CN117326101A - Pneumatic heat protection structure of ultra-low orbit aircraft solar cell array - Google Patents
Pneumatic heat protection structure of ultra-low orbit aircraft solar cell array Download PDFInfo
- Publication number
- CN117326101A CN117326101A CN202311372772.3A CN202311372772A CN117326101A CN 117326101 A CN117326101 A CN 117326101A CN 202311372772 A CN202311372772 A CN 202311372772A CN 117326101 A CN117326101 A CN 117326101A
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- Prior art keywords
- solar cell
- cell array
- ultra
- temperature
- protection structure
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Links
- 239000000758 substrate Substances 0.000 claims abstract description 22
- 239000012774 insulation material Substances 0.000 claims description 20
- 238000002955 isolation Methods 0.000 claims description 12
- 239000011810 insulating material Substances 0.000 claims description 6
- 229920001721 polyimide Polymers 0.000 claims description 5
- XAGFODPZIPBFFR-UHFFFAOYSA-N aluminium Chemical compound [Al] XAGFODPZIPBFFR-UHFFFAOYSA-N 0.000 claims description 4
- 229910052782 aluminium Inorganic materials 0.000 claims description 4
- 238000009413 insulation Methods 0.000 claims description 4
- 229920000049 Carbon (fiber) Polymers 0.000 claims description 3
- 239000004917 carbon fiber Substances 0.000 claims description 3
- VNWKTOKETHGBQD-UHFFFAOYSA-N methane Chemical compound C VNWKTOKETHGBQD-UHFFFAOYSA-N 0.000 claims description 3
- 239000003292 glue Substances 0.000 claims description 2
- 238000013461 design Methods 0.000 abstract description 24
- 238000000034 method Methods 0.000 abstract description 6
- 238000012360 testing method Methods 0.000 abstract description 4
- 238000012795 verification Methods 0.000 abstract description 4
- 230000000694 effects Effects 0.000 abstract description 2
- 239000000463 material Substances 0.000 description 3
- 230000009286 beneficial effect Effects 0.000 description 2
- 238000010586 diagram Methods 0.000 description 2
- 239000004744 fabric Substances 0.000 description 2
- 239000002994 raw material Substances 0.000 description 2
- 230000006978 adaptation Effects 0.000 description 1
- 239000002313 adhesive film Substances 0.000 description 1
- 230000006835 compression Effects 0.000 description 1
- 238000007906 compression Methods 0.000 description 1
- 230000017525 heat dissipation Effects 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
- 230000004048 modification Effects 0.000 description 1
- 238000004088 simulation Methods 0.000 description 1
- 230000001360 synchronised effect Effects 0.000 description 1
Classifications
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
- B64G1/00—Cosmonautic vehicles
- B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
- B64G1/52—Protection, safety or emergency devices; Survival aids
- B64G1/58—Thermal protection, e.g. heat shields
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
- B64G1/00—Cosmonautic vehicles
- B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
- B64G1/42—Arrangements or adaptations of power supply systems
- B64G1/44—Arrangements or adaptations of power supply systems using radiation, e.g. deployable solar arrays
- B64G1/443—Photovoltaic cell arrays
Abstract
The invention belongs to the technical field of satellite heat protection, and particularly relates to a pneumatic heat protection structure of a solar cell array of an ultra-low orbit aircraft. And (3) carrying out a substrate protection design and a single battery protection design on the front surface of the solar cell array, and carrying out an electric connector and grounding resistor protection design and a diode protection design on the back surface of the solar cell array. Through ground test verification and on-orbit verification, the protection effect of the patent design on the front and the back of the solar cell array is remarkable, the solar cell array can be effectively protected at a higher temperature caused by pneumatic friction in the process of 160 km-120 km derailment, and the solar cell array can be ensured to work normally in the severe environment of an ultralow orbit.
Description
Technical Field
The invention belongs to the technical field of satellite heat protection, and particularly relates to a pneumatic heat protection structure suitable for a solar cell array of an ultra-low orbit aircraft.
