A method for thermal equilibrium testing of CubeSats

Abstract

This invention discloses a method for thermal equilibrium testing of a CubeSat, comprising: thermally modifying a prototype CubeSat; calculating the characteristic point temperatures of the modified prototype CubeSat under high and low temperature conditions; constructing a thermal equilibrium testing system and placing the modified prototype CubeSat within the system; obtaining the thermal equilibrium test boundary temperatures under high and low temperature conditions based on the characteristic point temperatures of the modified prototype CubeSat under high and low temperature conditions, respectively; and conducting a thermal equilibrium test based on the thermal equilibrium test boundary temperatures under high and low temperature conditions. This invention has the advantages of accurate testing, low cost, high reliability, short cycle time, and large-scale production.

Description

Technical Field

[0001] This invention belongs to the field of satellite thermal control technology, and in particular relates to a method for thermal balance testing of CubeSats. Background Technology

[0002] CubeSats, with their advantages of short development cycles, low cost, and flexible launch methods, have become one of the three major hot topics and cutting-edge developments in the field of aerospace technology, especially nanosatellite swarms flying in formations or constellations, which are favored by the international aerospace community. CubeSats have small mass and low heat capacity, while their onboard electronic equipment is highly integrated, resulting in a very high heat flux density. This causes the satellite's temperature level to fluctuate rapidly due to changes in heat flux, posing a considerable challenge to thermal control design. Because of the relatively short design life of CubeSats, and considering the limitations of mass, space, and power, most CubeSats launched so far employ only passive thermal control measures. The relatively low international requirements for the lifespan of CubeSats are also due to the difficulty in designing thermal control systems; therefore, many CubeSats lack a thermal control system altogether.

[0003] Thermal equilibrium testing is an experiment conducted under the orbital thermal environment conditions of a space simulation chamber to verify the temperature distribution under equilibrium conditions during satellite orbital flight. Commonly used thermal equilibrium testing methods and their characteristics are as follows:

[0004] 1) Solar simulator: It uses a xenon lamp as a light source, which emits light with a spectral distribution close to that of sunlight after special filtering.

[0005] The technology is complex and costly, which contradicts the low-cost design philosophy of CubeSats.

[0006] 2) Infrared cage simulation: By installing electric heating belts on the cage, the external heat flow on each side of the satellite is simulated by the infrared radiation of the heating belts.

[0007] The small size of the cube star limits the size of the heating strips, which makes it impossible to meet the requirements for heat flow simulation of the heating zones. Furthermore, the heating interference between different zones is severe, resulting in a large error in the external heat flow simulation.

[0008] 3) Infrared lamp array simulator: It consists of infrared lamps and their supports forming an infrared lamp array. During the test, the lamp array is formed around the star to achieve radiative heating of the satellite surface.

[0009] It is difficult to determine the actual heat flux absorbed by the satellite, and there is the same external heat flux error problem as the infrared cage simulation;

[0010] 4) Surface contact electric heater: A thin film resistance heating element is attached to the surface of the spacecraft. When energized, the Joule heat generated by the element is equal to the heat flow from outside space absorbed by the surface.

[0011] The surface of a CubeSat is mostly body-mounted wings. Electric heating elements cannot be attached to the outer surface of a prototype CubeSat, and due to the dense and compact structure of the equipment on the inner surface, electric heating elements generally cannot be attached either. Summary of the Invention

[0012] The technical problem solved by this invention is to overcome the shortcomings of the prior art and provide a CubeSat thermal balance test method, which has the advantages of accurate testing, low cost, high reliability, short cycle and large batch.

[0013] The objective of this invention is achieved through the following technical solution: a method for thermal equilibrium testing of a CubeSat, comprising: thermally modifying a prototype CubeSat; calculating the characteristic point temperatures of the modified prototype CubeSat under high and low temperature conditions; constructing a thermal equilibrium testing system and placing the modified prototype CubeSat within the thermal equilibrium testing system; obtaining the thermal equilibrium test boundary temperatures under high and low temperature conditions based on the characteristic point temperatures of the modified prototype CubeSat under high and low temperature conditions, respectively; and conducting a thermal equilibrium test based on the thermal equilibrium test boundary temperatures under high and low temperature conditions.

