Design and Verification of Thermal Control System of Communication Satellite

Abstract

The multiple working modes, complex working conditions, frequent changes in external heat flux, and high power consumption of communication satellites all pose great difficulties to their thermal design. This paper mainly describes the design of a thermal control system for high-power communication satellites. Firstly, new efficient heat transfer technologies and thermal control materials for spacecraft are introduced. Secondly, the structure and internal heat source of the satellite are introduced. Thirdly, the external heat fluxes are analyzed, and the position of the heat dissipation surface and extreme conditions are confirmed. Then, a thermal control system is designed around the difficulties of thermal control. With heat pipes, the temperature uniformity of +Y deck, −Y deck, and +Z deck increased by 8 °C, 9.9 °C, and 34.2 °C, respectively. Furthermore, the maximum temperature of the power controller, secondary power supply, bidirectional frequency converter, and solid discharge decreased by 32.5 °C, 22.0 °C, 14.0 °C, and 164 °C, respectively. Finally, a thermal balance test is performed. The test results show that the temperatures of the solid-state power amplifier, on-board computer, power controller, secondary power supply, and bidirectional frequency converter meet the requirements of the thermal control indices. In addition, the temperature of thermal-sensitive components such as batteries and the storage tank also meets the requirements. The thermal design scheme is reasonable and feasible, and the thermal balance test verifies the correctness of the thermal design.

1. Introduction

As an important safeguard service subsystem of spacecraft, the mission of spacecraft thermal control subsystems is to provide a good working temperature environment to maintain normal operation of the spacecraft in orbit during its life span, and the performance and reliability of the thermal control system directly determines the success or failure of the satellite’s development and working life. With the rapid development of China’s space field, especially, in recent years, the rapid development of micro-nano-satellites, the spacecraft structure layout is more and more complex and the functions and number of instruments and equipment continue to increase. At the same time, with the development of micro-electro-mechanical system technology, the focus on the integration of instruments and equipment is increasingly high; spacecraft are moving towards the direction of the rapid development of micro-satellites, which leads to the development of low-cost thermal control technology with high precision, high heat flow, etc., and the development of thermal control technology puts forward a higher and higher level of technical development. The development of thermal control technology has thus put forward more and more requirements.

In order to meet people’s demand for communication services, especially the growing demand for ubiquitous connections, large-scale low Earth orbit (LEO) satellite constellation technology [1,2] has emerged. With the continuous maturity of small satellite manufacturing and satellite launch technology [3,4], the deployment of small satellites is accelerating globally. In the severe international situation of fierce competition in LEO internet constellations, the construction of LEO satellite internet communication engineering in China has also been put on the agenda, and the implementation of LEO satellite internet constellations is being vigorously promoted. By reducing latency, increasing bandwidth, and expanding service range, LEO satellites can achieve faster and more reliable internet access [5,6].

The multiple working modes, complex working conditions, frequent changes in external heat flux, and high power consumption of communication satellites all pose great difficulties to their thermal design. As an important subsystem of satellites, the thermal control system is designed to ensure that the temperature of each part of the satellite is within its specified temperature range [7,8]. Due to the harsh space thermal environment where spacecraft work (including the 4 K cold black vacuum environment, solar radiation, Earth infrared radiation, and Earth albedo) [9,10], its in-orbit temperature is prone to significant fluctuations, which will cause a deformation of the satellite’s structure [11]. Therefore, in order to reduce the impact of the space thermal environment and ensure the stability of satellite systems, strict thermal control design must be carried out [12,13].

In recent years, new high-efficiency heat transfer technologies of spacecraft mainly include ultra-high thermal conductive materials [14], mini-heat pipes, axial-grooved aluminum ammonia heat pipes, two-phase pumped-loop and cryogenic-loop heat pipes [15,16,17], a spray cooling system [18], and thermal control technologies based on MEMS [19]. All of these are the key means for spacecraft thermal control, including high-heat-flux heat dissipation, high-power heat transfer, separated heat source thermal dissipation, and cryogenic heat transfer.

