Thermal Optimization Design for a Small Flat-Panel Synthetic Aperture Radar Satellite

Abstract

This article introduces a small microwave remote sensing satellite weighing 310 kg, operating in low earth orbit (LEO). It is equipped with an X-band synthetic aperture radar (SAR) antenna, capable of a maximum imaging resolution of 0.6 m. To achieve the objectives of lower cost, reduced weight, minimized power consumption, and enhanced temperature stability, an optimized thermal design method tailored for satellites has been developed, with a particular focus on SAR antennas. The thermal control method of the antenna is closely integrated with structural design, simplifying the thermal design and its assembly process, reducing the resource consumption of thermal control systems. The distribution of thermal interface material (TIM) in the antenna assembly has been carefully calculated, achieving a zero-consumption thermal design for the SAR antenna. And the temperature difference of the entire antennas when powered on and powered off would not exceed 17 °C, meeting the specification requirements. In addition, to ensure the accuracy of antenna pointing, the support plate of antennas requires stable temperature. The layout of the heaters on the board has been optimized, reducing the use of heaters by 30% while ensuring that the temperature variation of the support board remains within 5 °C. Then, an on-orbit thermal simulation analysis of the satellite was conducted to refine the design and verification. Finally, the thermal test of the SAR satellite under vacuum conditions was conducted, involving operating the high-power antenna, verifying that the peak temperature of T/RM is below 29 °C, the temperature fluctuation amplitude during a single imaging task is 10 °C, and the lowest temperature point of the support plate is 16 °C. The results of the thermal simulation and test are highly consistent, verifying the correctness and effectiveness of the thermal design.

1. Introduction

With the rapid advancement of satellite and rocket manufacturing technology, particularly the emergence and application of reusable rockets, the average launch cost of a single satellite has seen a significant reduction, ushering the global aerospace industry into a new era [1,2,3]. A satellite constellation is a system in which multiple satellites work together as a whole, also known as a distributed satellite system (DSS). In DSS, two or more satellites are deployed in the same or multiple orbits to complete specific space tasks through mutual cooperation. Compared to a single satellite, DSS significantly enhances the satellite’s global coverage efficiency, also referred to as temporal resolution [4]. Small- and medium-sized satellites have become the primary components of DSSs, owing to their advantages in terms of lower cost, reduced weight, and shorter development time. LEO is frequently employed as the operational orbit for DSSs, featuring an orbital altitude not exceeding 2000 km. This implies reduced launch costs and offers satellites with high bandwidth and low communication latency. Currently, the scarcity of LEO orbits is intensifying competition among global companies for constellation deployment, leading to a production-centric approach in the design of constellation satellites [5,6].

SAR is an active earth observation system that emits continuous radio pulses towards the target scene, receives and records the echoes of each pulse, and achieves high-resolution microwave imaging. SAR can be mounted on various platforms, including airborne, spaceborne, and missile-borne platforms, and it finds widespread applications in numerous fields. SAR satellites are capable of imaging at night or in areas obscured by clouds, demonstrating all-weather and all-day imaging capabilities that are unachievable by optical remote sensing satellites [7]. Currently, typical small SAR satellite constellations include ICEYE, Capella, StriX, and so on [8,9,10], which almost achieve hourly revisits for a single target. In addition, to provide users with higher-precision data, medium and large SAR satellites are continuously being upgraded, such as Sentinel-1, FIA, CSG, RadarSat, and others [11,12,13,14].

The commonly used SAR satellite antennas include slotted waveguide antennas and microstrip phased array antennas, both of which are of a plate antenna structure. Additionally, there are deployable mesh reflector antennas and deployable paraboloid antennas [15]. For plate-type SAR antennas, a reasonable thermal design is necessary to ensure the temperature stability of high-power electronics and supporting structures. SAR satellites often employ methods such as multi-layer insulation (MLI), thermal control coatings, TIM, heat pipes, and heat conduction devices for their thermal design [16,17]. MLI is a vacuum thermal insulation component made of multiple reflective screens and spacers interlaced. The MLI forms a high thermal resistance to radiant heat flux through the layers of reflective screens, and, theoretically, its thermal conductivity can reach 10⁻⁵ W/(m·K) to achieve good thermal insulation. And TIM is a kind of gap filling material, which is used to increase the heat transfer between two contact surfaces. So far, numerous studies have proposed thermal control methods for SAR satellites, all emphasizing heat dissipation and temperature uniformity in the transmitter/receiver module (T/RM) [18,19].