[0014] The above-mentioned CubeSat thermal balance test method includes the following thermal modification of the prototype CubeSat: painting the structural plate frame of the prototype CubeSat; covering the exterior of the internal equipment of the prototype CubeSat with multiple layers of insulation; covering the exterior of the outer cabin plate of the prototype CubeSat with multiple layers of insulation; and attaching electric heaters to the outer surface of the prototype CubeSat.

[0015] In the above-mentioned CubeSat thermal balance test method, based on finite element thermal simulation analysis, the characteristic point temperature of the thermally modified prototype CubeSat under high temperature and low temperature conditions is calculated.

[0016] In the above-mentioned CubeSat thermal balance test method, the maximum external heat flow + maximum operating mode of the prototype CubeSat is the high-temperature condition; the minimum external heat flow + minimum operating mode of the prototype CubeSat is selected as the low-temperature condition.

[0017] In the above-mentioned CubeSat thermal balance test method, the characteristic point temperature is taken as the temperature at the satellite structural plate or non-heat source equipment.

[0018] In the above-mentioned CubeSat thermal balance test method, the thermal balance test system includes a vacuum simulation chamber, a simulation cabin, and a payload platform; wherein, the thermally modified prototype CubeSat is placed on the upper part of the payload platform; the simulation cabin is connected to the payload platform, and the thermally modified prototype CubeSat is located inside the simulation cabin; the simulation cabin is located inside the vacuum simulation chamber.

[0019] In the above-mentioned CubeSat thermal balance test method, the outer surface of the simulation chamber is covered with multiple layers of insulation to control the heat dissipation of the simulation chamber to the vacuum simulation chamber; the carrier platform is equipped with an electric heater, which can heat and control the temperature synchronously with the simulation chamber.

[0020] In the above-mentioned CubeSat thermal equilibrium test method, the thermal equilibrium test boundary temperature under low-temperature conditions is obtained by the following formula:

[0021] Q 2L =Q 3L -Q 1L ;

[0022]

[0023] Among them, Q 2L Q represents the total heat dissipation of satellite equipment under low-temperature operating conditions. 3L Q represents the heat flow radiated from a satellite into space under cryogenic conditions. 1L To allow the satellite to absorb space heat flux under cryogenic conditions, ε0 is the system emissivity, and A S Let T be the surface area of ​​the satellite, σ be the Boltzmann constant, and T be the surface area of ​​the satellite. C,L T represents the boundary temperature for the thermal balance test of the satellite under low-temperature operating conditions. S,L This refers to the boundary temperature of the thermal balance test in the simulation chamber under low-temperature conditions.

[0024] In the above-mentioned CubeSat thermal equilibrium test method, the thermal equilibrium test boundary temperature under high-temperature conditions is obtained by the following formula:

[0025] Q 2H =Q 3H -Q 1H ;

[0026]

[0027] Among them, Q 2H Q represents the total heat dissipation of satellite equipment under high-temperature operating conditions. 3H Q represents the heat flow radiated from a satellite into space under high-temperature operating conditions. 1H To allow the satellite to absorb space heat flux under high-temperature operating conditions, ε0 is the system emissivity, and A S Let T be the surface area of ​​the satellite, σ be the Boltzmann constant, and T be the surface area of ​​the satellite. C,H T represents the boundary temperature for the thermal balance test of the satellite under high-temperature conditions. S,H This refers to the boundary temperature of the thermal balance test in the simulation chamber under high-temperature conditions.

[0028] In the above CubeSat thermal equilibrium test method, the system emissivity ε0 is:

[0029]

[0030] Among them, A S Let A be the surface area of ​​the satellite, σ be the Boltzmann constant, and A be the surface area of ​​the satellite. C To simulate the internal surface area of ​​the cabin, ε S To simulate the average emissivity of the cabin surface area, ε CTo simulate the emissivity of the cabin's internal surface.