The thermal control material is an important medium to realize the thermal control function of spacecraft, and is also the basis for the development of spacecraft thermal control technology. The wide application of multilayer insulation materials, thermal conductive materials, thermal coatings, and interface materials in the spacecraft thermal control system complete the thermal control of spacecraft with the thermophysical properties of the materials [20,21]. According to an analysis of international spacecraft development trends, the future demand for thermal control materials is mainly reflected in three aspects [22]: (1) Efficient insulation materials, high thermal conductive materials, high-emissivity thermal control coatings, and other thermal control materials in low-temperature environment of 20~100 K. (2) Thermal control materials such as insulation materials and high-temperature-resistant coating materials in a high-temperature environment of 750~2000 K. (3) The development demand for high-thermal-conductive materials, high-thermal-conductive interface materials, intelligent thermal control coating materials, and thermal storage materials is proposed in response to the common requirements of thermal conductive enhancement and efficient thermal dissipation for spacecraft with increasing integration.

The research object of this paper is a test satellite, and its thermal design difficulties are mainly reflected in the following aspects: (1) The β-angle alternates between −90° and +90°, causing the satellite’s surfaces to take turns being exposed to sunlight and experiencing severe fluctuations in external heat fluxes. When the β-angle is greater than 58°, the track is in the full illumination range. Within one year, there can be 2–3 consecutive light periods on the track, with a maximum illumination time of about 49 days. (2) The power consumption of the solid-state power amplifier is 448 W, and its structure is relatively complex, with certain space constraints. (3) The aluminum honeycomb sandwich structures have low thermal conductivity and an uneven temperature distribution. (4) The battery has a special structure, a narrow operating temperature range, and insulation requirements. The above issues pose great challenges to the thermal control design of the entire satellite.

This paper mainly focuses on the above difficulties to develop a communication satellite thermal control design scheme, and the correctness and rationality of the thermal control design have been verified through ground experiments. Firstly, new efficient heat-transfer technologies and thermal control materials for spacecraft are introduced. Secondly, the structure and internal heat source of the satellite are introduced. Thirdly, the external heat fluxes are analyzed, and the position of the heat dissipation surface and extreme conditions are confirmed. Then, a thermal control system is designed around the difficulties of thermal control. With heat pipes, the temperature uniformity of the +Y deck, −Y deck, and +Z deck increased by 8 °C, 9.9 °C, and 34.2 °C, respectively. Furthermore, the maximum temperature of the power controller, secondary power supply, bidirectional frequency converter, and solid discharge decreased by 32.5 °C, 22.0 °C, 14.0 °C, and 164 °C, respectively. Finally, a thermal balance test is performed. The test results show that the temperatures of the solid-state power amplifier, on-board computer, power controller, secondary power supply, and bidirectional frequency converter meet the requirements of the thermal control indices.

2. Introduction to Satellite

2.1. Orbit

One satellite operates in a circular orbit with a track height of 1175 km and an orbit inclination of 86.5°. A three-axis stable attitude towards the ground is adopted by the satellite, with its forward flight direction defined as the +X direction and the direction towards the ground defined as the +Z direction.

2.2. Structure

To accommodate the entire satellite mission payload equipment and fully utilize the carrier envelope space, the satellite body is designed as a regular trapezoidal box structure, including six decks, named +X, −X, +Y, −Y, +Z, and −Z. The satellite structure is composed of aluminum honeycomb structural panels, and the maximum size of the satellite launch state is 140 cm × 175.9 cm × 102.3 cm. The payloads on the satellite mainly include a Ka-transmitting antenna, Ka-receiving antenna, solid-state power amplifier, secondary power, and bidirectional frequency converter. The layout of the devices on the satellite is shown in Figure 1. The devices labeled in Figure 1 are mainly high-power-consumption devices and devices with special temperature requirements. It should be highlighted that some of these high-power-consumption devices have heat-dissipation needs and temperature-compensation needs.

To obtain sufficient lighting conditions, a dual-axis SADA is adopted to drive the solar wing. The A-axis is used to drive the solar wing to rotate with a rotation range from 0° to 360°, parallel to the Y-axis direction of the satellite. The B-axis is used to drive the unilateral solar wing to swing perpendicular to the direction of the solar wing, with a swing range from −75° to 75°.

