Spacecraft high-efficiency thermal test method

Abstract

The application provides a spacecraft high-efficiency thermal test method, comprising the following steps: S1: setting a transfer orbit balance working condition, verifying a substitute heater design, carrying out a test according to a device state of the spacecraft in the transfer orbit, and obtaining a transfer orbit balance working condition temperature; S2: low-temperature working condition 1 setting: not simulating an on-orbit low-temperature working condition of the spacecraft, and setting a balance working condition meeting an overall satellite thermal vacuum cycle low-temperature test condition; S3: high-temperature working condition 1 setting; not simulating an on-orbit high-temperature working condition of the spacecraft, and setting a balance working condition meeting an overall satellite thermal vacuum cycle high-temperature test condition; S4: entering high- and low-temperature working conditions 2-4: the sequence is: low-temperature working condition 2→ high-temperature working condition 2→ low-temperature working condition 3→ high-temperature working condition 3→ low-temperature working condition 4→ high-temperature working condition 4. The method provided by the application can reduce waiting temperature balance stable time while ensuring thermal balance and the effectiveness of thermal vacuum test data acquisition, and achieves the purpose of optimizing an overall satellite thermal test process.

Description

Technical Field

[0001] This invention relates to the field of aerospace testing and verification, and in particular to a high-efficiency thermal testing method for spacecraft. Background Technology

[0002] Spacecraft thermal testing includes thermal balance testing and thermal vacuum testing, typically conducted first. The purpose of thermal balance testing is to verify the correctness and effectiveness of the spacecraft's thermal control design and implementation; it also verifies the correctness of the thermal control calculations and analysis, and modifies the thermal mathematical model based on the test results, enabling the prediction of satellite flight temperatures using the revised model. The purpose of thermal vacuum testing is to apply more severe temperature stress to the spacecraft under vacuum conditions than its normal operating environment, exposing design and quality defects and verifying the spacecraft's ability to operate normally under the applied environmental conditions. Both thermal balance and thermal vacuum tests are important spacecraft verification tests.

[0003] Both thermal equilibrium tests and thermal vacuum tests require data acquisition or performance testing after the temperature has reached equilibrium and stabilized. The temperature equilibrium stability criterion for a thermal equilibrium test is that the monotonic temperature change is less than 1°C over 5 consecutive hours. The temperature equilibrium stability criterion for a thermal vacuum test is that the monotonic temperature change or fluctuation is less than 1°C / hour.

[0004] For geostationary orbit satellites, the original whole-satellite thermal balance test typically includes four typical test conditions: Thermal balance condition 1 - transfer orbit, used to verify the satellite's alternative heater design; Thermal balance condition 2 - cryogenic condition for the whole satellite; Thermal balance condition 3 - high temperature condition for the satellite's north plate; and Thermal balance condition 4 - high temperature condition for the satellite's south plate. The whole-satellite thermal vacuum test is conducted after the thermal balance test. The original thermal vacuum test consists of four temperature cycles, corresponding to thermal vacuum conditions 1 through 4. Each cycle involves stopping at either the high or low temperature end to test satellite performance. See the cross-sectional diagrams of the thermal balance test and thermal vacuum test. Figure 1 .

[0005] With the rapid development of the aerospace industry, the types and numbers of spacecraft have increased significantly. Ensuring the timely completion of spacecraft development schedules places higher demands on the efficiency of thermal testing technology. Existing test design methods are gradually becoming inadequate to meet mission requirements, mainly in the following aspects:

[0006] (1) The original spacecraft thermal test design strictly followed the order of conducting thermal balance test first and then thermal vacuum test. The two tests were decoupled from each other, and the acquisition and analysis of test data such as temperature acquisition were independent, without being merged and optimized.

[0007] (2) The original spacecraft thermal testing work was carried out by thermal balance test personnel and thermal vacuum test personnel respectively. The test plan, test execution and test summary were all carried out separately and independently, resulting in duplication of work and waste of manpower, which meant that the existing human resources could not support the development of a larger number of spacecraft.