Background
A test aircraft of a certain model runs in a 120-280 km solar synchronous orbit, and the design life of a satellite is 1 year, so that the test aircraft is the lowest running orbit in all satellite models known in China at present. Wherein pneumatic friction exists in the process of 160 km-120 km derailment, and the pneumatic friction can cause the temperature of the protrusion of the star to rise (the maximum temperature of the protrusion can reach 300 ℃ in the windward direction) so as to influence the normal operation of the solar cell array. Therefore, special protection design is carried out for the solar cell array heat influence in the track lowering process.
1. Thermal environment of windsurfing area
According to the simulation result, under the action of 120km aerodynamic heat, the surface temperature of the positive windward area of the sailboard and the protrusions with the surface of 6mm can reach more than 300 ℃. The pneumatic heat flows and wall temperatures of different track heights are shown in table 1.
TABLE 1
2. Thermal environment of windsurfing area with non-frontal side surface of sailboard
The thermal environment summary of the windsurfing areas is shown in Table 2:
TABLE 2
The conditions of the temperature profile of the windsurfing board of fig. 1 are: the plane of the sailboard is parallel to the incoming flow direction, the sailboard is not deflected, the pneumatic heat of the side face of the low rail is 100W/m < 2 >, the maximum temperature of the body-mounted sailboard is 121 ℃, and the maximum temperature of the unfolded sailboard is 85 ℃.
3. Solar cell array element and raw material heat-resistant temperature
The heat resistance temperature statistics of the solar cell array element and the raw materials are shown in table 3.
TABLE 3 Table 3
The heat resistance temperature statistics of the polyimide film and the adhesive film of the substrate are shown in Table 4.
TABLE 4 Table 4
Disclosure of Invention
Aiming at the problems existing in the prior art, the pneumatic heat protection structure of the solar cell array of the ultra-low orbit aircraft is provided, and special protection design is carried out aiming at the heat influence of the solar cell array in the orbit descending process, wherein pneumatic friction is considered in the orbit descending process of 160 km-120 km, and the pneumatic friction can cause the temperature of a protrusion of a star to rise (the maximum windward temperature can reach 300 ℃), so that the normal operation of the solar cell array is influenced.
The technical scheme of the invention is as follows: the utility model provides a pneumatic thermal protection structure of ultralow orbit aircraft solar cell array, all installs the arc weather strip that can cover the solar cell substrate windward side completely at solar cell substrate windward side, the weather strip back is connected with solar cell substrate windward side through the heat insulating column, weather strip openly windward side is provided with high temperature resistant multilayer insulating material.
Further, the wind shielding strip is a carbon fiber wind shielding strip.
Further, the electrical connectors and resistors on the solar array panel are covered and insulated with a high temperature resistant multilayer insulating material.
Further, polyimide films are arranged on the outer surfaces of the high-temperature-resistant multilayer heat insulation materials covered on the electric connector and the resistor.
Further, high-temperature-resistant multilayer heat insulation materials are adhered to +Z and +X directions of the isolated diode module on the back surface of the solar cell substrate, and are adhered to the surrounding area of the diode module through high-temperature glue after being supported by the L-shaped aluminum supporting blocks and are grounded with the star in a high-resistance manner.
Further, the pasting height of the high-temperature-resistant multilayer heat insulation material pasted around the isolation diode module on the X-wing is not less than 12mm, and the length is the length of the isolation diode module.
Further, the pasting height of the high-temperature-resistant multilayer heat insulation material pasted around the isolation diode module on the +X wing is not less than 16mm, and the length is the length of the isolation diode module.
Further, the thickness of the high-temperature-resistant multilayer heat insulation material is 5.5-6.5mm.
Furthermore, when the single batteries of the solar cell array are distributed, no piece is distributed in the area of the front pressing point phi 60mm of the solar cell substrate, and the distance between the front pressing point phi and the edge of the hinge is 20 mm.
The beneficial effects of the invention are as follows: a pneumatic thermal protection structure suitable for a solar cell array of an ultra-low orbit aircraft is provided. And (3) carrying out a substrate protection design and a single battery protection design on the front surface of the solar cell array, and carrying out an electric connector and grounding resistor protection design and a diode protection design on the back surface of the solar cell array. Through ground test verification and on-orbit verification, the protection effect of the patent design on the front and the back of the solar cell array is remarkable, the solar cell array can be effectively protected at a higher temperature caused by pneumatic friction in the process of 160 km-120 km derailment, and the solar cell array can be ensured to work normally in the severe environment of an ultralow orbit.