In terms of thermal design, the on-orbit working time of the payloads on optical remote sensing satellites and SAR satellites is quite similar. Specifically, each orbital cycle only includes short imaging tasks, during which the heat consumption of related electronics is high. Therefore, the excellent thermal design concepts of small optical remote sensing satellites can serve as a reference. To address the heat dissipation issue of the radiation detector for the ChubuSat-2, Daeil Park proposes high-thermal-conductive graphite sheets to be used as a flexible radiator fin for the radiation detector [20]. Graphite sheets have lighter weight and higher thermal conductivity than traditional fins. AM Elshaer has designed an aluminum thermal storage panel for the battery of a small satellite using phase change materials (PCM). This design enables the storage and utilization of equipment heat between high- and low-temperature conditions in orbit, enhancing the temperature stability of the equipment and reducing the overall power consumption of the small satellite [21]. Casper proposed that proper surface treatment of existing structures, such as the payload shell, can directly dissipate the heat from high-power components, thereby avoiding increased thermal coupling and more complex hardware stacking, ultimately contributing to reducing the size and weight of small satellites [22]. In addition, it is necessary to conduct reasonable tests to verify the correctness of the thermal design and the effectiveness of the onboard electronics [23,24,25].

This paper introduces a 310 kg high-resolution microwave remote sensing satellite equipped with an X-band SAR antenna, capable of achieving a maximum imaging resolution of 0.6 m. To ensure the normal operation and imaging quality of the satellite, some thin-film heaters are used for temperature control to ensure that the temperature of the antenna support plate is between 15 and 25 °C, and to ensure that the temperature of the T/RMs is not too low. To fulfill the design concept of reduced weight and lower power consumption, while ensuring the temperature stability of the SAR antenna [26], several strategies have been proposed, encompassing both the SAR antenna and the satellite platform. The SAR antenna is a microstrip phased array antenna, adopting a structure–thermal integration design. The use of a high thermal conductivity structural substrate and a uniform component layout ensures the temperature consistency of the T/RM. This paper reasonably calculates the effective area of the TIM, controls the overall temperature level of the antenna, and ensures its operating temperature with zero-power consumption. The SAR antenna is connected to the satellite platform through a structural plate, which is called a support plate. To ensure that the temperature variation of the antenna support plate remains within 5 °C and to minimize the utilization of heaters, an optimized layout method for heaters has been proposed, and the maximum spacing for heater placement has been calculated. As the satellite is an initial version, some additional heat pipes are added to ensure its higher reliability, which is not necessary for the design of subsequent satellites.

In the next generation SAR satellites, in order to achieve a longer imaging time, the fluid loops will be used to dissipate the heat of the working SAR antenna. Fluid loops have the ability to transport heat in large capacity over long distances. Moreover, when the antenna is not working, the fluid circuit could also transport the waste heat of the satellite platform to keep the antenna warm.

This paper also elaborates on the thermal vacuum (TVAC) testing process of the SAR satellite. During the test, high-power emission of the satellite is achieved through an absorber box, which can be positioned within the vacuum chamber and possesses temperature control capabilities. This ensures the integrity of the satellite development process [27].

2. Satellite Overview

Table 1 displays parameters of the satellite. The satellite has an orbit altitude of 515 km, with its SAR antenna operating in the X-band frequency. It supports stripmap modes with resolutions of 0.6 m, 1 m, and 2 m, as well as a ScanSAR mode with a resolution of 10 m. In the stripmap mode with a resolution of 2 m, the imaging duration for each orbit cycle of the satellite can exceed 100 s. The satellite maintains an attitude in which its SAR antenna points towards the Earth during standby mode and swings sideways by 20° to 40° during imaging.

Figure 1 illustrates the layout of the satellite, which appears flat overall. The SAR antenna is positioned at the top of the satellite platform, which consists of six external panels and two internal panels. The SAR antenna is exclusively connected to the +Z side support plate, and the satellite platform design prioritizes the SAR antenna, eliminating the movable components like hinges and support rods. This ensures superior flatness of the antenna plane, thereby attaining higher imaging quality. The satellite flies in the +X direction in orbit, which has a very small windward area, minimizing the flying resistance of the satellite. Designing the platform with the payload as the center will result in the platform size being close to the load, and the electronics distribution within the platform is loose. This provides possibilities for optimizing the layout design of the heater.

The SAR antenna is divided into three arrays, each with identical layout and function. The array is composed of stacked layers, featuring twenty-four radiation units (RUs) on the outermost (+Z side) of the SAR antenna and an antenna mounting plate with an aluminum alloy honeycomb structure in the middle. Inside the antenna, there are several devices, including twenty-four T/RMs paired with RUs, and eight power modules (PMs), each responsible for energy distribution to three T/RMs. Each array is controlled by a wave position control unit (WPCU). Both T/RM and PM have high heat consumption during imaging. In addition, if the temperature difference between T/RM components on the SAR antenna is beyond 5 °C, the imaging quality is adversely affected.