[0008] (3) Spacecraft thermal tests need to be conducted in a vacuum chamber. During the test, in order to ensure that the set temperature can be reached, liquid nitrogen needs to be continuously circulated in the chamber and the temperature-controlled lamp array needs to be turned on. Liquid nitrogen and lamp array are both high-consumption materials / energy sources. Conducting thermal balance and thermal vacuum tests independently will inevitably lead to cost waste. Summary of the Invention

[0009] To address the problems existing in the prior art, this invention provides a high-efficiency thermal testing method for spacecraft, which, while ensuring the effectiveness of thermal balance and thermal vacuum test data acquisition, reduces the waiting time for temperature equilibrium stabilization, thereby optimizing the overall satellite thermal testing process.

[0010] To achieve the above objectives, the present invention adopts the following solution:

[0011] This invention provides a high-efficiency thermal testing method for spacecraft, comprising the following steps:

[0012] S1: Set the transfer orbit equilibrium conditions, verify the alternative heater design, conduct experiments based on the spacecraft's equipment status in transfer orbit, and obtain the transfer orbit equilibrium temperature T. t-trans ;

[0013] S2: Cryogenic Condition 1 Setting: Does not simulate the cryogenic conditions of the spacecraft in orbit, but sets it to a balanced condition that meets the cryogenic test conditions of the whole satellite thermal vacuum cycle;

[0014] S3: High-temperature condition 1 setting; does not simulate the high-temperature conditions of spacecraft in orbit, but is set to a balanced condition that meets the high-temperature test conditions of the whole satellite thermal vacuum cycle;

[0015] S4: Entering high and low temperature operating conditions 2-4:

[0016] The sequence is: Low temperature condition 2 → High temperature condition 2 → Low temperature condition 3 → High temperature condition 3 → Low temperature condition 4 → High temperature condition 4.

[0017] In some embodiments, the present invention further includes the following technical features:

[0018] In step S1, the transfer orbit equilibrium operating temperature T t-trans The determination of temperature balance requires achieving the required stability, i.e., a change of less than 1°C over 5 consecutive hours. The criterion for temperature balance stability is that the monotonic change in temperature is less than 1°C over 5 consecutive hours.

[0019] Step S2 specifically involves, based on the transfer orbit equilibrium condition, shutting off some loads to maintain consistency with the thermal control analysis state under low-temperature conditions. The temperature equilibrium stability criterion is that the monotonic temperature change is less than 1°C over 5 consecutive hours, thus obtaining the low-temperature equilibrium temperature T. t-low After the temperature balance and stability criteria are met, low-temperature operating condition tests are carried out.

[0020] Obtain the low-temperature equilibrium operating temperature T t-low The specific method is as follows:

[0021] The temperature T at the temperature measurement point under the cryogenic conditions in orbit of the spacecraft was obtained through thermal control calculations and analysis. a-low In T a-low Based on the test temperature stress T add-low The temperature T under low-temperature equilibrium condition is obtained. t-low =T a-low +T add-low Among them, the test temperature stress T add-low The equipment testing requirements were obtained by consulting the documentation.

[0022] Obtain the T a-low The specific method is as follows:

[0023] Based on the state settings of the transfer orbit thermal control analysis, some payloads were shut down, ensuring that at least 80% of the spacecraft's equipment temperature was lower than the equipment temperature T calculated from the transfer orbit thermal control analysis. a-trans Reduce the temperature by 5–10°C, and ensure that all equipment does not exceed the lower limit of the acceptance temperature T required by the equipment. acpt-low The thermal analysis temperature T at the temperature measurement point under the cryogenic conditions in orbit of the spacecraft was obtained. a-low T acpt-low ≤T a-low <T a-trans .

[0024] Step S3 specifically involves setting up a high-temperature equilibrium condition, energizing the north and south light arrays, adjusting the heat flux of the light arrays to the high-temperature condition thermal control analysis calculation value, and determining the temperature equilibrium stability criterion as a monotonic temperature change of less than 1°C over 5 consecutive hours, thus obtaining the high-temperature condition temperature T. t-high After the temperature balance and stability criteria are met, high-temperature operating condition tests are carried out.

