Thermal control of satellites

Abstract

The present invention relates to a satellite comprising a satellite body including at least one radiation surface configured to radiate heat of the satellite into outer space, and an attitude determination and control (ADCS) system for controlling the orientation of the satellite orbiting around the Earth, wherein the ADCS system is configured to be able to change the orientation of the satellite with respect to the sun between a first position where the radiation surface is away from the sun and a second position where the radiation surface is exposed to the sun. The present invention further relates to a method for controlling heat radiation heat transfer of an artificial satellite.

Description

Background Art

[0001] This application relates to the thermal control of spacecraft, particularly artificial satellites.

[0002] In the design of a spacecraft, the role of the thermal control system (TCS) is to keep all component systems of the spacecraft within the allowable temperature range during all mission phases. An allowable temperature range that must be maintained is defined for all components of the spacecraft to meet the operating requirements. Exposure to extreme temperatures can cause damage and significant impacts on performance.

[0003] Since satellites in orbit operate in a vacuum, the temperatures of hot spots and cold areas are not balanced by the convective heat flux of air. Therefore, extreme temperatures are a very difficult problem for satellites. Satellites are exposed to extreme temperatures in space due to large changes in the external environment, such as the extreme cold in the shadow of deep space and the intense heat of direct sunlight. Furthermore, satellites orbiting in space generate heat according to the power consumption and efficiency of their components.

[0004] Temperature can be adjusted by passive or active thermal management techniques. Passive thermal management techniques aim to maintain the temperature of components without using a power source. Such systems are generally low-cost, but they increase in volume and weight. In conventional satellite thermal control systems, the panels of the satellite body may be used as radiators to release excess heat by thermal radiation heat transfer. However, as the requirements for thermal control increase, the body panels become larger, which is a factor leading to an undesirable increase in the size of the satellite. Equipping a small spacecraft with a radiator and dissipating excess heat by thermal radiation heat transfer can be difficult due to volume constraints.

[0005] Active thermal control methods generally require power for operation and are thus more difficult to incorporate into small satellites in terms of power, mass, and volume. Therefore, the scenarios where active thermal control methods can be used in small spacecraft are limited. However, these methods are generally more effective than passive types in more precisely managing the temperature of components. Representative active thermal control methods that can be mounted on large-scale spacecraft include electric resistance heaters, cryocoolers, thermoelectric coolers, fluid loops, etc. For example, a resistive electric heater controlled by a thermostat is used to keep the temperature of equipment above the lower limit during the cold period of a mission.

[0006] Some known thermal control technologies may not be applicable to small spacecraft due to their surface area, volume, and power constraints. Therefore, there is a need for a thermal management system and method applicable to small spacecraft.

[0007] Some embodiments of the present invention described below solve some of these problems. However, the present invention is not limited to solving these problems, and some embodiments of the present invention solve other problems.

Summary of the Invention

[0008] This summary is provided to present, in a simplified form, a selection of concepts that are further described in the following detailed description. This summary is not intended to identify key features or essential features of the claimed subject matter, nor is it intended to limit the scope of the claimed subject matter.

[0009] The present invention provides an artificial satellite and a method for controlling the heat radiation heat transfer of the artificial satellite.

[0010] In a first aspect of the present invention, the present disclosure provides a satellite body including at least one radiation surface configured to radiate heat of the satellite into outer space, and an attitude determination and control (ADCS) system for controlling the orientation of the satellite orbiting around the Earth, wherein the ADCS system is configured to change the orientation of the satellite with respect to the sun between a first position where the radiation surface is away from the sun and a second position where the radiation surface is exposed to the sun.

[0011] In some embodiments, the ADCS system is configured to change the orientation of the satellite by rotating the satellite around a rotation axis extending substantially perpendicular to the direction of travel of the satellite.

[0012] In some embodiments, a surface in the normal direction to the radiation surface extends away from the satellite body in a first direction substantially perpendicular to the rotation axis.

[0013] In some embodiments, in the second position, the angle between the first direction and the solar radiation incident on the radiation surface is 90° or less, optionally 60° or less, optionally 30° or less, or optionally 10° or less.

[0014] In some embodiments, the satellite further includes one or more temperature sensors for determining the temperature of the satellite and / or the temperature of one or more components of the satellite, and the ADCS system is configured to change the orientation of the satellite in response to signals from the one or more temperature sensors.

[0015] In some embodiments, the ADCS system is configured to maintain the satellite in the second position in order to absorb incident heat from solar radiation on the radiation surface and control the temperature of the satellite.