To achieve superior anti-vibration properties and enhanced on-orbit attitude control capabilities, the components on board are positioned as close to the center of mass as feasible. Two internal compartments divide the platform into three sections. The middle section houses a digital control unit (DCU), a power conditioning and distribution unit (PCDU), and a reaction wheel (RW), all of which exhibit significant long-term heat consumption. On both sides of the platform, propulsion systems (PSs) and batteries are arranged, which have a narrow temperature adaptation range. Additionally, there are antenna control units (ACUs), data storage units (DSUs), and data transmission control units (DTCUs), which are devices related to imaging and data transmission tasks and exhibit short-term heat consumption. The navigation receiving unit (NRU) and the measurement and control unit (MCU) have low heat generation.

3. Thermal Design Strategy

The primary goal of thermal design is to ensure that the SAR antenna can capture images for over 100 s in the stripmap mode with a resolution of 2 m within each orbit cycle. At the same time, the temperature difference of all T/RMs on the SAR antenna should be within 5 degrees Celsius at all times. In addition, under extreme space environments and satellite operating modes, the thermal control system (TCS) must ensure that all components remain within the permissible temperature range (Table 2), while minimizing the power consumption of the heaters.

The surface of RUs needs to maintain a flatness of ±2 mm; however, thermal deformation is a significant factor affecting the flatness. During the assembly, integration, and testing (AIT) process of the SAR satellite, the ambient temperature stands at 20 °C. To eliminate the temperature variation of the structure between its in-orbit and AIT states, which could potentially degrade the shape and positional accuracy of the antenna mounting surface, it has been decided to maintain the +Z plate’s in-orbit temperature at 20 °C. Furthermore, a temperature control accuracy of 5 °C is proposed to avoid excessive thermal deformation caused by uneven temperatures at the mounting points.

3.1. Overall Thermal Design

Taking into account the satellite layout and thermal control objectives, we have determined the basic strategy for thermal design. The planned overall heat transfer path of the satellite is illustrated in Figure 2. Let us begin with the antenna’s support plate. The SAR antenna located above the support plate exhibits instantaneous high heat consumption and must adhere to temperature consistency requirements. Consequently, MLI is strategically placed on the Z-side and in the circumferential direction of the antenna, aiming to minimize disturbances from the external environment on that side. (Given that the Z-side of the SAR antenna is larger than the support plate, it is susceptible to the space thermal environment). The heat generated on the antenna is radiated to the space environment through the +Z surface, which is coated with polyimide germanium film and has good wave transmission capability, providing necessary protection for the antenna. Place 20 units of MLI between the support plate and antenna to minimize the impact of the SAR antenna on the temperature control of the support plate. The −Z direction of the support plate is oriented towards the satellite platform. Compared to SAR antennas, the platform offers greater flexibility in arranging heat dissipation surfaces. Through reasonable thermal design, a stable temperature level can be maintained inside the platform. Therefore, an MLI with only 1/4 of the conventional thickness is arranged between the support plate and the platform. By arranging appropriate coating radiators on the +Y side, −Y side, and −Z side of the satellite platform to maintain an average temperature of 20 °C, the required power of the heaters for the support plate can be less. For PS and the battery, their sections are relatively spacious, and the panels are wrapped in MLI, making their thermal control relatively straightforward.

3.2. Methods to Reduce Heaters

The dimensions of the support plate are 0.78 m × 2.77 m, and it needs to be maintained at 20 °C. A dense arrangement of heaters and thermostats are capable of achieving a temperature control of 20 °C; however, it is not an optimized design and will inevitably result in a waste of satellite resources. The improved method is presented in Figure 3. The connection points between the support plate and the SAR antenna, as well as other panels, are located at its circumferential positions, resulting in significant heat leakage at these locations. Twelve groups of heaters are arranged in the circumferential direction of the support plate, with a controlled temperature of 20 °C, leaving the center area of the support plate blank (without heaters). Assuming that both the antenna and the platform are in a low-temperature state, we take the longitudinal section of the support plate and simplify it into an infinite flat plate model; that is, assuming there is no temperature gradient in the thickness and width directions within the analysis region, it can be regarded as a one-dimensional model with a single-direction temperature gradient. The temperature at a distance of x from the edge of the heater is defined as T_x, so the heat conduction differential equation is expressed as follows [28].

(1)${\displaystyle \frac{d^2{T}_x}{d{x}^2}}={\displaystyle \frac{[K_{equ}(T-T_{outu})+K_{eqd}(T-T_{outd})]}{{\delta }_p{\lambda }_p}}$

(2)$K_{eq}={\epsilon }_p\sigma (T_0^2+T_{out}^2)(T_0+T_{out})$

where $K_{eq}$ represents the equivalent heat transfer coefficient of the plate surface, T_out is the ambient temperature outside the support plate, δ signifies the thickness, λ stands for thermal conductivity, ε indicates infrared emissivity, σ represents the Stefan–Boltzmann constant, the subscript u designates the upper surface of the flat plate, while the subscript d designates the lower surface, the subscript p designates the support plate, and T₀ represents the temperature of heaters. The boundary conditions of this model can be described as follows.