2.3. Internal Heat Source

The payload of the communication satellite is the largest heat source, with a working time of no more than 30 min per orbit. The total power consumption during operation is 653.7 W, including 448 W for the solid-state power amplifier, 117.3 W for the secondary power, 78.4 W for the bidirectional frequency converter, and 10 W for the Ka-transmitting antenna. The power consumption of the entire satellite is about 800 W, and the high-power devices other than the payload mainly include the power controller (50 W), on-board computer (OBC, 28 W), and S measurement and control machine (12.5 W).

How to dissipate the heat from high-power devices will become a key focus in the thermal design, and the heat dissipation of the solid-state power amplifier will become a key technology.

3. Heat Flux Analysis

The space orbit heat fluxes mainly include three parts: direct solar heat flux, Earth’s albedo heat flux, and Earth’s infrared radiation heat flux [22,23,24,25,26]. Among them, the Earth’s radiation infrared heat flux is only related to the orbit height and does not change much during its lifespan. Therefore, the changes in space heat fluxes are mainly reflected in the direct solar heat flux and Earth’s albedo heat flux, which are caused by the periodic motion of the sun in the ecliptic plane, resulting in the β-angle between the solar light vector and the satellite’s orbital plane changing within a certain range [27,28]. Therefore, the changes in the comprehensive energy of orbit heat fluxes during satellite operation in orbit depend on the β-angle [9,29,30].

The satellite mission cycle is 3 years, and Figure 2 shows the change diagram of the satellite’s β-angle within 7 years. As seen in Figure 3, the β-angle changes between −90° and 90°. When the β-angle is greater than 58 °, the orbit is in the full-day illumination segment.

To determine the direction of the heat dissipation surface and extreme conditions, the external heat flux on each surface of the satellite is calculated when β = ±90°, β = ±58°, and β = 0°. Figure 3 shows the calculation results of the external heat flux on each surface of the satellite when β = ±90°, β = ±58°, and β = 0°.

To comprehensively evaluate the external heat fluxes absorbed by the satellite, the average external heat fluxes of the satellite deck surfaces are calculated separately in the early and late stages of the white paint, as shown in

Table 1. The average external heat fluxes of the satellite deck surfaces (units: W/m²).

Name	β-Angle	+X	+Y	+Z	−X	−Y	−Z	Total
Early	90°	38.7	9.0	147.1	38.9	252.6	0.2	576.5
	58°	84.5	9.3	197.2	87.5	217.5	47.2	701.2
	0°	118.1	47.5	176.1	118.0	48.5	79.9	588.1
	−58°	90.5	217.5	196.7	91.1	9.3	47.9	595
	−90°	39.0	252.6	147.4	39.1	9.0	0.4	397.5
Late	90°	40.3	9.0	152.2	40.4	618.1	0.4	950.4
	58°	154.7	9.7	277.2	162.1	530.4	117.9	1310
	0°	238.6	105.2	224.5	238.3	107.8	199.8	1114.2
	−58°	169.7	530.2	276.0	171.1	9.7	119.7	1218.4
	−90°	41.0	618.1	152.8	41.1	9.0	1.1	773.1

.

From the above Figures and Table, it can be seen that the +Y surface is exposed in the long term to sunlight, resulting in a higher external heat flux, when the β-angle ispostive. But the +Y surface is a shaded area with a smaller external heat flux when the angle is negative. The change in the +Y surface is opposite to the −Y surface. The external heat fluxes on the +Z surface and −Z surface are relatively stable. While the angle is 0 °, the external heat fluxes on the +X surface and −X surface are at their maximum, with significant periodic variations.

Due to the high power of the satellite device, some or all of the outer surfaces of each deck are used as heat dissipation surfaces. Considering that there are moments of high external heat flux on both ±Y surfaces, but they are not exposed to sunlight at the same time, four heat pipes are adopted to connect the ±Y and ±Z heat dissipation surfaces to improve heat dissipation efficiency.

Thermal conduction is the main way of heat transfer between satellite equipment and structural panels and heat pipes, which mainly depends on parameters such as thermal conductivity and heat capacity. Thermal radiation is the main way of heat dissipation between the satellite and the deep cold space, and the satellite thermal simulation of heat dissipation mainly considers the radiation characteristics of the satellite surface, which mainly depends on the emissivity and absorption rate of the surface of the satellite and so on.