[0025] Obtain the high-temperature operating temperature T t-high The specific method is as follows:

[0026] The temperature T measured under high-temperature conditions in orbit was obtained through thermal control calculations and analysis. a-high In T a-high Based on the test temperature stress T add-high The temperature T under low-temperature equilibrium condition is obtained. t-high =T a-high +Tadd-high Among them, the test temperature stress T add -hi g h. Refer to the equipment testing requirements to obtain them.

[0027] Obtain the T a-high The specific method is as follows:

[0028] Thermal control analysis was conducted under conditions where both the south and north plates of the spacecraft experienced high temperatures simultaneously, ensuring that at least 80% of the equipment temperature across the entire satellite was lower than the baseline data T calculated in the previous step. b Increase the temperature by another 5-10°C, and ensure that all equipment does not exceed the upper limit of the acceptance temperature T. acpt-high The thermal analysis temperature T at the temperature measurement point under the high-temperature operating conditions in orbit of the spacecraft was obtained. a-high T b <T a-high ≤T acpt-high .

[0029] Obtain the T b The specific method is as follows:

[0030] Thermal control analysis was conducted on the north plate of the spacecraft under high-temperature conditions to obtain the thermal analysis temperature T. summ Thermal control analysis was conducted on the south plate of the spacecraft under high-temperature conditions, and the thermal analysis temperature T was obtained. wint , get T b =max{T summ ,T wint}

[0031] In step S4,

[0032] Low temperature conditions 2-4: The temperature setting is consistent with that of low temperature condition 1; the equilibrium stability criterion is the monotonic change value of temperature or the fluctuation is less than 1℃ / hour. After the temperature equilibrium stability criterion is reached, the low temperature condition test is carried out.

[0033] High-temperature conditions 2-4: The temperature setting is consistent with that of high-temperature condition 1; the equilibrium stability criterion is a monotonic temperature change or fluctuation of less than 1℃ / hour. After the temperature equilibrium stability criterion is reached, high-temperature condition tests are conducted.

[0034] The beneficial effects of this invention are:

[0035] Shorten the proportion of the waiting time for temperature equilibrium stabilization in spacecraft thermal vacuum tests relative to the total test duration, optimize data acquisition and analysis during thermal testing, and reduce labor and material costs. Attached Figure Description

[0036] Figure 1 This is a cross-sectional diagram of the thermal test process;

[0037] Figure 2 This is a cross-sectional view of the optimized thermal testing process in an embodiment of the present invention. Detailed Implementation

[0038] To make the technical solutions and advantages of the present invention clearer, the technical solutions of the embodiments of the present invention will be fully described below with reference to the accompanying drawings. Obviously, the described embodiments are only some, not all, of the embodiments of the present invention. All other embodiments obtained by those skilled in the art based on the embodiments of the present invention without creative effort are within the scope of protection of the present invention.

[0039] Embodiments of the present invention are described in detail below. Examples of these embodiments are shown in the accompanying drawings, wherein the same or similar reference numerals denote the same or similar elements or elements having the same or similar functions throughout. The embodiments described below with reference to the accompanying drawings are exemplary and are only used to explain the present invention, and should not be construed as limiting the present invention.

[0040] By setting new thermal equilibrium and thermal vacuum test conditions, some conditions from the original thermal equilibrium and thermal vacuum tests are integrated. Temperature equilibrium stabilization times are set for low-temperature condition 1 and high-temperature condition 1 to meet the criterion for temperature equilibrium stability in thermal equilibrium tests (temperature monotonic change less than 1°C over 5 consecutive hours). This achieves the purpose of verifying the thermal analysis model and providing temperature data for the correction of the thermal mathematical model, while simultaneously meeting the equipment temperature testing conditions for thermal vacuum tests. This approach not only verifies the thermal analysis model and provides temperature data for its correction, but also reduces the waiting time for temperature equilibrium in spacecraft thermal vacuum tests, thus improving test efficiency.