(3)$x=0,T_x=T_0;x=0.5d_{\max },{\displaystyle \frac{dT}{dx}}=0$

In the formula, d_max represents the maximum distance left blank. The excess temperature θ_x is defined as follows.

(4)$\theta =T_x-{\displaystyle \frac{K_{equ}{T}_{outu}+K_{eqd}{T}_{outd}}{K_{equ}+K_{eqd}}}$

Formula (1) can be rewritten as follows:

(5)${\displaystyle \frac{d^2\theta }{d{x}^2}}=n^2{\theta }_x$

(6)$n=\sqrt{{\displaystyle \frac{K_{equ}+K_{eqd}}{{\delta }_p{\lambda }_p}}}$

n is a constant. The boundary conditions of the model are described as follows.

(7)$x=0,{\theta }_x={\theta }_0=T_0-{\displaystyle \frac{K_{equ}{T}_{outu}+K_{eqd}{T}_{outd}}{K_{equ}+K_{eqd}}};x=0.5d_{\max },{\displaystyle \frac{dT}{dx}}=0$

Solving differential equations yields the following results.

(8)${\theta }_x={\displaystyle \frac{{\theta }_0}{ch(nx)}}$

By utilizing the aforementioned analytical method, we can determine the maximum distance d_max for leaving blank space when the lowest temperature at the center of the support plate exceeds 15 °C. This approach reduces the heaters layout area by 30%, indicating that more heaters and thermostats can be allocated towards ensuring the temperature of other components, such as batteries.

3.3. Optimization of Thermal Design for the SAR Antenna

For large-area antennas, it is of great significance to adopt simplified thermal control measures as much as possible, as any simplification will have an exponential impact on the entire antenna manufacturing process. Each antenna subarray is equipped with four columns of T/RMs, all with the same column spacing. Additionally, two PMs are positioned on one side of each column of T/RMs, ensuring consistent boundary conditions for each column of T/RMs.

The antenna mounting plate is made of an aluminum skin–aluminum core honeycomb board, ensuring both low weight and high strength, while maintaining excellent planar and longitudinal thermal conductivity. This enables the heat generated by the components to be quickly transferred to the mounting board and dispersed after the imaging process. In addition, when MLI is uniformly arranged in the circumferential direction, the boundary of the antenna mounting plate can be considered as an adiabatic state, and the temperature variation across the entire plane can be regarded as a quasi-static process. Due to the hysteresis effect of heat transfer, T/RMs remain unaffected by the temperature rise of PM during the imaging process. In addition, a double-channel aluminum–ammonia heat pipe is positioned beneath each column of T/RMs, serving to mitigate edge effects and reduce temperature disparities arising from inconsistent heat generation among the T/RMs. Figure 4 shows an overview of the antenna mounting plate.

There are gaps between the antenna mounting plate, T/RMs, and RUs, and heaters can be arranged within these gaps to ensure thermal uniformity of mounting plate. By reasonably designing the TIM within the gap, the temperature of T/RMs can be maintained in an ideal state. Specifically, no heating is required for the antenna under low-temperature conditions of the whole satellite, and the peak temperature of T/RMs under high-temperature conditions remains below the maximum permissible value, with a certain temperature margin maintained under both conditions. This design reduces satellite energy consumption which is called zero-consumption thermal control. To design an appropriate area for TIM, in this article, the power consumption of short-time imaging is averaged across the entire orbital period. Assuming that the average temperature (T_avg) of T/RMs under extreme high-temperature conditions is 20 °C, and defining the control temperature (T_con) under extreme low-temperature conditions as −10 °C, the area of TIM can be determined based on the thermal equilibrium equation.

Taking a single RU as the research object, and assuming there is little area of contact between the mounting plate and RU, the thermal equilibrium equation of the T/RMs is as follows.

(9)${\displaystyle \frac{Q_{-zh}-\sigma T_{avg}^4}{1+\frac{1-{\epsilon }_{MLI}}{A_{RU}{\epsilon }_{MLI}}+\frac{1}{A_{T/RM}{X}_{T/RM-MLI}}+\frac{1-{\epsilon }_{T/RM}}{A_{T/RM}{\epsilon }_{T/RM}}}}+{\displaystyle \frac{T_{RUh}-T_{avg}}{\frac{1}{h_{T/RM-MP}}+\frac{{\delta }_{MP}}{{\lambda }_{MP}}+\frac{1}{h_{RU-MP}}}}={\displaystyle \frac{P\tau }{{\tau }_0{A}_{T/RM}}}$

