CN110949696A - Satellite engine thermal protection performance verification system - Google Patents
Satellite engine thermal protection performance verification system Download PDFInfo
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- CN110949696A CN110949696A CN201911142455.6A CN201911142455A CN110949696A CN 110949696 A CN110949696 A CN 110949696A CN 201911142455 A CN201911142455 A CN 201911142455A CN 110949696 A CN110949696 A CN 110949696A
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- thermal protection
- protection performance
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- satellite engine
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
- B64G7/00—Simulating cosmonautic conditions, e.g. for conditioning crews
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- Testing Of Devices, Machine Parts, Or Other Structures Thereof (AREA)
- Testing Of Engines (AREA)
Abstract
The invention provides a satellite engine thermal protection performance verification system in the technical field of satellite transmitter thermal protection, which comprises a space environment simulator, a high-temperature measuring device and a power supply, wherein the space environment simulator is connected with the high-temperature measuring device; the space environment simulator is characterized in that an infrared lamp array, a reflective heat shield and a thermal protection component sample are arranged in the space environment simulator, and the reflective heat shield, the infrared lamp array and the thermal protection component sample are sequentially arranged in parallel; the infrared lamp array is connected with the power supply, and the thermal protection assembly sample is connected with the high-temperature measuring device. The invention provides a test system for verifying the temperature resistance and the heat insulation performance of a thermal protection component of a satellite engine, which can verify the reasonability and the reliability of the structural design of the thermal protection component and improve the working safety of the satellite engine.
Description
Technical Field
The invention relates to the technical field of satellite transmitter thermal protection, in particular to a ground verification system for verifying temperature resistance and heat insulation performance of a satellite engine thermal protection assembly.
Background
With the development of the aerospace industry, the search radius of human beings is increased, the orbit height of the satellite is higher and higher, and the weight requirement of the satellite is more and more severe from the sun synchronization, the earth synchronization to the moon, the mars and the like. On the other hand, the number of engines is increased for maintaining the cruise segment and the orbit of the high-orbit satellite and the deep space exploration satellite, the wall-fixing radiation and the plume temperature are higher than 1000 ℃ when the engines work, the products and parts on the satellite need to be thermally protected, and the weight of the thermal protection is obviously increased due to the increase of the number of the engines.
In order to reduce the weight of the thermal protection of the engine, the structure of the thermal protection has to be optimized, whether the optimized thermal protection assembly can meet the thermal protection requirements of satellite products and components is not enough to eliminate the huge safety risk brought to the satellite by thermal protection failure only by simulation, and the high-altitude thermal test run test performed by adopting a real engine is very expensive.
Therefore, the design of a test system for ground verification of temperature resistance and heat insulation performance of a satellite engine thermal protection assembly meets the reasonability and reliability of the structural design of the thermal protection assembly, and the improvement of the safety of a satellite when a satellite transmitter works is a problem which is currently faced by and urgently needs to be solved by technical personnel in the field.
Disclosure of Invention
Aiming at the defects in the prior art, the invention aims to provide a satellite engine thermal protection performance verification system.
The invention provides a satellite engine thermal protection performance verification system, which comprises a space environment simulator, a high-temperature measuring device and a power supply, wherein the space environment simulator is connected with the high-temperature measuring device;
the space environment simulator is characterized in that an infrared lamp array, a reflective heat shield and a thermal protection component sample are arranged in the space environment simulator, and the reflective heat shield, the infrared lamp array and the thermal protection component sample are sequentially arranged in parallel;
the infrared lamp array is connected with the power supply, and the thermal protection assembly sample is connected with the high-temperature measuring device.
In some embodiments, the spatial simulator vacuum level is 1.33 × 10-3Pa, the surface emissivity of the heat sink is more than 0.9, and the temperature of the heat sink is lower than 100K.
In some embodiments, a camera is mounted within the space simulator.
In some embodiments, the power supply is a power adjustable dc power pack.
In some embodiments, the high temperature radiation temperature range of the infrared lamp array is 300 ℃ to 1100 ℃.
In some embodiments, the infrared lamp array consists of 10 tungsten filament quartz lamps with a power of 500W, with a 30mm spacing between each quartz lamp.
In some embodiments, the reflective heat shield is fabricated from polished stainless steel foil, the reflective heat shield having dimensions of 400mm x 400 mm.
In some embodiments, the reflective heat shield has an emission power greater than 0.9.
In some embodiments, a thermocouple is adhered to the inside of the sample of the thermal protection assembly, and the thermocouple is a K-type thermocouple.
In some embodiments, a floating freezing point is arranged in the space environment simulator, and the floating freezing point is coated with a plurality of layers of heat insulation components.
Compared with the prior art, the invention has the following beneficial effects:
1. the invention can simulate the high-temperature radiation of the engine during working, verify the reasonability and reliability of the design of the thermal protection structure and improve the working safety of the satellite engine.