[0041] This invention proposes a thermal test temperature control scheme with five different thermal test conditions:

[0042] (1) Transfer track balance condition: Verify the alternative heater design.

[0043] (2) Cryogenic Condition 1: This condition does not simulate the cryogenic conditions of a spacecraft in orbit; instead, it is set as a balanced condition that meets the cryogenic testing conditions of the entire satellite's thermal vacuum cycle. The specific method is as follows:

[0044] The temperature T measured under the cryogenic conditions of the spacecraft in orbit was obtained through thermal control calculations and analysis. a-low In T a-low Based on the test temperature stress T add-low The temperature T under low-temperature equilibrium condition is obtained. test-low =T a-low +T add-low Among them, the test temperature stress T add-low The equipment testing requirements were obtained by consulting the documentation.

[0045] (3) High-temperature condition 1: Instead of simulating the high-temperature conditions of a spacecraft in orbit, this condition is set to a balanced condition that meets the requirements of a full-satellite thermal vacuum cycle high-temperature test. The specific method is as follows:

[0046] The temperature T measured under high-temperature conditions in orbit was obtained through thermal control calculations and analysis. a-high In T a-high Based on the test temperature stress T add-high The temperature T under low-temperature equilibrium condition is obtained. test-high =T a-high +T add-high Among them, the test temperature stress T add-low The equipment testing requirements were obtained by consulting the documentation.

[0047] Based on the above method, the setting conditions for low-temperature equilibrium conditions and high-temperature equilibrium conditions are obtained from the analysis of the pre-test thermal mathematical model to ensure that they meet the temperature conditions for thermal vacuum testing. After the conditions reach the temperature equilibrium requirements, the spacecraft thermal vacuum test is carried out.

[0048] (4) Low temperature conditions 2-4: The equilibrium and stable temperature is set to T test-low .

[0049] (5) High-temperature operating conditions 2-4: The equilibrium and stable temperature is set to T test-high .

[0050] The optimized thermal test profile is attached. Figure 2 .

[0051] The technical solution of the present invention will be described below with reference to a specific embodiment.

[0052] 1. Thermal control analysis implementation

[0053] (1) Transfer orbit thermal control analysis: Based on the equipment status of the spacecraft in the transfer orbit, thermal control analysis of the spacecraft is carried out to obtain the transfer orbit thermal analysis temperature T. a-trans .

[0054] (2) Low-temperature thermal control analysis: Based on the state settings of the transfer orbit thermal control analysis, some payloads were shut down, so that no less than 80% of the spacecraft's equipment temperature was lower than the equipment temperature T calculated by the transfer orbit thermal control analysis. a-trans Reduce the temperature by 5–10°C, and ensure that all equipment does not exceed the lower limit of the acceptance temperature T required by the equipment. acpt-low The thermal analysis temperature T at the temperature measurement point under the cryogenic conditions in orbit of the spacecraft was obtained. a-low T acpt-low ≤T a-low <T a-trans .

[0055] (3) Thermal control analysis at the summer solstice and winter solstice in orbit: Thermal control analysis was carried out under the high temperature (summer solstice) conditions on the north plate of the spacecraft, and the thermal analysis temperature T was obtained. summ Thermal control analysis was conducted on the south plate of the spacecraft under high-temperature (winter solstice) conditions, and the thermal analysis temperature T was obtained. wint , get T b =max{T summ ,T wint}

[0056] (4) Thermal control analysis under high-temperature conditions: Thermal control analysis was conducted under the condition that the south and north plates of the spacecraft were simultaneously at high temperatures, ensuring that the temperature of no less than 80% of the equipment on the entire satellite was lower than the baseline data T calculated in the previous step. b Increase the temperature by another 5-10°C, and ensure that all equipment does not exceed the upper limit of the acceptance temperature T. acpt-high The thermal analysis temperature T at the temperature measurement point under the high-temperature operating conditions in orbit of the spacecraft was obtained. a-high T b <T a-high ≤T acpt-high .