Table 2. Thermal control parameters for surface coating.

Protective Layer	Solar Absorptivity		Infrared Emissivity
	BOL	EOL
Aluminum black anodized, E51-M black lacquer	0.85	—	0.85
Aluminum alloy conductive oxidation	0.25	—	0.85
Aluminum conductive oxidation	0.14	—	0.36
Aluminum alloy natural color anodic oxidation	0.55	—	0.7
Double-sided aluminized polyester film	0.12	—	0.05
SR107-ZK white paint	0.17	0.5	0.87
Single-sided aluminized polyimide film	0.35	0.65	0.69
Conductive F46 silver-plated secondary surface mirror	0.13	0.4	0.67

shows the thermal control parameters of the satellite and the surface coatings of each device (for devices for which surface radiation parameters have been given, the data provided by them are used as design inputs). For the modeling of each piece of equipment, the material’s physical parameters are used as equivalent density to ensure that the simulation model equipment is the same weight as the real equipment.

Given the values of the external heat fluxes absorbed by the combined heat dissipation surfaces of the satellite, when the β-angle is 0°, +58°, and −58°, the external heat fluxes absorbed by all surfaces are at a relatively high level, which can be used as a high-temperature condition. When the β-angle is 90°and −90°, the external heat fluxes absorbed by all surfaces are the lowest, which is considered as a low-temperature condition.

4. Thermal Design Scheme

The thermal control system not only needs to maintain laser communication within its specified temperature range, but also provides sufficient heat dissipation for high-power and long-term electronic devices. How to break through the limitations of structure and creatively solve the heat dissipation problem of high-power solid state power amplifiers will become the most critical technology in the thermal design. How to solve the problem of uneven distribution and severe fluctuations of heat fluxes outside the satellite’s surfaces will become a difficult point in the overall thermal design. How to solve the problem of low thermal conductivity and uneven temperature distribution in the aluminum honeycomb sandwich structures will also become a difficult point in thermal design. How to meet the temperature requirements of the battery and also meet its insulation requirements will become a key focus in the thermal design. This article mainly takes the above issues as the starting point to design the entire satellite thermal control system, as follows.

4.1. Thermal Design of Deck

The aluminum honeycomb sandwich structures are used as satellite’s decks, with a total thickness of 25 mm. Some outer part surfaces of the +X deck and −Z deck are covered with 20 units of multi-layer insulation assembly (MLI), whose facial mask is a polyimide single-sided aluminized secondary surface mirror.

All outer surfaces of the ±Y, +Z, and −X, as well as some outer surfaces of the +X and −Z, are used as heat dissipation surfaces, which are sprayed with SR107-ZK white paint. Except for the installation surface, the inner surfaces of the satellite are sprayed with E51-M black paint.

To improve the heat transfer capacity of the deck and flatten the temperature of the deck, 13 heat pipes are pre-embedded inside the deck, with 2 heat pipes embedded in the +Y deck, 2 heat pipes embedded in the −Y deck, 6 heat pipes embedded in the +Z deck, and 3 heat pipes embedded in the +X deck, as shown in Figure 4 (green part).

Considering that there are moments of high external heat flux on both ±Y surfaces, but they are not exposed to sunlight at the same time, four heat pipes are adopted to connect the ±Y and ±Z heat dissipation surfaces to improve heat dissipation efficiency. In addition, a heat pipe is used to connect the −X deck and +Z deck to balance the effect of external heat fluxes, as shown in Figure 4 (blue part).

4.2. Thermal Design of High-Power Devices

High-power devices mainly include the solid-state power amplifier (448 W), secondary power (117.3 W), bidirectional frequency converter (78.4 W), power controller (50 W), and OBC (28 W).

The solid-state power amplifier mainly consists of 16 modules, each with a power consumption of 28 W. Due to their own structural limitations, conventional heat pipes cannot be directly used for heat transfer. To solve the heat dissipation problem, a truss heat pipe is innovatively designed based on its structural characteristics. The heat is transferred to the nearby deck through four truss heat pipes, and finally dissipated through the deck, as shown in Figure 5. The solid-state power amplifier, truss heat pipes, and deck are all installed with thermal conductivity, and the installation interfaces are filled with thermal conductivity fillers.