[0016] In some embodiments, the ADCS system is configured to maintain the satellite in the second position for at least 5 minutes, optionally at least 10 minutes, or optionally at least 15 minutes.

[0017] In some embodiments, the ADCS system is configured to maintain the satellite at a second position until a temperature setpoint is reached.

[0018] In some embodiments, the ADCS system is configured to be able to maintain the satellite at a first position in order to control the temperature of the satellite by releasing the heat of the satellite into space through a radiation surface.

[0019] In some embodiments, the solar absorptance of the radiation surface is between 10% and 60%, optionally between 20% and 50%, or optionally between 30% and 40%.

[0020] In some embodiments, the radiation surface has an infrared (IR) emissivity of at least 0.85, optionally at least 0.9, and a solar absorptance of at least 0.3, or optionally at least 0.35.

[0021] In some embodiments, the radiation surface is at least 0.05 m 2 , optionally at least 0.1 m 2 , or optionally at least 0.15 m 2 in area.

[0022] In some embodiments, the radiation surface includes a part of a radiation panel that forms a part of an existing structure of the satellite body.

[0023] In some embodiments, the radiation surface includes an adhesive tape made of polyvinyl fluoride film.

[0024] In some embodiments, the satellite further includes one or more solar panels attached to the satellite body, and the plane perpendicular to the sun of the solar panel extends away from the satellite body in a direction substantially opposite to the first direction.

[0025] In some embodiments, the ADCS system is configured to be able to alternately control the orientation of the satellite between a first position and a second position in order to selectively enable heating and cooling of the satellite.

[0026] In some embodiments, the satellite further comprises a SAR antenna attached to the satellite body, and the SAR antenna faces the Earth at the first and second positions.

[0027] In some embodiments, the satellite is a nanosatellite or a microsatellite.

[0028] In some embodiments, the satellite is an Earth observation satellite and optionally a SAR satellite.

[0029] In a second aspect of the present invention, a method for controlling the heat radiation heat transfer of a satellite, wherein the satellite comprises a satellite body including at least one radiation surface configured to emit the heat of the satellite into outer space, and an attitude determination and control (ADCS) system for controlling the orientation of the satellite orbiting around the Earth, which is optionally a satellite according to the first aspect, and the method executed by the ADCS system includes changing the orientation of the satellite with respect to the sun between a first position where the radiation surface faces away from the sun and a second position where the radiation surface is exposed to the sun.

[0030] In some embodiments, changing the orientation of the satellite includes rotating the satellite around a rotation axis that extends substantially perpendicular to the direction of travel of the satellite.

[0031] In some embodiments, the method further includes maintaining the satellite at the second position in order to absorb the incident heat from the solar radiation on the radiation surface and control the temperature of the satellite.

[0032] In some embodiments, the method further includes maintaining the satellite at the second position for at least 5 minutes, optionally at least 10 minutes, or optionally at least 15 minutes.

[0033] In some embodiments, the method further includes maintaining the satellite at the second position until a temperature set point is reached.

[0034] In some embodiments, the method further includes maintaining the satellite at a first position to control the satellite temperature by radiating the satellite's heat into space through a radiating surface.

[0035] In some embodiments, the method further includes alternately changing the orientation of the satellite between a first position and a second position to selectively heat and cool the satellite.

[0036] In some embodiments, the method replaces redundant lines of the thermal control system.

[0037] In some embodiments, the method is performed during an imaging mission.

[0038] In some embodiments, the method is performed according to a predetermined position, a predetermined time, and / or a measured temperature in orbit.

[0039] Also, in embodiments of the present invention, when implemented in a computing system that forms part of a satellite operating system, there is provided a computer-readable medium containing instructions in the form of an algorithm to cause the system to execute any of the methods described herein.

[0040] Preferred features may be appropriately combined as would be apparent to those skilled in the art and may be combined with any aspect of the present invention. A satellite according to a first aspect can be developed to have the features described in a method for controlling radiative heat transfer according to a second aspect. The method according to the second aspect can be further developed using the features described in the context of the satellite according to the first aspect.

[0041] Embodiments of the present invention are described by way of example with reference to the following drawings.

Brief Description of the Drawings

[0042]

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[0043] Throughout the drawings, like features are represented using common reference numerals.

DETAILED DESCRIPTION OF THE INVENTION

[0044] Hereinafter, embodiments of the present invention will be described as an example. These examples represent the best methods currently known to the applicant for practicing the present invention, but are not the only methods by which this can be achieved. The description shows the functions of the examples and a series of procedures for constructing and operating the examples. However, the same or equivalent functions and sequences can be realized by different examples.