[0031] Compared with the prior art, the present invention has the following advantages:

[0032] (1) This invention reduces the production of infrared cages and process plates, thereby reducing the cost of CubeSat production and development;

[0033] (2) In addition to the extra-satellite multilayer, the experimental satellite and the launched satellite are in the same state, which shortens the satellite assembly time during the development process;

[0034] (3) This invention reduces the difficulty of model correction caused by inaccurate simulation of external heat flow in heat balance tests. Attached Figure Description

[0035] Various other advantages and benefits will become apparent to those skilled in the art upon reading the following detailed description of preferred embodiments. The accompanying drawings are for illustrative purposes only and are not intended to limit the invention. Furthermore, the same reference numerals denote the same parts throughout the drawings. In the drawings:

[0036] Figure 1 This is a flowchart of the CubeSat thermal equilibrium test method provided in the embodiments of the present invention;

[0037] Figure 2 This is a schematic diagram of the thermal balance test system provided in an embodiment of the present invention. Detailed Implementation

[0038] Exemplary embodiments of the present disclosure will now be described in more detail with reference to the accompanying drawings. While exemplary embodiments of the present disclosure are shown in the drawings, it should be understood that the present disclosure may be implemented in various forms and should not be limited to the embodiments set forth herein. Rather, these embodiments are provided to enable a more thorough understanding of the present disclosure and to fully convey the scope of the disclosure to those skilled in the art. It should be noted that, unless otherwise specified, the embodiments and features described herein can be combined with each other. The present invention will now be described in detail with reference to the accompanying drawings and embodiments.

[0039] Figure 1 This is a flowchart of the CubeSat thermal equilibrium test method provided in an embodiment of the present invention. Figure 1 As shown, the method includes the following steps:

[0040] Hot modification of the prototype CubeSat;

[0041] Calculate the characteristic point temperatures of the thermally modified prototype CubeSat under high and low temperature conditions;

[0042] A thermal balance test system was set up, and the thermally modified prototype CubeSat was placed inside the thermal balance test system;

[0043] The thermal equilibrium test boundary temperatures under high and low temperature conditions were obtained based on the characteristic point temperatures of the thermally modified prototype CubeSat under high and low temperature conditions, respectively.

[0044] The thermal balance test is conducted based on the thermal balance test boundary temperatures under high and low temperature conditions.

[0045] Specifically, S1, thermal modification of the prototype CubeSat. Depending on the specific thermal design method, optional modifications to the CubeSat may include, but are not limited to, painting of the structural plate frame, pasting and wiring of electric heaters, pasting and wiring of thermistors / thermocouples, multi-layer wrapping of internal equipment, and multi-layer wrapping of external cabin plates, so that the experimental satellite is consistent with the final launch state.

[0046] S2. Based on finite element thermal simulation analysis, calculate the temperature of characteristic points of the CubeSat under high and low temperature conditions;

[0047] Preferably, the finite element simulation analysis can use the thermal simulation model used in the thermal design;

[0048] Preferably, the satellite's maximum external heat flux + maximum operating mode is the high-temperature condition, and the satellite's minimum external heat flux + minimum operating mode is the low-temperature condition.

[0049] Preferably, the characteristic point temperature is taken as the temperature at the satellite structural plate or on a non-heat source device.

[0050] S3. Construct a thermal balance test system;

[0051] like Figure 2 As shown, the thermal balance test system includes a vacuum simulation chamber, a simulation module, an experimental satellite, and a payload platform; among which,

[0052] The thermally modified prototype CubeSat is positioned on the upper part of the carrying platform; the simulation cabin is connected to the carrying platform, and the thermally modified prototype CubeSat is located inside the simulation cabin; the simulation cabin is located inside the vacuum simulation chamber.

[0053] The thermal balance test system also includes a measurement system.

[0054] The simulation chamber is a sealed enclosure welded from aluminum plates. A temperature-controlled heater is attached to the outer surface to control the temperature of the simulation chamber, and a coating is sprayed on the inner surface to ensure that the infrared emissivity is not less than 0.85.

[0055] Optionally, the outer surface of the simulation chamber is covered with multiple layers to control the heat dissipation of the simulation chamber to the vacuum simulation chamber;

[0056] The experimental satellite was a hot-modified experimental satellite;

[0057] The cargo platform is equipped with an electric heater, which can be synchronously heated and temperature-controlled with the simulation cabin.

[0058] The measurement system is connected to the test satellite via a cable used for on-board signal transmission and for the transmission of signals from the thermal test electric heater and thermistors / thermocouples.

[0059] S4. Calculate the boundary temperature of the thermal balance test under low temperature and high temperature conditions based on the characteristic point temperature described in S2.

[0060] Taking a certain type of satellite as an example, the thermal test boundary temperature is calculated by analyzing the boundary temperature.