(10)${\displaystyle \frac{Q_{-zc}-\sigma T_{con}^4}{1+\frac{1-{\epsilon }_{MLI}}{A_{RU}{\epsilon }_{MLI}}+\frac{1}{A_{T/RM}{X}_{T/RM-MLI}}+\frac{1-{\epsilon }_{T/RM}}{A_{T/RM}{\epsilon }_{T/RM}}}}+{\displaystyle \frac{T_{RUc}-T_{con}}{\frac{1}{h_{T/RM-MP}}+\frac{{\delta }_{MP}}{{\lambda }_{MP}}+\frac{1}{h_{RU-MP}}}}=0$

where A represents the thermal radiation area, T stands for the temperature of the components, Q represents the space heat flux, X denotes the radiation angle coefficient, and h represents the thermal conductance of the interface, which is a parameter proportional to the area it occupies. P stands for the heat consumption of a single T/RM, τ represents the working duration of T/RMs within one orbital period, τ₀ signifies the duration of one orbital period, the subscript −z represents the antenna −Z side, the subscript h designates extreme high-temperature conditions, the subscript c indicates extreme low-temperature conditions, and the subscript MP refers to the antenna mounting plate.

RU is composed of multiple thin plates, where heat is exchanged between the plates via radiation. A layer of polyimide germanium-coated film is situated on the outer surface of RU. The thermal equilibrium equation for RU is as follows.

(11)${\displaystyle \frac{T_{RUh}-T_{avg}}{\frac{1}{h_{T/RM-MP}}+\frac{{\delta }_{MP}}{{\lambda }_{MP}}+\frac{1}{h_{RU-MP}}}}={\displaystyle \frac{Q_{+zh}-\sigma T_{RUh}^4}{\frac{1}{{\epsilon }_{RU1}}+\frac{1}{{\epsilon }_{RU2}}+\cdots +\frac{1}{{\epsilon }_{RUn}}+1}}$

(12)${\displaystyle \frac{T_{RUc}-T_{con}}{\frac{1}{h_{T/RM-MP}}+\frac{{\delta }_{MP}}{{\lambda }_{MP}}+\frac{1}{h_{RU-MP}}}}={\displaystyle \frac{Q_{+zc}-\sigma T_{RUc}^4}{\frac{1}{{\epsilon }_{RU1}}+\frac{1}{{\epsilon }_{RU2}}+\cdots +\frac{1}{{\epsilon }_{RUn}}+1}}$

In the formula, the subscript +z represents the antenna’s +Z side.

By solving Equations (9)–(12) simultaneously, h_T/RM-MP and h_RU-MP can be obtained. Subsequently, the adapted area of the TIM can be obtained.

The derivation process of Formula (4) is as follows.

Under extreme high-temperature conditions of the satellite, the operation of T/RMs is intermittent, and the temperature fluctuates. The temperature fluctuation amplitude (∆T_h) of T/RM can be estimated according to Equation (13).

(13)$\Delta T_h={\displaystyle \frac{P\tau C_{MP}+0.5h_{1}P{\tau }^2}{C_{MP}{C}_{T/RM}+0.5h_{T/RM-MP}\tau (C_{MP}+C_{T/RM})}}$

In the formula, C represents the heat capacity, and C_MP denotes the average heat capacity of the antenna mounting plate per individual T/RM. Following detailed design parameter iteration, the calculated value of the fluctuation amplitude ∆T_h is 17 °C.

After the aforementioned optimization analysis, it is evident that T/RMs can achieve zero-heating-power thermal control under reasonable TIM settings, while ensuring their temperature remains within the required range.

Other components, such as PM and WPCU, do not necessitate special thermal design methods and can be installed using conventional thermal conductivity methods. The reason is that, compared with T/RM, PM and WPCU have high heat capacity, a larger installation area, and lower heat consumption.

4. Thermal Analysis

The satellite orbit is at an altitude of 515 km, with an inclination angle of 97.42° and a local time of 6:00 AM at the descending node. The beta angle of the orbit varies between 59° and 87° annually. This implies that most of orbital cycles are in full sunlight, with a small number experiencing brief periods of ground shadow (Figure 5). This enables the solar array to harvest sufficient energy even when the SAR antenna remains pointed towards the Earth for a long time, thereby reducing the satellite’s attitude adjustment time during and after imaging and data transmission.

Statistics were compiled on the orbital heat flux of the satellite in all directions (Table 3), solely using the Beta angle as the comparison variable.

According to the thermal design strategy, the antenna is primarily influenced by the heat flux in the +Z direction, whereas the platform is predominantly affected by the heat flux in the +Y, −Y, and −Z directions. Owing to the constant temperature of the support plate, the antenna and the platform are fully thermally decoupled. It is evident that the extreme operating conditions of the antenna and the platform are not aligned. Therefore, the satellite’s extreme high- and low-temperature operating conditions are determined by combining the orbits of the two Beta angles, which, in reality, does not occur.