Drawings
FIG. 1 is a diagram of a temperature profile of a windsurfing board in the prior art;
FIG. 2 is a schematic diagram of a solar array weather strip;
FIG. 3 is a thermal protection design of a solar array connector;
FIG. 4 is a rear layout view of a solar array;
FIG. 5 is a thermal protection design of a solar array diode;
FIG. 6 is a thermal protection design of a solar array diode;
fig. 7 is a front layout view of a solar array.
In the figure: 1 is a solar cell array, 2 is a heat insulation column, 3 is a wind shielding strip, 4 is a high-temperature-resistant multilayer heat insulation material, 5 is an on-board electric connector and a grounding resistor, 6 is a polyimide film, 7 is a diode, 8 is an X wing, 9 is a body plate, 10 is a +X wing, 11 is an L-shaped aluminum supporting block, 12 is a compression point, and 13 is a hinge.
Detailed Description
The invention is further described below with reference to the accompanying drawings.
1. Solar cell array 1 substrate protection design
As shown in fig. 2, arc-shaped wind shielding strips 3 capable of completely covering the windward side of the solar cell substrate (relative to the flying direction of the satellite) are arranged on the windward side of the solar cell substrate to prevent the substrate from being damaged by heat, the back sides of the wind shielding strips are connected with the windward side of the solar cell substrate through heat insulation columns 2, and the windward side of the front side of the wind shielding strips is provided with high-temperature-resistant multilayer heat insulation materials 4. The wind shielding strip is a carbon fiber wind shielding strip.
2. Solar cell array panel upper electric connector and ground resistance protection design
As shown in fig. 3, the on-board electrical connector and the ground resistor 5 are directly covered with a multi-layered insulating material for heat insulation. The outer surfaces of the high-temperature-resistant multilayer heat insulation material covered on the electric connector and the resistor are provided with polyimide films 6.
3. Solar cell array diode protection design
Because the diode 7 heats during operation, and the base temperature of the substrate under the influence of aerodynamic heat is superimposed, the protection design of the diode is not suitable for a mode of covering and insulating by using a plurality of layers of heat insulating materials, and the design is not beneficial to the heat dissipation of the diode. In this way, it is proposed to adhere a plurality of layers of heat insulating material around the isolation diode on the back surface of the substrate to shield the aerodynamic heat on the windward side, and to ensure that the isolation diode is not affected by the aerodynamic heat even when the attitude is deflected by 6 °.
The diode module mounting area and dimensions are shown in fig. 4.
Since the satellite flying gesture is in +X or +Z direction, the aerodynamic heat of the diode in +Z and +X directions is protected by pasting multiple layers of materials.
Considering the worst satellite attitude, under the condition that the attitude deflects by 6 degrees, calculating the shielding height of the needed multilayer materials according to the size of the diode module, and pasting C-shaped multilayer heat insulation materials around the diode module.
A plurality of layers of heat insulation materials with the height not lower than 12mm are adhered on the X wing 8, the length is the length of an isolation diode module (-254 mm for the lateral length of an X outer plate, 305mm for the lateral length of an X inner plate and 82mm for the longitudinal length);
and (3) pasting a multi-layer heat insulation material with the height not lower than 16mm on the +X wing 10, wherein the length is the length of the isolation diode module (+X outer plate transverse length is 101mm and 152mm respectively, +X inner plate length is 101mm,101mm and 152mm respectively, and the longitudinal length is 82 mm).
The attachment position is shown in fig. 5.
The thickness of the high-temperature-resistant multi-layer heat insulation material is about 6mm, and the high-temperature-resistant multi-layer heat insulation material is stuck to the surrounding area of the diode module only on the leeward side due to the fact that the temperature of the windward side is too high. The multi-layer material is supported by the L-shaped aluminum support block 11 and then needs to be grounded, and is grounded with the star body in a high resistance way. As shown in fig. 6.