The secondary power, bidirectional frequency converter, power controller, and OBC are all installed with thermal conductivity on the deck, and the installation interfaces are filled with thermal conductivity fillers. In addition, to the improve heat dissipation capacity, the pre-embedded heat pipes on the deck are designed near the installation positions of the high-power equipment and decks.

4.3. Thermal Design of Active Heating Zone

Active thermal control is implemented to further improve and maintain satellite temperature levels on the basis of passive thermal control, while adopting a heating compensation method to control the reasonable distribution of temperature.

When the payload is turned off, to ensure the normal operation of the satellite, an active heating zone is set up on the satellite deck and equipment with special temperature requirements, including +X deck (4 channels), +Y deck (1 channel), +Z deck (3 channels), −X deck (2 channels), −Y deck (3 channels), −Z deck (1 channel), SADA (2 channels), battery (4 channels), storage tank (2 channels), power controller (1 channel), and S measurement and control machine (1 channel). In addition, to increase reliability, backup heating zones were designed for both the battery and storage unit.

4.4. Thermal Design of Special Devices

Special devices include the battery and storage tank, both of which have high-temperature requirements. The working temperature of the battery is 10~30 °C, and the working temperature of the storage tank is 20~45 °C. Due to the unique characteristics of the battery, insulation design should also be considered during thermal design. A 100 μm polyimide film is set up on the outer side of the battery, and the outer side of the film is covered with 20 MLI, of which the facial mask is a double-sided aluminized polyester film. The contact surface between the battery and the deck should be insulated with polyimide film for installation. In addition, the battery is equipped with 4 active heating zones, including 2 main and 2 backup heating zones.

The thermal insulation between the storage tank and deck is carried out with 5 mm thick polyimide. The storage tank’s outer surface is covered with 20 MLI, whose facial mask is a double-sided aluminized polyester film. Two active heating zones are arranged in the storage tank, including one main active heating zone and one backup active heating zone.

4.5. Thermal Design of Other Devices

To improve the uniformity of temperature, all inner equipment surfaces of the satellite are treated with blackening except for the installation surface. Passive devices and low-power devices are installed directly on the deck. The outer side of the star sensor is covered with 20 MLI to isolate the impact of the space environment.

The momentum wheels and brackets are installed with thermal conductivity, and the brackets and decks are installed with thermal conductivity. The installation interfaces are filled with thermal conductivity fillers.

5. Evaluation of Heat Dissipation Effect of Heat Pipes

To study the heat dissipation effect of heat pipes, two schemes with and without heat pipes are analyzed, mainly evaluating from two dimensions: the temperature uniformity of the deck itself and the temperature of the high-power devices. A thermal analysis model (Figure 6) is established with NX-2022/SST-2021 software, and thermal simulation calculations are performed with and without heat pipes. Due to the limited scope of the embedded heat pipe, only the temperature of the deck around the heat pipe is displayed (Figure 7 and Figure 8), and the temperature difference data of the deck are shown in

Table 3. Statistics of temperature difference between decks.

Name	With Heat Pipes	Without Heat Pipes	${∆\mathit{T}}_2$$-{∆\mathit{T}}_1$
	$\mathbf{Temperature} \mathbf{Difference} \left({∆\mathit{T}}_1\right)$	$\mathbf{Temperature} \mathbf{Difference} \left({∆\mathit{T}}_2\right)$
+Y deck	10.2	18.2	8
−Y deck	14.2	24.1	9.9
+Z deck	29.3	63.5	34.2

. In addition, the highest temperatures of some high-power devices are also calculated, as shown in

Table 4. Maximum temperature statistics for high-power devices.

Name	With Heat Pipes	Without Heat Pipes	T2max − T1max
	Temperature (T1max)	Temperature (T2max)
Power controller	15.0	47.5	32.5
Secondary power	16.0	38.0	22.0
Bidirectional frequency converter	35.0	49.0	14.0
Power amplifier	51.0	215.0	164.0

.