Table 4 outlines the assumed extreme operating conditions for the orbit and the power consumption of the satellite in numerical thermal analysis. Apart from the entries listed in the table, the disparities between the two extreme operating conditions also encompass differences in the surface coating, variations in the shunt of the solar array, and differences in the satellite’s attitude. The extreme working conditions of the satellite are the superposition of various conditions, such as the power consumption and heat flux in space. These environmental conditions are applied in simulation analysis and thermal test verification. The thermal design should ensure that the temperature margin of each component on the board is at least 5 °C under extreme conditions.

Figure 6 illustrates the thermal analysis model for the satellite antenna and the platform, which are independently modeled and simulated. During the modeling process, non-critical areas are simplified. In the analysis model, the satellite platform utilized 9843 nodes, while the antenna employed 8649 nodes. When conducting thermal analysis of the antenna, the temperature distribution of the platform has no impact on the antenna, as the temperature of the support plate remains constant. On the other hand, when conducting thermal analysis of the platform, the temperature of the antenna influences the thermal control power of the support plate. Therefore, under extreme low-temperature conditions of the platform, a constant temperature boundary of −10 °C is set for the antenna, while under extreme high-temperature conditions, a constant temperature boundary of 20 °C is set for the antenna.

Figure 7 illustrates the simulated temperatures of various components on the SAR antenna under extreme high-temperature conditions, including the coldest and hottest points on the TR/M, PM, WPCU, and the antenna mounting plate of all arrays. The figure depicts a working condition where imaging tasks are performed for 10 orbital cycles, followed by a period of 5 orbital cycles without any imaging tasks. The antenna attains thermal equilibrium after the 7th cycle, indicating that the temperature ceases to rise further. The peak temperature of the T/RM reaches 33 °C, and the temperature fluctuation amplitude during a single imaging task is 16 °C. Notably, the temperature consistency of all components, including the T/RM, surpasses the design specifications by maintaining a variance of less than 3 °C. Upon comparing the four sets of curves, it is evident that, despite variations in the temperature rise among components during the imaging process, all components are capable of rapidly cooling down to temperatures comparable to that of the antenna mounting plate post-task. This is primarily attributed to the exceptional thermal diffusion capabilities exhibited by the mounting plate. It should be noted that during the establishment of the numerical model, the lumped parameter method was applied to each component, simplifying them into a sheet-like structure while neglecting internal details. The thermal load was directly applied to the shell of each component, which deviates from the actual location of the thermal load. Consequently, there may be slight discrepancies between the simulated and measured results regarding the temperature peak.

Figure 8a illustrates the temperature distribution of the support plate under low-temperature conditions. The temperature level remains within the design specifications. The lowest temperature point is observed at the center of the structure, reaching a minimum of 15.4 °C, which aligns with the heater layout. Figure 8b illustrates the simulated temperature range of the platform components under extreme operating conditions (represented by the red section), while also presenting the corresponding permissible temperature range (represented by the gray section). Thanks to the reasonable arrangement of heat dissipation surfaces, the temperature of the components inside the platform is maintained at approximately 20 °C. This ensures ample design margin for each device while minimizing the thermal control power required for the support plate. Under extreme high-temperature conditions, the total thermal control power of the support plate is 31.8 W, while under extreme low-temperature conditions, this power increases to 60.2 W, which occupies little power of the satellite to maintain the support plate at 20 °C.

5. Verification of Thermal Design

To verify the correctness of the thermal design, TVAC testing was conducted on the SAR antenna and the satellite, respectively. Temperature-controlled absorbing devices were employed to absorb the high-power microwaves emitted by the antenna, ensuring that the equipment was not damaged. Additionally, these devices also served the purpose of simulating space heat flux.

Figure 9 demonstrates the space heat flux simulation method employed for TVAC testing. The SAR antenna is positioned facing a device equipped with absorbing materials, and heaters are affixed to the underside of the device to mimic the space heat flux along the normal direction of the antenna plane (i.e., the satellite’s +Z direction). The thermal test is carried out in a vacuum chamber with controllable temperature, where a high-emissivity black coating is applied to the chamber walls. By adjusting the wall temperature, the chamber can be transformed into a heat flux simulation device [29]. During the thermal testing of the antenna, the heat flux in the Z-direction of the antenna is simulated utilizing this particular approach. During the TVAC testing of the entire satellite, an infrared light array was employed as an additional measure to simulate both the direct solar radiation and the heat flux generated by the solar arrays on the satellite.

Since the satellite is mounted upside down on the absorbing device via a supporting structure, the thermal deformation of the structure during thermal testing may potentially impact the satellite. Consequently, temperature control measures were implemented on the supporting structure to maintain them close to ambient temperature.

