the statistical model of rarefied gas dynamics, as follows:

[0071]

[0072] Where, S is the molecular velocity, σ is the reflection coefficient (σ=1 when the wall reflection is completely diffuse, and σ=0 when it is completely specular), Tr is the reflected particle temperature, Ta is the atmospheric temperature, θ is the vehicle attack angle, erf is the error function, and R is the gas constant The statistical method is more suitable for the study of free molecular flow, and there is sufficient reason to believe that the super low orbit vehicle drag calculation under the statistical model is more accurate.

[0073] In this study, the vehicle attack angle θ is simplified to 0; the reflected particle temperature T r is related to the reflection coefficient, and when it is specular, it is T a ​Same (no energy loss); the aircraft wind area A is set to 0.3m 2 , 0.5m 2 and 1m 2 .

[0074] As Figure 9 shown, if the reflection state is full mirror reflection: when the orbit height is 70km, the system with a wind area of 0.3m2 receives a resistance of 11.5mN; when the orbit height is 90km, the system with a wind area of 0.3m2 receives a resistance of 0.84mN; when the orbit height is 70km, the system with a wind area of 1m2 receives a resistance of 38.3mN; when the orbit height is 90km, the system with a wind area of 1m2 receives a resistance of 2.8mN.

[0075] The air-breathing electric propulsion device in the above embodiment does not need to carry propellant itself, can take gas from the surrounding thin atmosphere as propellant, and can work on orbit for a long time. In particular, the self-cooling structure designed can fully cool the high-temperature part of the air-breathing electric propulsion device, increase the working stability and service life of the air-breathing electric propulsion device, and improve the energy utilization rate of the air-breathing electric propulsion device.

[0076] The details of the present application are known.

[0077] The technical features of the above embodiments can be combined arbitrarily, and in order to make the description simple, all possible combinations of the technical features in the above embodiments are not described, however, as long as the combination of the technical features does not exist contradictory, it should be considered that it is within the scope of the present application.

[0078] The above-described embodiments only express several implementation manners of the present application, and the description is more specific and detailed, but it should not be understood as a limitation on the scope of the patent. It should be pointed out that for ordinary skilled in the art, without departing from the concept of the present application, a number of modifications and improvements can be made, which are within the scope of the present application. Therefore, the protection scope of the patent of the present application should be subject to the appended claims.

Claims

1. A solar powered air-breathing electric propulsion drone, characterized in that, The unmanned aerial vehicle comprises a fuselage, a ring-shaped truss type photoelectric conversion device mounted on the fuselage, a ring-shaped truss unfolding mechanism and a foldable thin film type photoelectric conversion array supported and connected by the ring-shaped truss unfolding mechanism, and the unfolding or folding of the thin film type photoelectric conversion array is realized through the ring-shaped truss unfolding mechanism.

2. The solar-powered, air-breathing electric-propulsion drone of claim 1, wherein The fuselage adopts a flying wing type aerodynamic layout.

3. The solar-powered, air-breathing electric-propulsion drone of claim 1 or 2, wherein, The ring-shaped outer frame is a double-layer structure, each layer of the ring-shaped outer frame is formed by a plurality of cross bars connected end to end through self-locking hinges, and the double-layer ring-shaped outer frames are rotationally connected through linkage shafts at every same or different interval self-locking hinges, so as to realize the connection and synchronous linkage rotation, unfolding or folding of the double-layer ring-shaped outer frames.

4. The solar-powered, air-breathing electric-propulsion drone of claim 3, wherein, An unfolding mechanical arm is arranged between the fuselage and the ring-shaped truss unfolding mechanism, and is used for supporting the ring-shaped truss unfolding mechanism at a certain height at a certain angle.

5. The solar powered air-breathing electric propulsion drone of claim 1, wherein, A plurality of air-breathing electric propulsion devices are mounted on the fuselage, and the air-breathing electric propulsion devices generate thrust to drive the unmanned aerial vehicle, so as to ensure long-time flight of the unmanned aerial vehicle.

6. The solar-powered, air-breathing electric-propulsion drone of claim 5, wherein, An air inlet unit is arranged at the front end of the fuselage, the air inlet unit is communicated with a gas guide channel in the fuselage, and the captured gas is sent to each air-breathing electric propulsion device through the gas guide channel.

7. The solar-powered, air-breathing electric-propulsion drone of claim 6, wherein, The air-breathing electric propulsion device comprises a cooling flow channel, an ionization chamber and a nozzle, the ionization chamber is communicated with the nozzle, the cooling flow channel is arranged on the periphery of the ionization chamber, the gas taken by the air inlet unit is divided into two parts, one part of the gas directly enters the ionization chamber for ionization, and the other part of the gas enters the cooling flow channel through the cooling flow channel inlet, the cooling flow channel is communicated with the ionization chamber through the cooling flow channel outlet, the gas entering the cooling flow channel through the cooling flow channel inlet cools the ionization chamber, and then enters the ionization chamber through the cooling flow channel outlet for ionization, and the two gas streams entering the ionization chamber are accelerated through the nozzle, so that the thruster generates thrust.

8. The solar-powered, air-breathing electric-propulsion drone of claim 7, wherein, The air-breathing electric propulsion device is a coaxial structure.

9. The solar-powered, air-breathing electric-propulsion drone of claim 8, wherein, The inlet of the air inlet unit is a honeycomb structure formed by a plurality of air inlet holes arranged in a honeycomb shape.

10. The solar-powered, air-breathing electric-propulsion drone of claim 9, wherein, The air inlet unit is in the shape of a horn, and an aluminum-coated reflective material is coated on the inner side wall of the air inlet unit, and the air inlet unit is in a parabolic cross-sectional tapered structure along the direction of the air flow.

11. The solar-powered, air-breathing, electric-propulsion drone of claim 7 or 8 or 9 or 10, wherein, The ionization chamber and the nozzle adopt an assembled split design or an integrated design.

12. The solar-powered, air-breathing electric-propulsion drone of claim 11, wherein, The radio frequency coil outside the ionization chamber is located in the cooling flow channel, the electromagnetic coil outside the rear section of the ionization chamber and the front section of the nozzle is located in the cooling flow channel, and the gas flowing through the cooling flow channel can be fully cooled.

Images