From Table 3, it can be seen that heat pipes can improve the temperature uniformity of the decks. With heat pipes, the temperature uniformity of the +Y deck, −Y deck, and +Z deck increased by 8 °C, 9.9 °C, and 34.2 °C, respectively.

From Table 4, it can be seen that the temperature of high-power devices decreases significantly with heat pipes, and the heat dissipation effect is significantly improved. The maximum temperature of the power controller, secondary power supply, bidirectional frequency converter, and solid discharge decreased by 32.5 °C, 22.0 °C, 14.0 °C, and 164 °C, respectively. In particular, the designed truss heat pipes can effectively dissipate heat for the power amplifier and meet its working requirements.

6. Thermal Balance Test

At present, thermal simulation methods cannot be used to solve the thermal design problems of spacecraft alone. To verify the correctness of the thermal design, thermal balance tests must be conducted [31,32,33]. The thermal balance test is a supplement to and validation of thermal analysis. To obtain temperature distribution data under the working conditions of the satellite and establish an accurate thermal model, it is necessary to conduct thermal balance tests on the satellite.

6.1. Test Scheme

In the thermal balance test of the satellite, a vacuum tank is used to simulate the space environment, and the absorbed heat flux is simulated with an infrared heating cage. During the test, the temperature of the heat sink is required to be below 100 K, and the vacuum degree inside the tank is required to be better than 1.3 × 10⁻³ Pa. Figure 9 is a physical layout of the satellite’s thermal balance test, including the satellite, vacuum tank, and infrared heating cage. The simulation of the heat balance test consists of three main parts, vacuum, low-temperature and black background, and external heat flow. In this test, the vacuum tank is used to simulate the vacuum and low-temperature and black background, and the external heat flux is simulated by the infrared cage. The heat flux (external heat flow) in this paper is divided into different regions, which will be different in different working conditions, and these are imposed according to different working condition definitions. This heat balance test fully follows the requirements of the test and accurately reflects the real orbital conditions from an engineering point of view. The diffraction of waves, which has a limited effect on the heat flux, can be completely ignored.

6.2. Test Conditions

Three extreme conditions are simulated by varying the external heat fluxes, working mode, distribution of internal heat sources, and thermal properties of the thermal control layer. The specific settings are listed in

Table 5. Thermal balance test conditions.

Number	Condition Name	Main Setting Conditions
1	Low-temperature condition (LT)	The β-angle is 90° and the solar constant is 1322 W/m2. The performance parameter of MLI is αs/ε = 0.36/0.69, and the performance parameter of SR107-ZK white paint is αs/ε = 0.17/0.87. The active thermal control works continuously. Platform equipment is long-term operational. Payload turns off.
2	High-temperature condition 1 (HT1)	The β-angle is 58° and the solar constant is 1414 W/m2. The performance parameter of MLI is αs/ε = 0.64/0.69, and the performance parameter of SR107-ZK white paint is αs/ε = 0.5/0.87. The active thermal control works continuously. Platform equipment is long-term operational. Payload turns on.
3	High-temperature condition 2 (HT2)	The β-angle is 0° and the solar constant is 1414 W/m2. The performance parameter of MLI is αs/ε = 0.64/0.69, and the performance parameter of SR107-ZK white paint is αs/ε = 0.5/0.87. The active thermal control works continuously. Platform equipment is long-term operational. Payload turns on.

, mainly including the β-angle, solar constant, performance parameter of the MLI, and performance parameter of the white paint (SR107-ZK). The active thermal control works continuously.

In the test, the KM2 heat sink temperature is required to be below 100 K, simulating the space background of the satellite in orbit; the vacuum degree in the tank is required to be better than 1.3 × 10⁻⁴ Pa, simulating the vacuum environment of the satellite in orbit. Considering the configuration characteristics and working mode of the satellite, an infrared cage is used to simulate the external heat flow in the heat balance test.

6.3. Test Results

According to the designed conditions, the thermal balance test is completed in the order from low temperature to high temperature. The degree of vacuum during the thermal balance test is 6.5 × 10⁻⁴ Pa, and the heat sink temperature is 95 K, which is the same as space’s thermal environment. The test results are shown in Figure 10 as temperature curves of the key components of the satellite under different conditions.