[0057] 2. Thermal test implementation

[0058] (1) Transfer track balancing condition:

[0059] Experiments were conducted based on the equipment status of the spacecraft in transfer orbit to obtain the equilibrium operating temperature T in transfer orbit. t-trans The requirement is to achieve the stability required for thermal equilibrium, i.e., a change of less than 1°C over 5 consecutive hours. The criterion for temperature equilibrium stability is that the monotonic temperature change is less than 1°C over 5 consecutive hours.

[0060] (2) Low Temperature Condition 1: Based on the transfer track equilibrium condition, some loads are shut off, maintaining consistency with the low temperature condition thermal control analysis state. The temperature equilibrium stability criterion is that the monotonic temperature change is less than 1℃ over 5 consecutive hours, thus obtaining the low temperature condition temperature T. t-low After the temperature equilibrium stability criterion is met, low-temperature operating condition tests are carried out.

[0061] (3) High-Temperature Condition 1: Set up a high-temperature equilibrium condition, power on the north and south light arrays, and adjust the heat flow of the light arrays to the high-temperature condition thermal control analysis calculation value. The temperature equilibrium stability criterion is that the monotonic temperature change is less than 1℃ within 5 consecutive hours, and the high-temperature condition temperature T is obtained. t-high After the temperature equilibrium stability criterion is reached, high-temperature operating condition tests are carried out.

[0062] (4) Entering high and low temperature operating conditions 2-4:

[0063] The sequence is: Low temperature condition 2 → High temperature condition 2 → Low temperature condition 3 → High temperature condition 3 → Low temperature condition 4 → High temperature condition 4.

[0064] Low-temperature operating conditions 2-4: The temperature setting is consistent with that of low-temperature operating condition 1. The equilibrium stability criterion is a monotonic temperature change or fluctuation of less than 1℃ / hour. After the temperature equilibrium stability criterion is reached, low-temperature operating condition tests are carried out.

[0065] High-temperature conditions 2-4: The temperature settings are consistent with those of high-temperature condition 1. The equilibrium stability criterion is a monotonic temperature change or fluctuation of less than 1℃ / hour. After the temperature equilibrium stability criterion is met, high-temperature condition testing is conducted.

[0066] By comparing the time costs before and after optimization, the optimization can shorten the test cycle by approximately 3 days, as shown in Table 1. Simultaneously, the test data analysis process has been integrated and optimized, meeting the task requirements.

[0067] Table 1. Estimated thermal test time before and after optimization.

[0068]

[0069] In the description of this specification, references to terms such as "an embodiment" and "example" refer to specific features, structures, or characteristics described in connection with that embodiment or example, which are included in at least one embodiment or example of the invention. In this specification, the illustrative expressions of the above terms are not necessarily intended to refer to corresponding embodiments or examples in a suitable manner.

[0070] It must be pointed out that the above description of the embodiments is not intended to limit the invention but only to help understand the core idea of ​​the invention. For those skilled in the art, any improvements to the invention and equivalent alternatives made to the invention without departing from the principle of the invention are also within the scope of protection of the claims of the invention.

Claims

1. A high-efficiency thermal testing method for spacecraft, characterized in that, Includes the following steps: S1: Set the transfer orbit equilibrium conditions, verify the alternative heater design, conduct experiments based on the spacecraft's equipment status in transfer orbit, and obtain the transfer orbit equilibrium temperature. ; S2: Cryogenic Condition 1 Setting: Does not simulate the cryogenic conditions of the spacecraft in orbit, but sets it to a balanced condition that meets the cryogenic test conditions of the whole satellite thermal vacuum cycle; S3: High-temperature condition 1 setting; does not simulate the high-temperature conditions of spacecraft in orbit, but is set to a balanced condition that meets the high-temperature test conditions of the whole satellite thermal vacuum cycle; S4: Entering high and low temperature operating conditions 2~4: The sequence is: Low temperature condition 2 → High temperature condition 2 → Low temperature condition 3 → High temperature condition 3 → Low temperature condition 4 → High temperature condition 4; In step S4, Low temperature conditions 2~4: The temperature setting is consistent with that of low temperature condition 1; the equilibrium stability criterion is the monotonic change value of temperature or the fluctuation is less than 1℃ / hour. After the temperature equilibrium stability criterion is reached, the low temperature condition test is carried out. High-temperature conditions 2-4: The temperature setting is consistent with that of high-temperature condition 1; the equilibrium stability criterion is the monotonic change value of temperature or the fluctuation is less than 1℃ / hour. After the temperature equilibrium stability criterion is reached, the high-temperature condition test is carried out.