[0045] Satellite 100 moves along the traveling direction 101. The satellite 100 shown in FIG. 1 is a nanosatellite for earth observation. However, the present disclosure is equally applicable to other spacecraft such as small satellites and large spaceships. In the following description, for simplicity, reference is made to artificial satellites.

[0046] The satellites of some embodiments of the present invention are orbiting within the low Earth orbit or are configured to orbit within the low Earth orbit. In some embodiments, satellite 100 is controllable by a ground segment (not shown in FIG. 1) composed of all ground-based elements for a spacecraft system. In some embodiments, there are ground stations and / or computing systems configured for the purpose of controlling satellites moving in orbits around the Earth and / or for the purpose of performing some of the operations described in the present disclosure.

[0047] As shown in FIG. 1, the satellite 100 is composed of a satellite body 110, one or more solar panels 150, and wings 160. One or more SAR antennas can be attached to the wings 160 of the satellite. The satellite 100 further includes a propulsion system 190 attached as shown to the surface of the satellite body 110 on the side opposite to the solar panel 150. The propulsion system 190 can be composed of a plurality of thrusters (e.g., four), one of which is illustrated as 194 in FIG. 1 and is part of a system for operating and maneuvering the satellite to properly position the satellite for taking SAR images of the Earth. The satellite body may house a computing system configured to perform some of the operations described herein.

[0048] The satellite 100 further includes an Attitude Determination and Control System (ADCS) unit. The ADCS unit (not shown) is typically disposed within the satellite body 110. The ADCS unit is used to control the satellite orbiting in space, i.e., to direct the satellite 100 in a desired direction and maintain it. The ADCS unit will be further described with reference to FIG. 12. The ADCS unit may include one or more reaction wheels (not shown). The reaction wheels may exert a torque force on the satellite body 110. Additionally or alternatively, the ADCS unit may be composed of a plurality of torque rods. These plurality of torque rods are typically operated to maintain the satellite 100 in a specific attitude, and the operation is controlled by an ADCS controller.

[0049] In some embodiments, satellite 100 includes one or more optical devices (such as star tracker 170 and / or sun sensors) for measuring the position of the star using a photovoltaic cell or a camera. The star tracker 170 may be attached to the upper surface of the satellite body 110. The star tracker 170 may be used to determine the orientation of the satellite with respect to the sun. The star tracker generally takes pictures of space and compares the images with an internally stored catalog to determine the azimuth with respect to the sun. In the embodiment shown in FIG. 1, two star trackers 170 are used, and even if one of them is blocked by sunlight, the other one can still be visible. The satellite 100 may also include one or more sun sensors (not shown) as navigation devices for detecting the position of the sun. Based on one or more star trackers and / or one or more sun sensors, the satellite can determine the orientation of the satellite with respect to the sun. This information is provided to the ADCS system for control so that the orientation of the satellite 100 can be changed between the first position and the second position, as will be further described below.

[0050] As shown in FIG. 1, the artificial satellite 100 is composed of a radiation surface 140. The radiation surface 140 may be a dedicated surface for dissipating excess heat by heat radiation heat transfer, and may also be called a radiator. The radiation surface 140 is configured to release the heat of the satellite into space. This is usually done by emitting infrared (IR) rays from the surface. The radiation surface 140 shown in FIG. 1 is mainly circular. However, in an alternative example, the radiation surface 140 may have a different shape, such as mainly rectangular. The radiation surface has an area of at least 0.05 m 2 or at least 0.1 m 2 or at least 0.15 m 2 The circular radiation surface 140 shown in FIG. 1 has an area of about 0.2 m 2 and a diameter of about 0.5 m. Depending on the internal temperature and the heat dissipation amount, the radiation surface radiates heat between 280 W / m 2 ~480 W / m 2

[0051] ​The present invention is based on the discovery that the radiator or radiating surface of a satellite can also be used to absorb incident heat from solar radiation. By adjusting the orientation of the satellite with respect to the sun, the radiator or radiating surface can be controlled to perform heat absorption or heat radiation, respectively. To perform active thermal control, the ADCS system of the satellite can be used to change the orientation of the satellite between a first position where the radiating surface faces away from the sun and a second position where the radiating surface is exposed to the sun. In this case, since the ADCS system is integrated into the satellite and no additional elements are required to control the thermal behavior, the cost, integration time, and number of components of the satellite can be reduced. This is particularly advantageous for small spacecraft with strict mass and volume constraints. For example, if the power required to control the thermal behavior using the ADCS system is reduced, it may be possible to downsize the solar panels in the satellite design. Additionally, or alternatively, the satellite may be equipped with one electric heating system instead of two redundant electric heating systems. When changing the orientation of the satellite to utilize solar heat, the power consumption for this purpose is less than the power required for an electric stove.