2. The invention can simulate the high-temperature radiation of different engines, and saves time and test cost.
3. The invention reduces the high-altitude hot test and saves the development cost of the engine thermal protection.
Drawings
Other features, objects and advantages of the invention will become more apparent upon reading of the detailed description of non-limiting embodiments with reference to the following drawings:
FIG. 1 is a schematic diagram of a satellite engine thermal protection performance verification system according to the present invention.
Detailed Description
The present invention will be described in detail with reference to specific examples. The following examples will assist those skilled in the art in further understanding the invention, but are not intended to limit the invention in any way. It should be noted that it would be obvious to those skilled in the art that various changes and modifications can be made without departing from the spirit of the invention. All falling within the scope of the present invention.
The present invention will be described in more detail below with reference to the accompanying drawings, which illustrate embodiments of the invention. This invention may be embodied in many different forms and should not be construed as limited to the embodiments set forth herein.
The invention provides a satellite engine thermal protection performance verification system, as shown in figure 1, comprising a space environment simulator 1, a high-temperature measuring device 4 and a power supply 5;
the vacuum degree of the space simulator 1 is 1.33 multiplied by 10-3Pa, the surface emissivity of the heat sink is more than 0.9, the temperature of the heat sink is lower than 100K, and the space environment simulator ensures that the working environment of the thermal protection assembly sample is basically consistent with the in-orbit environment of the satellite.
Be provided with infrared lamp battle array 2, reflection heat screen 3, thermal protection subassembly sample 6 in the space environment simulator 1, reflection heat screen 3 infrared lamp battle array 2 and thermal protection subassembly sample 6 three are parallel placement in space environment simulator 1 in proper order, and infrared lamp battle array 2 is radiant heating, can simulate the high temperature of engine during operation to keep apart with experimental sample physics, guaranteed experimental security, design placed in the middle of infrared lamp battle array 2 simultaneously, make the heat of infrared lamp battle array 2 partial radiation can reflect to thermal protection subassembly sample 6 through reflection heat screen 3 on, so that the high temperature on simulation thermal protection sample surface.
The infrared lamp array is connected with a ground direct current power supply outside the space environment simulator 1 through a transition plug of the space environment simulator 1. Install high temperature sensor on hot protection component sample 6, its transition plug through the space environment simulator is connected with the thermoscope in the high temperature measuring device 4 on ground, constitutes temperature measuring system, gathers the temperature of hot protection component sample through temperature measuring system, confirms its thermal-insulated and temperature resistance ability to confirm rationality and the reliability of hot protection structural design, this experimental system is ground experimental system simultaneously, can effectively reduce high altitude hot test run test, saves the development cost of engine heat protection.
And a thermocouple is adhered inside the thermal protection component sample 6 and is a K-type thermocouple. And a camera is arranged in the space simulator 1. The heat insulation performance of the thermal protection component is verified through the temperature of the thermocouple, and the temperature resistance performance of the thermal protection component is verified by combining a camera in the space simulator.
The power supply 5 is a power-adjustable direct current power supply set, and the radiation heat of the infrared lamp array 2 is adjusted by adjusting the output power of the power supply, so that the high-temperature radiation temperature range of the infrared lamp array 2 reaches 300-1100 ℃, and the actual radiation temperature range of the satellite engine in the space operation process is maximally reached.
The reflective heat shield 3 is made of polished stainless steel foil, and the emission power of the reflective heat shield is greater than 0.9. The transmitting power of the transmitting heat shield 3 is matched with the heat sink surface emissivity of the space simulator 1, so that the experiment needs are better met, the experiment tends to a true value more, meanwhile, the reflecting heat shield adopts polished stainless steel foil, the high reflectivity of the stainless steel foil is utilized, and the radiant heat of the infrared lamp array is reflected to the surface of a test sample, so that the infrared lamp array can simulate the high temperature of thousands of degrees.
A floating freezing point is arranged in the space environment simulator 1, and a plurality of layers of heat insulation assemblies are coated outside the floating freezing point.
The invention preferably selects the manufacturing methods and corresponding data provided in the following examples:
in the test of the thermal protective member, first, a test specimen having the same structure as the thermal protective member was prepared, and the size of the test specimen of this example was 300mm × 300mm, and 5K-type thermocouples were attached to the inside of the test specimen.
An infrared lamp array is designed according to the size of a sample, the infrared lamp array consists of 10 tungsten filament quartz lamps, the distance is 30mm, the power of each quartz lamp is 500W, and the infrared lamp array is installed on a stainless steel frame.
The high temperature reflecting screen is designed according to the sample size, the reflecting heat shield is made of polished stainless steel foil, the emissivity is more than 0.9, the size is 400mm multiplied by 400mm, and the thickness is 50 μm.
The infrared lamp array is connected with 10 ground direct current power supplies of 120V and 5A outside the simulator through a transition plug of the space environment simulator, and the radiation heat of the infrared lamp array is adjusted by adjusting the output power of the power supplies, so that the high temperature of 700 ℃ on the surface of the test piece is simulated.