At the same time, according to the test results, the temperature data of the key devices under various condition are also compiled, as shown in

Table 6. Temperature of key devices (units: °C).

Name	LT		HT1		HT2		Thermal Control Indicators
	Min	Max	Min	Max	Min	Max	Min	Max
Star Sensor	−11.4	−9.2	−7.3	−7.1	−8.6	−8.3	−40	45
Gyroscope	−7.2	−3.6	5.8	21.2	3.0	18.0	−20	45
Magnetic Torquer	−5.2	−2.1	5.8	18.8	3.9	16.5	−20	55
Storage Tank	24.4	28.3	24.7	28.6	24.7	28.7	20	45
Power Controller	−10.1	−9.9	5.0	5.2	2.2	4.1	−20	55
S Measurement and Control Machine	−10.1	−6.6	6.7	6.8	2.9	3.3	−20	55
Secondary Power	−11.3	−11.1	5.0	5.1	−2.9	−2.0	−25	45
On-Board Computer	−4.2	−4.0	15.5	15.7	14.6	16.1	−20	55
GNSS Receiver	−3.6	−2.7	21.8	23.0	18.7	19.6	−20	55
Wave Filter	−9.3	−6.9	16.5	19.4	16.5	20.1	−40	80
Bidirectional Frequency Converter	−5.2	−4.7	21.1	23.5	15.1	15.4	−15	55
Ka-Transmitting Antenna	−15.4	−15.2	4.4	4.6	2.6	2.9	−90	90
Ka-Receiving Antenna	−9.1	−8.7	5.8	6.0	4.4	4.7	−90	90
Solid-State Power Amplifier	−8.9	−7.4	−2.7	−2.5	3.0	25.1	−20	50
Battery	19.0	23.1	18.9	22.8	19.0	23.1	10	30
SADA	−8.5	−8.4	2.4	2.6	1.4	1.8	−15	45
Momentum Wheel	−6.0	−4.6	3.9	6.9	1.5	6.2	−20	50

.

From the above Figures and Tables, it can be seen that the temperature of each heating zone can be controlled near the target temperature under low-temperature conditions, indicating that the design of the active heating zone is reasonable. In high-temperature conditions, the temperature of high-power devices, including the solid-state power amplifier, secondary power, bidirectional frequency converter, power controller, and OBC, operates normally, indicating that the heat dissipation surface can effectively dissipate heat and the heat dissipation path is normal. In the test, the temperature of each device can meet the requirement of the thermal control indicator, suggesting that the thermal design scheme is reasonable and feasible.

7. Discussion

In general, passive thermal control measures such as heat pipes, MLI, white paint, black paint, insulation pads, thermal fillers, and blackening treatment are mainly used in the thermal design process for thermal conductivity, insulation, and heat dissipation. Meanwhile, thin-film electric heaters are used for temperature compensation.

To solve the heat dissipation problem of the solid-state power amplifier, a truss heat pipe is innovatively designed based on its structural characteristics, which can be used to transfer its heat to the deck for heat dissipation. A thermal control scheme has been specially developed for a special device—the battery—which can meet both its temperature requirement and insulation requirement. A large number of heat pipes are embedded inside the deck for isothermal design to compensate for the low thermal conductivity of the aluminum honeycomb sandwich structures. The surface of the deck is equipped with coupled surface-mounted heat pipes to balance the external heat fluxes and improve heat dissipation efficiency.

To verify the correctness of the thermal design, the satellite is subjected to a thermal balance test according to the three extreme conditions laid out. The test results show that all devices meet the requirements of the thermal control indicators.

8. Conclusions

A thermal control system is designed based on the orbit, attitude, and thermal control development technology requirements of the satellite so that the satellite can meet its required operating temperature indicators in different orbital environments. The thermal balance test verifies the correctness of the thermal design, which can provide a basis for the subsequent thermal design of communication satellites.

Due to the extremely harsh orbital environment where the satellite is located, this paper has designed its thermal control system according to the worst-case scenario, and the satellite can almost adapt to any low Earth orbit.