[0052] In some embodiments, the radiating surface 140 and the solar panel 150 are arranged on opposite sides of the satellite body 110. The surface perpendicular to the radiating surface 140 may extend away from the satellite body in the first direction Y. The surface perpendicular to the solar panel 150 may extend away from the satellite body in a direction substantially opposite to the first direction Y. This direction is shown as -Y in FIG. 1.

[0053] During normal operation of the satellite, the solar panels can always be oriented towards the sun as much as possible. Therefore, the position of satellite 100 can be controlled such that the solar panel 150 faces the sun, as shown in FIG. 1. The direction of the sun is indicated by the arrow "S" in FIG. 1, which is also called the solar vector. It is well known in the art that more energy can be captured if the solar panel is facing the sun directly. However, there are exceptions when the satellite is oriented in a specific direction for the satellite to exert its original capabilities, such as imaging the earth by synthetic aperture radar (SAR) or other imaging methods. For imaging, it is necessary to align the length direction of the SAR antenna along the moving direction 101. However, during normal operation of satellite 100 shown in FIG. 1, the radiation surface 140 is continuously facing away from the sun. Therefore, satellite 100 is in a position where the radiation surface 140 faces deep space. The surrounding space environment functions as a heat sink, and the radiation surface 140 may be used to remove the waste heat generated in the satellite.

[0054] In the first position of the satellite, the radiation surface faces the direction opposite to the sun. In other words, in the first position, the normal vector of the radiation surface extending in the first direction Y points away from the sun. Satellite 100 shown in FIG. 1 is in the first position. The satellite may be in the first position during normal operation. The first position indicates a plurality of positions where the radiation surface is not exposed to the sun. The first position is different from the second position of the satellite where the radiation surface is exposed to the sun. In other words, in the second position, the normal vector of the radiation surface points towards the sun. In the second position, the outer surface of the radiation surface faces the sun. The second position indicates a plurality of positions where the radiation surface faces the sun. By changing the orientation of the satellite to the second position and exposing the radiation surface to the sun, the radiation surface can be utilized as a heat source. By changing the orientation of the satellite with respect to the sun, the heat of the satellite can be actively controlled via the radiation surface 140.

[0055] Hereinafter, the heat radiation heat transfer control of a satellite will be described using a series of still images shown in FIGS. 1 to 11. In the series of still images, the satellite is moving at a speed v along the traveling direction 101. Further, the ADCS subsystem is configured to be able to change the orientation of the satellite with respect to the sun. In some embodiments, the ADCS system is configured to rotate the satellite along the rotation axis Z and change and / or maintain the orientation of the satellite with respect to the sun. The rotation axis Z may penetrate the satellite body 110. The rotation axis Z may extend in a direction substantially perpendicular to the first direction Y. Due to their small size and high agility, it is easy for nanosatellites and small satellites to rotate the entire satellite to change the orientation with respect to the sun.

[0056] In this example, the payload mounted on the main wing 160 is a synthetic SAR antenna, and since it is quite sensitive to temperature, especially high temperature, even when the satellite is not imaging, the direction around the Z axis is selected for rotation. Therefore, the SAR antenna receives infrared heating from the earth even when not imaging. For this reason, it is optimal to keep the SAR antenna generally downward so as not to be directly exposed to the cold of deep space or the extreme heat of direct sunlight. In other examples of different types of satellites and payloads, this may not be a constraint. In that case, it may be selected to rotate the satellite about an axis other than the Z axis to change and / or maintain the orientation of the satellite with respect to the sun.

[0057] As shown in FIGS. 1 to 5, the satellite 100 is oriented at a first position where the radiation surface 140 is not exposed to the sun S. The radiation surface 140 at the first position can radiate the heat of the satellite into outer space. First, as indicated by the arrows in FIGS. 1 to 3, the ADCS system rotates the satellite 100 counterclockwise about the rotation axis Z so that the solar panel 150 continuously faces the sun. In FIG. 4, the rotation around the rotation axis Z of the satellite is reversed. In other words, the satellite is no longer tracking the sun to maximize the amount of solar power generation through the solar panel 150. FIG. 4 shows the start of a so-called "sun kick maneuver", and the change of position from the first position to the second position is started.