[0061] According to the law of conservation of energy

[0062] Q1 + Q2 = Q3

[0063] Q1 is the heat flux absorbed by the satellite in space (solar radiation, solar radiation reflected by the Earth, infrared radiation from the Earth, and cold space background radiation), Q2 is the total heat consumption of the satellite equipment, and Q3 is the heat flux radiated by the satellite into space.

[0064] The thermal analysis boundary temperature of the satellite under high and low temperature conditions was obtained using thermal analysis software, which was T. L and T H (Using subscripts L and H to represent low-temperature and high-temperature operating conditions respectively, the same applies below), the corresponding operating condition satellite is heat consumption Q. 2L and Q 2H Considering the thermal experiment, the heat balance equation is:

[0065] Low temperature operating conditions:

[0066] Q 2L =Q 3L -Q 1L

[0067]

[0068] Among them, Q 2L Q represents the total heat dissipation of satellite equipment under low-temperature operating conditions. 3L Q represents the heat flow radiated from a satellite into space under cryogenic conditions. 1L To allow the satellite to absorb space heat flux under cryogenic conditions, ε0 is the system emissivity, and A S Let T be the surface area of ​​the satellite, σ be the Boltzmann constant, and T be the surface area of ​​the satellite. C,L T represents the boundary temperature for the thermal balance test of the satellite under low-temperature operating conditions. S,L This refers to the boundary temperature of the thermal balance test in the simulation chamber under low-temperature conditions.

[0069] The boundary temperature of the thermal equilibrium test under high-temperature conditions is obtained by the following formula:

[0070] Q2H =Q 3H -Q 1H ;

[0071]

[0072] Among them, Q 2H Q represents the total heat dissipation of satellite equipment under high-temperature operating conditions. 3H Q represents the heat flow radiated from a satellite into space under high-temperature operating conditions. 1H To allow the satellite to absorb space heat flux under high-temperature operating conditions, ε0 is the system emissivity, and A S Let T be the surface area of ​​the satellite, σ be the Boltzmann constant, and T be the surface area of ​​the satellite. C,H T represents the boundary temperature for the thermal balance test of the satellite under high-temperature conditions. S,H This refers to the boundary temperature of the thermal balance test in the simulation chamber under high-temperature conditions.

[0073] ε0 represents the system emissivity, which is calculated using the radiation formula for a closed system:

[0074]

[0075] In the formula A S A represents the surface area of ​​a satellite. C The area represents the simulated cabin surface area, and σ represents the Boltzmann constant, which is taken as 5.67 × 10⁻⁶. -8 ;

[0076] The surface area A of a certain type of CubeSat S =0.08m 2 Simulated cabin surface area A C =1.12m 2 Surface average emissivity ε S =0.74, simulated cabin interior surface emissivity ε C =0.9, substituting into equation (2), we find the system emissivity ε0 = 0.735.

[0077] Using measuring points on the surface of the structural plate as reference temperature points, the temperature T under low-temperature conditions was calculated using thermal analysis software. L =1.2℃, Q 2L =4W; T under high temperature conditions H =22℃, Q 2H =6.2W, substituting into equation (1), the thermal test boundary temperatures under low temperature and high temperature conditions are respectively

[0078] T C,L = -14.6℃, T C,H =1.9℃

[0079] S5. Set the thermal balance test boundary temperature as described in S4 and conduct the thermal balance test.

[0080] By adjusting the power of the electric heaters in the simulation chamber and the cargo platform, the temperatures of the simulation chamber and the cargo platform are controlled to the thermal test boundary temperatures, and a thermal balance test is conducted. It should be noted that the process of conducting a thermal balance test based on the thermal balance test boundary temperatures under high and low temperature conditions is well known and will not be described in detail here.

[0081] This invention reduces the production of infrared cages and process boards, thus lowering the cost of CubeSat manufacturing and development. In addition to the multi-layered outer layer, the experimental satellite maintains the same state as the launched satellite, shortening the satellite assembly time during development. This invention also reduces the difficulty of model correction due to inaccurate external heat flow simulation during thermal balance testing. This invention has the advantages of accurate testing, low cost, and high reliability, making it suitable for CubeSat applications requiring low cost, short cycle time, and large-scale production.