2. The method according to claim 1, characterized in that, In step S1, the temperature of the transfer track balancing condition. The determination of temperature balance requires achieving the required stability, i.e., a change of less than 1°C over 5 consecutive hours. The criterion for temperature balance stability is that the monotonic change in temperature is less than 1°C over 5 consecutive hours.

3. The method according to claim 1, characterized in that, Step S2 specifically involves, based on the transfer orbit equilibrium condition, shutting off some loads to maintain consistency with the thermal control analysis state under low-temperature conditions. The temperature equilibrium stability criterion is that the monotonic temperature change is less than 1°C over 5 consecutive hours, thus obtaining the low-temperature equilibrium temperature. After the temperature balance and stability criteria are met, low-temperature operating condition tests are carried out.

4. The method according to claim 3, characterized in that, Obtain the temperature of the low-temperature equilibrium condition. The specific method is as follows: The temperature at the temperature measurement point under the cryogenic conditions of the spacecraft in orbit was obtained through thermal control calculations and analysis. ,exist Based on the test temperature stress The temperature under low-temperature equilibrium condition is obtained. Among them, the test temperature stress The equipment testing requirements were obtained by consulting the documentation.

5. The method according to claim 4, characterized in that, Obtain the The specific method is as follows: Based on the state settings of the transfer orbit thermal control analysis, some payloads were shut down, ensuring that at least 80% of the spacecraft's equipment temperatures were lower than those calculated from the transfer orbit thermal control analysis. Reduce the temperature by 5-10°C, and ensure that all equipment does not exceed the lower limit of the acceptance temperature required by the equipment. The thermal analysis temperature of the temperature measurement point under the cryogenic conditions in orbit of the spacecraft was obtained. , .

6. The method according to claim 1, characterized in that, Step S3 specifically involves setting up a high-temperature equilibrium condition, powering on the north and south light arrays, adjusting the heat flux of the light arrays to the high-temperature condition thermal control analysis calculation value, and determining the temperature equilibrium stability criterion as a monotonic temperature change of less than 1°C over 5 consecutive hours to obtain the high-temperature condition temperature. After the temperature balance and stability criteria are met, high-temperature operating condition tests are carried out.

7. The method according to claim 6, characterized in that, Obtain the temperature under the high-temperature operating condition. The specific method is as follows: The temperature of the temperature measurement point under high-temperature conditions in orbit was obtained through thermal control calculations and analysis. ,exist Based on the test temperature stress The temperature under high-temperature equilibrium condition is obtained. Among them, the test temperature stress The equipment testing requirements were obtained by consulting the documentation.

8. The method according to claim 7, characterized in that, Thermal control analysis was conducted on the north plate of the spacecraft under high-temperature conditions to obtain the thermal analysis temperature. Thermal control analysis was conducted on the south plate of the spacecraft under high-temperature conditions to obtain the thermal analysis temperature. ,get .

9. The method according to claim 8, characterized in that, Obtain the The specific method is as follows: Thermal control analysis was conducted under conditions where both the south and north plates of the spacecraft experienced high temperatures simultaneously, ensuring that at least 80% of the equipment temperatures across the entire satellite were lower than the baseline data calculated in the previous step. Increase the temperature by another 5-10°C, ensuring that all equipment does not exceed the upper limit of the acceptance temperature. The thermal analysis temperature of the measurement point under the high-temperature operating conditions of the spacecraft in orbit was obtained. , .

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