Figure 10 exhibits the temperature curves of 24 T/RMs during the satellite TVAC testing. When comparing the results with Figure 7a, it is observed that under identical thermal conditions as those used in the simulation analysis, the temperature trough of the T/RM is approximately 2 °C higher, while the temperature peak is roughly 4 °C lower. This discrepancy, which arises due to the simplifications made in the simulation analysis, poses no significant impact on the satellite’s operational performance in orbit. During the imaging process, the maximum temperature difference among T/RMs is 3.6 °C, exceeding the simulation analysis results. This discrepancy arises from the fact that the implementation of thermal control measures falls short of the ideal outcomes anticipated during thermal analysis.

Based on the testing results, the energy consumption for temperature control of the antenna is indeed zero, attributable to the rational design of the TIM usage area. This is called a zero-power thermal design approach.

Figure 11 illustrates the temperature curve on the support plate during the TVAC testing. Point 1 in the figure represents the temperature measurement point near the DCU and PCDU. Due to the operating components, the temperature of Point 1 is slightly higher than 20 °C under high-temperature conditions. Point 2 denotes the temperature measurement point at other undisturbed positions, while Point 3 and Point 4 are the temperature measurement points at the center of the support plate, which exhibit lower temperatures. Overall, the temperature of the support plate meets the temperature difference requirement of less than 5 °C, and the temperature curve in Figure 11a aligns with the simulated temperature.

Table 5 displays the equilibrium temperature of platform components during TVAC testing, where the temperatures of multiple components of the same type are grouped together. Compared to Figure 7b, the testing results exhibit a high degree of agreement with the simulation analysis, achieving a temperature prediction error of less than 3 °C. This demonstrates that each component possesses relatively accurate thermal condition inputs. Furthermore, this achievement relies on the highly implemented thermal control components and precise space heat flux simulation methods utilized during the thermal testing.

Overall, the temperature results for T/RMs and other components on the satellite during the TVAC testing are quite satisfactory. During the test, auxiliary temperature measurements is arranged on the antenna mounting plate. Upon observing the temperature variations, it became evident that the aluminum honeycomb mounting plate exhibited superior thermal diffusion capabilities than anticipated. Consequently, the heat pipe positioned beneath the T/RMs may not be essential, and we intend to simplify the design in subsequent satellites.

6. Conclusions

In this study, a thermal design strategy has been proposed for a high-resolution microwave remote sensing satellite, weighing 310 kg and orbiting in LEO. This strategy aims to achieve low weight, minimize energy consumption, and ensure high thermal stability. Utilizing a stacked antenna structure, an integrated structure–thermal design is employed for the antenna. This approach involves arranging active and passive thermal control elements within the structural gaps, along with the utilization of TIM to regulate the temperature levels of the antenna, thereby achieving a zero-power thermal design for the antenna. Simultaneously, it is imperative to maintain the temperature of the antenna support plate on the satellite platform at a constant 20 °C to minimize thermal deformation of the antenna. By adopting optimized design methods, we have successfully reduced the number of heaters required by 30%. Furthermore, by designing the temperature of the platform to hover around 20 °C, we have minimized the power consumption for thermal control of the support plate. A thermal simulation analysis was conducted on the satellite, revealing that the TCS can effectively ensure imaging in the typical 2 m stripmap mode for 100 s during each orbit cycle, even under extreme in-orbit conditions. Notably, all onboard components remain within the permissible temperature range, and the temperature variance among the core components of the antenna, specifically the T/RMs, remains minimal. The utilization of temperature-controlled microwave-absorbing devices guarantees the successful conduct of TVAC testing for high-power launches of the satellite. During the testing, the peak temperature of the T/RMs reached 28 °C, with a temperature variance not exceeding 3.6 °C, thus validating the efficacy of the thermal design strategy. The testing results also confirm that the thermal control design of the antenna indeed does not incur any energy consumption. This research approach will facilitate the simplification of thermal design for small SAR satellites and communication satellites with comparable configurations.

Author Contributions

Conceptualization, T.B., L.K. and Y.Z.; methodology, T.B. and Y.Z.; software, Y.Z.; validation, T.B., Y.Z. and J.Y.; formal analysis, T.B. and Y.Z.; data curation, T.B. and Y.Z.; writing—original draft preparation, T.B. and Y.Z.; writing—review and editing, T.B., L.K. and H.A.; supervision, L.K., L.Z. and H.A.; project administration, L.K. and H.A.; funding acquisition, J.Y. All authors have read and agreed to the published version of the manuscript.

Data Availability Statement

Data are available on request from the authors.

Conflicts of Interest

Authors Tian Bai, Yuanbo Zhang, Lin Kong, Jisong Yu and Lei Zhang were employed by the company Chang Guang Satellite Technology Co., Ltd. The remaining authors declare that the re-search was conducted in the absence of any commercial or financial relationships that could be construed as a potential conflict of interest.

Footnotes

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Figure 1. Structural layout of SAR01A: (a) overview of in-orbit status, (b) view from −Y side (PS: propulsion system; DTCU: data transmission control unit; DSU: data storage unit; ACU: antenna control unit; DCU: digital control unit; PCDU: power conditioning and distribution unit; RW: reaction wheel; NRU: navigation receiving unit; MCU: measurement and control unit).