[0082] Although the present invention has been disclosed above with reference to preferred embodiments, it is not intended to limit the present invention. Any person skilled in the art can make possible changes and modifications to the technical solutions of the present invention by utilizing the methods and techniques disclosed above without departing from the spirit and scope of the present invention. Therefore, any simple modifications, equivalent changes and alterations made to the above embodiments based on the technical essence of the present invention without departing from the content of the technical solutions of the present invention shall fall within the protection scope of the technical solutions of the present invention.

Claims

1. A method of thermal balance testing of a CubeSat, characterized by The method comprises the following steps: thermally modifying a positive sample cube star; calculating characteristic point temperatures of the thermally modified positive sample cube star under high-temperature and low-temperature working conditions; building a heat balance test system and placing the thermally modified positive sample cube star in the heat balance test system; obtaining heat balance test boundary temperatures under the high-temperature and low-temperature working conditions according to the characteristic point temperatures of the thermally modified positive sample cube star under the high-temperature and low-temperature working conditions respectively; performing heat balance tests according to the heat balance test boundary temperatures under the high-temperature and low-temperature working conditions.

2. The method of claim 1, wherein: The step of thermally modifying the positive sample cube star comprises the following steps: spraying paint on a structural plate frame of the positive sample cube star; coating a plurality of thermal insulation layers on an outer surface of in-satellite equipment of the positive sample cube star; coating a plurality of thermal insulation layers on an outer surface of an out-of-satellite cabin plate of the positive sample cube star; and sticking an electric heater on an outer surface of the positive sample cube star.

3. The method of claim 1, wherein: The characteristic point temperatures of the thermally modified positive sample cube star under the high-temperature and low-temperature working conditions are calculated based on finite element thermal simulation analysis.

4. The method of claim 3, wherein: The maximum external heat flow + maximum working mode of the positive sample cube star is the high-temperature working condition. The minimum external heat flow + minimum working mode of the positive sample cube star is the low-temperature working condition.

5. The method of claim 3, wherein: The characteristic point temperatures are temperatures at satellite structural plates or on non-thermal source equipment.

6. The method of claim 1, wherein: The heat balance test system comprises a vacuum simulation chamber, a simulation cabin and a load platform. The thermally modified positive sample cube star is arranged on an upper portion of the load platform. The simulation cabin is connected to the load platform, and the thermally modified positive sample cube star is arranged in the simulation cabin. The simulation cabin is arranged in the vacuum simulation chamber.

7. The method of claim 6, wherein: The simulation cabin is coated with a plurality of thermal insulation layers on an outer surface thereof, and is used for controlling heat dissipation of the simulation cabin to the vacuum simulation chamber. The load platform is stuck with an electric heater, and can be synchronously heated and temperature-controlled with the simulation cabin.

8. The method of claim 1, wherein: The heat balance test boundary temperature under the low-temperature working condition is obtained by the following formula: Q 2L = Q 3L - Q 1L ; Wherein, Q 2L is the total heat consumption of the satellite equipment under low temperature condition, Q 3L is the space radiation heat flow of the satellite under low temperature condition, Q 1L is the space heat flow absorbed by the satellite under low temperature condition, ε0 is the system emissivity, A S is the surface area of the satellite, σ is the Boltzmann constant, T C,L is the thermal balance test boundary temperature of the satellite under low temperature condition, T S,L is the thermal balance test boundary temperature of the simulation cabin under low temperature condition.

9. The method of claim 1, wherein: The heat balance test boundary temperature under the high-temperature working condition is obtained by the following formula: Q 2H = Q 3H - Q 1H ; Wherein, Q 2H is the total heat consumption of the satellite equipment under high temperature working condition, Q 3H is the space radiation heat flow of the satellite under high temperature working condition, Q 1H is the space heat flow absorbed by the satellite under high temperature working condition, ε0 is the system emissivity, A S is the surface area of the satellite, σ is the Boltzmann constant, T C,H is the thermal balance test boundary temperature of the satellite under high temperature working condition, T S,H is the thermal balance test boundary temperature of the simulation cabin under high temperature working condition.

10. The method of claim 8 or 9, wherein: The system emissivity ε0 is: where A is the surface area of the satellite, σ is the Boltzmann constant, A S is the surface area of the satellite, σ is the Boltzmann constant, A C is the surface area of the satellite, σ is the Boltzmann constant, A S is the surface area of the satellite, σ is the Boltzmann constant, A C is the surface area of the satellite, σ is the Boltzmann constant, A

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