5K-type thermocouples pasted in the test piece are connected with a 2700-temperature measuring instrument on the ground through transition plugs of the space environment simulator to form a temperature measuring system, and a floating freezing point is placed in the space environment simulator and coated with a plurality of layers of heat insulation assemblies.
During the test, the environment of the space simulator is set to be that the vacuum degree is 1.33 multiplied by 10-3Pa, the surface emissivity of the heat sink is more than 0.9, and the temperature of the heat sink is lower than 100K.
In conclusion, the high-temperature radiation simulation device can simulate the high-temperature radiation of the engine during working, verify the reasonability and reliability of the design of the thermal protection structure, and improve the working safety of the satellite engine; the invention can simulate the high-temperature radiation of different engines, thereby saving time and test cost; the invention reduces the high-altitude hot test and saves the development cost of the engine thermal protection.
The foregoing description of specific embodiments of the present invention has been presented. It is to be understood that the present invention is not limited to the specific embodiments described above, and that various changes or modifications may be made by one skilled in the art within the scope of the appended claims without departing from the spirit of the invention. The embodiments and features of the embodiments of the present application may be combined with each other arbitrarily without conflict.
Claims (10)
1. A satellite engine thermal protection performance verification system is characterized by comprising a space environment simulator (1), a high-temperature measuring device (4) and a power supply (5);
the space environment simulator (1) is internally provided with an infrared lamp array (2), a reflective heat shield (3) and a thermal protection component sample (6), wherein the reflective heat shield (3), the infrared lamp array (2) and the thermal protection component sample (6) are sequentially arranged in parallel;
the infrared lamp array (2) is connected with the power supply (5), and the thermal protection component sample (6) is connected with the high-temperature measuring device (4).
2. The satellite engine thermal protection performance validation system of claim 1, wherein the system is configured to perform thermal protection performance validationCharacterized in that the vacuum degree of the space simulator (1) is 1.33 multiplied by 10-3Pa, the surface emissivity of the heat sink is more than 0.9, and the temperature of the heat sink is lower than 100K.
3. The satellite engine thermal protection performance verification system according to claim 2, characterized in that a camera is installed in the space simulator (1).
4. The satellite engine thermal protection performance verification system according to claim 1, wherein the power supply (5) is a power adjustable dc power supply pack.
5. The satellite engine thermal protection performance verification system according to claim 4, wherein the high temperature radiation temperature range of the infrared lamp array (2) is 300 ℃ to 1100 ℃.
6. The satellite engine thermal protection performance verification system according to claim 5, wherein the infrared lamp array (2) is composed of 10 tungsten filament quartz lamps with power of 500W, and the distance between every two quartz lamps is less than or equal to 30 mm.
7. The satellite engine thermal protection performance verification system according to claim 1, wherein the reflective heat shield (3) is made of polished stainless steel foil, and the reflective heat shield (3) has dimensions of 400mm x 400 mm.
8. The satellite engine thermal protection performance validation system of claim 7, wherein the emissivity of the reflective heat shield (3) is > 0.9.
9. The satellite engine thermal protection performance verification system according to claim 1, wherein a thermocouple is pasted inside the thermal protection assembly test sample (6), and the thermocouple is a K-type thermocouple.
10. The satellite engine thermal protection performance verification system according to claim 1, wherein a floating freezing point is arranged in the space environment simulator (1), and a multi-layer heat insulation assembly is coated outside the floating freezing point.
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| CN201911142455.6A CN110949696A (en) | 2019-11-20 | 2019-11-20 | Satellite engine thermal protection performance verification system |
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Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN113928604A (en) * | 2021-10-19 | 2022-01-14 | 上海卫星装备研究所 | Device and method for testing performance of high-temperature heat shield of deep space exploration spacecraft |
| CN114212283A (en) * | 2021-12-17 | 2022-03-22 | 重庆哈丁环境试验技术股份有限公司 | Space environment simulation test box |
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| CN104015942A (en) * | 2014-06-16 | 2014-09-03 | 北京卫星环境工程研究所 | Ultrahigh-temperature thermal current simulation system used for spacecraft vacuum thermal test |
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2019
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Patent Citations (4)
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| RU2182105C2 (en) * | 2000-01-17 | 2002-05-10 | Открытое акционерное общество "Ракетно-космическая корпорация "Энергия" им. С.П. Королева" | Method of control of process of simulation of solar illumination of spacecraft by infra-red radiators and system for realization of this method |
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| Publication number | Priority date | Publication date | Assignee | Title |
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| CN113928604A (en) * | 2021-10-19 | 2022-01-14 | 上海卫星装备研究所 | Device and method for testing performance of high-temperature heat shield of deep space exploration spacecraft |
| CN113928604B (en) * | 2021-10-19 | 2024-04-09 | 上海卫星装备研究所 | Device and method for testing performance of high-temperature heat shield of deep space exploration spacecraft |
| CN114212283A (en) * | 2021-12-17 | 2022-03-22 | 重庆哈丁环境试验技术股份有限公司 | Space environment simulation test box |
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