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Figure 2. Heat transfer pathway. In which dotted arrows stands for the direction of heat transfer.

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Figure 3. Temperature control strategy for support plate.

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Figure 4. Thermal design strategy for antennas.

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Figure 5. The orbit schematic (from the perspective of the sun).

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Figure 6. Thermal analysis model: (a) antenna; (b) platform.

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Figure 6. Thermal analysis model: (a) antenna; (b) platform.

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Figure 7. The simulated temperature of the antenna: (a) T/RM, (b) PM, (c) WPCU, and (d) antenna mounting plate.

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Figure 7. The simulated temperature of the antenna: (a) T/RM, (b) PM, (c) WPCU, and (d) antenna mounting plate.

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Figure 8. The simulated temperature for the satellite: (a) support plate; (b) platform component.

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Figure 8. The simulated temperature for the satellite: (a) support plate; (b) platform component.

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Figure 9. Heat flux simulation system with absorbing material.

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Figure 10. T/RMs’ temperature during TVAC testing.

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Figure 11. Temperature curve of the support plate: (a) cold case; (b) hot case.

Table 1

System parameters of the SAR satellite.

Item	Parameter
Operating orbit	515 km SSO
Platform size	3.17 m × 1.24 m × 1.7 m
Antenna size	3.17 m × 1.06 m
Weight	310 kg
Operating frequency	X-band
Resolution @ imaging width	Stripmap: 0.6 m@5 km
	Stripmap: 1 m@10 km
	Stripmap: 2 m@20 km
	ScanSAR: 10 m@100 km
Imaging duration	>100 s@2 m (Stripmap)
Lifespan	>5 years

Table 2

Permissible temperature limit for components.

Components List		Permissible Temperature [°C]
Category	Item	Tmin	Tmax
SAR antenna	T/RM	−20	45
	PM	−20	45
	WPCU	−20	45
Platform	PS	10	30
	Battery	10	30
	DSU	−10	45
	DCU	−10	45
	MCU	−10	45
	RW	−10	45
	PCDU	−10	55
	DTCU	−20	45
	ACU	−20	55
	NRU	−20	55

Table 3

Orbital conditions and heat flux statistics.

Item	Beta Angle 59°		Beta Angle 87°
Solar constant [W/m2]	1367
Albedo	0.3
Planetary flux [W/m2]	239.2
	Solar	Infrared radiation	Solar	Infrared radiation
+X side heat flux [W/m2]	221.8	52.3	29.6	54.6
−X side heat flux [W/m2]	206.8	50.4	29.8	48.9
+Y side heat flux [W/m2]	10.9	57.6	0.1	58.1
−Y side heat flux [W/m2]	926.8	60.2	1380.1	58.6
+Z side heat flux [W/m2]	140.6	204.0	37.8	203.5
−Z side heat flux [W/m2]	41.6	1.0	27.8	2.1

Table 4

Details of extreme operating conditions.

Category	Item	Hot Case	Cold Case
Orbital heat flux	Solar constant [W/m2]	1414	1323
	Albedo	0.3	0.3
	Planetary flux [W/m2]	247.4	231.5
	Extreme orbital parameter for antenna	Beta angle 59°	Beta angle 87°
	Extreme orbital parameter for platform	Beta angle 87°	Beta angle 59°
Component power consumption	T/RM [W]	39.6 (Imaging)	-
	PM [W]	33.6 (Imaging)	-
	WPCU [W]	9 (Imaging)	-
	Battery [W]	476 (Imaging)	3 (Constant)
	DSU [W]	86 (Imaging)55 (Transmission)	-
	DCU [W]	25 (Imaging)25 (Transmission)	15 (Constant)
	MCU [W]	9.5 (Constant)	5.5 (Constant)
	RW [W]	81 (Maneuver)	2.5 (Constant)
	PCDU [W]	108 (Imaging)	60 (Constant)
	DTCU [W]	16 (Transmission)	-
	ACU [W]	80 (Imaging)	-
	NRU [W]	8 (Constant)	2.5 (Constant)

Table 5

Platform components’ temperature during the TVAC testing.

Equipment	Hot Case [°C]	Cold Case [°C]
PS	17.0~19.2	16.7~17.2
Battery	15.1~16.8	13.9~14.3
DSU#1#2	11.4~17.4	0.3~4.2
DCU	32.2~33.3	20.1~20.7
MCU	23.3~24.0	9.4~9.7
RW#1#2#3#4	22.6~24.9	2.5~6.2
PCDU	30.4~32.3	25.1~25.6
DTCU#1#2	13.8~21.2	2.8~4.3
ACU	15.3~19.4	1.9~2.1
NRU	13.8~14.4	−1.5~−1.3

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