4. Protection design of single battery of solar cell array
The bulges of the front solar cell unit and the lead do not exceed the coverage of the wind shielding strip, and no extra protection design is needed. Only when the design of the cloth piece, the mechanical protrusion on the solar cell substrate is avoided, and the heat conduction is avoided to influence the operation of the solar cell array. The front pressing point is specifically designed to be a non-cloth piece in a region of 12 phi 60mm, and the distance between the front pressing point and the edge of the hinge 13 is 20 mm. See figure 7 for a front layout of a solar array.
The foregoing is merely a preferred embodiment of the present invention and it should be noted that modifications and adaptations to those skilled in the art may be made without departing from the principles of the present invention, which are intended to be comprehended within the scope of the present invention.
Claims (9)
1. A pneumatic heat protection structure of ultra-low orbit aircraft solar cell array, its characterized in that: the solar cell comprises a solar cell substrate, wherein the windward side of the solar cell substrate is provided with arc-shaped wind shielding strips which can completely cover the windward side of the solar cell substrate, the back sides of the wind shielding strips are connected with the windward side of the solar cell substrate through heat insulation columns, and the windward side of the front side of the wind shielding strips is provided with high-temperature-resistant multilayer heat insulation materials.
2. The aerodynamic heat protection structure of an ultra-low orbit aircraft solar cell array according to claim 1, wherein: the wind shielding strip is a carbon fiber wind shielding strip.
3. The aerodynamic heat protection structure of an ultra-low orbit aircraft solar cell array according to claim 1, wherein: and covering and insulating the electric connector and the resistor on the solar cell array panel by using a high-temperature-resistant multilayer insulating material.
4. A pneumatic thermal protection structure for an ultra-low orbit aircraft solar cell array according to claim 3, wherein: polyimide films are arranged on the outer surfaces of the high-temperature-resistant multilayer heat insulation material covered on the electric connector and the resistor.
5. The aerodynamic heat protection structure of an ultra-low orbit aircraft solar cell array according to claim 1, wherein: and pasting high-temperature-resistant multilayer heat insulation materials in the +Z direction and the +X direction of the isolated diode module on the back surface of the solar cell substrate, wherein the high-temperature-resistant multilayer heat insulation materials are pasted in the surrounding area of the diode module through high-temperature glue after being supported by the L-shaped aluminum supporting block and are grounded with the star in a high-resistance manner.
6. The aerodynamic heat protection structure of an ultra-low orbit aircraft solar cell array according to claim 5, wherein: the high-temperature-resistant multilayer heat insulation material stuck around the isolation diode module on the X wing is stuck at a height not less than 12mm, and the length is the length of the isolation diode module.
7. The aerodynamic heat protection structure of an ultra-low orbit aircraft solar cell array according to claim 5, wherein: the high-temperature-resistant multilayer heat insulation material stuck around the isolation diode module on the +X wing is not less than 16mm in sticking height, and the length is the length of the isolation diode module.
8. The aerodynamic thermal protection structure of an ultra-low orbit aircraft solar cell array according to claim 1, 3 or 5, wherein: the thickness of the high-temperature-resistant multilayer heat insulation material is 5.5-6.5mm.
9. The aerodynamic heat protection structure of an ultra-low orbit aircraft solar cell array according to claim 1, wherein: when the single batteries of the solar cell array are distributed, no piece is distributed in the area of 60mm of the front pressing point phi of the solar cell substrate, and the distance between the front pressing point phi and the edge of the hinge is 20 mm.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202311372772.3A CN117326101A (en) | 2023-10-23 | 2023-10-23 | Pneumatic heat protection structure of ultra-low orbit aircraft solar cell array |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202311372772.3A CN117326101A (en) | 2023-10-23 | 2023-10-23 | Pneumatic heat protection structure of ultra-low orbit aircraft solar cell array |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN117326101A true CN117326101A (en) | 2024-01-02 |
Family
ID=89282836
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202311372772.3A Pending CN117326101A (en) | 2023-10-23 | 2023-10-23 | Pneumatic heat protection structure of ultra-low orbit aircraft solar cell array |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN117326101A (en) |
-
2023
- 2023-10-23 CN CN202311372772.3A patent/CN117326101A/en active Pending
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