[0058] The ADCS system controls the satellite to rotate in the clockwise direction until the radiation surface 140 reaches the second position in FIG. 6 where it is (at least partially) exposed to the sun S. The angle between the first direction Y and the solar radiation incident on the radiation surface 140 is approximately 90°. As shown in FIGS. 6 to 9, at the second position, the radiation surface absorbs the incident heat from the solar radiation. The amount of heat that can be absorbed is determined by the time the satellite is in the second position, the angle between the incident solar radiation and the radiation surface, the size of the radiation surface, and its absorption rate characteristics.

[0059] As shown in FIGS. 6 to 8, the orientation of the satellite is maintained at the second position. The orientation of the satellite 100 with respect to the sun changes as the satellite moves along the traveling direction 101. In FIG. 7, the angle between the first direction Y perpendicular to the radiation surface 140 and the solar radiation incident on the radiation surface is less than 60°. In FIG. 8, the angle between the first direction Y and the solar radiation incident on the radiation surface is less than 30°. In FIG. 9, the satellite 100 is still in the second position, but is controlled to rotate in the counterclockwise direction to return the satellite from the second position to the first position. According to one example, the satellite is maintained in the second position for at least 5 minutes, optionally at least 10 minutes, or optionally at least 15 minutes, raising the temperature of the satellite by several degrees Celsius. Whether the satellite can control its temperature to be constant strongly depends on the time (duration) the satellite is maintained in the second position. By adjusting the time the satellite is in the second position, the absorbed heat can be controlled with an accuracy of up to 1 watt. Instead of using a fixed time, it is also possible to control based on a temperature set point. In one embodiment, the temperature set point is determined by the upper limit temperature of the battery or other components of the satellite. For example, when the temperature of the battery reaches the upper limit temperature, e.g., 18°C, it is possible to prevent the satellite from being maintained in the second position. In some embodiments, the satellite can be controlled to stay in the second position until a predetermined satellite temperature is reached.

[0060] By maintaining the satellite in the second position, the incident heat of solar radiation received through the radiation surface can be absorbed, and the temperature of the satellite can be controlled. This can be referred to as the heating stage. The amount of electric power generated by solar photovoltaics through the radiation surface can be controlled based on a predetermined duration, required power input, or desired temperature. Therefore, it is also possible to implement a heating function that is actively controlled by the ADCS system. Similarly, it is possible to implement a heating function that does not directly heat the satellite using power. The opposite side of the radiation surface 140 can also be used to radiate the received sunlight input toward the internal elements and heat them. The radiation surface 140 can also transfer heat to the internal elements through the connection members and heat them. As an example, the radiation surface 140 may be part of a panel made of aluminum with good thermal conductivity.

[0061] In FIGS. 10 and 11, the satellite 100 is facing the first position again, and in this position, the radiation surface 140 is used to release heat. This can be referred to as the cooling stage. By maintaining the satellite in the first position, the heat of the satellite can be released into space through the radiation surface, and the temperature of the satellite can be controlled. Therefore, a cooling function that is actively controlled by the ADCS system can be implemented. Since the solar panels need to face the sun as much as possible to supply power to the spacecraft and charge the batteries, the cooling stage is generally longer.

[0062] Based on the above heating stage and cooling stage, by adjusting the duration of both stages, a thermal control method can be designed to maintain the average temperature inside the satellite body within an acceptable temperature range or a preferred temperature range. This method depends on the size of the radiation surface required to cool the satellite in the worst-case scenario and the time the satellite is in the second position to keep the satellite above the minimum acceptable temperature. By alternately changing the orientation of the satellite between the first position and the second position, the heating and cooling of the satellite can be selectively performed as needed.

[0063] The cooling and heating functions depend on the optical properties of the radiation surface, i.e., the solar absorptance and the infrared emissivity. The solar absorptance determines how much of the incident heat from solar radiation is absorbed by the spacecraft, and the infrared emissivity determines how much heat the spacecraft emits into space and how much heat it receives from the Earth. The surface properties of the radiation surface can be changed according to specific requirements of the satellite, for example, by adding an adhesive tape. Other methods such as coating the radiation surface are also possible.

[0064] The solar absorptance of the radiation surface may be arbitrarily between 10% and 60%, between 20% and 50%, or between 30% and 40%. Within this range, the incident heat absorbed from solar radiation through the radiation surface is sufficient to control the temperature of the satellite and is suppressed to such an extent that it does not exceed the allowable temperature range or the preferred temperature range of the components in a short time.

[0065] The tape can be added in the latter half of the assembly process, is easy to apply, and is relatively inexpensive. In some embodiments, the surface properties of the radiation surface can be changed by adding an adhesive tape to increase the solar heat absorptance at the radiation surface 140 and also increase the heat emission rate of the spacecraft. In some embodiments, an adhesive tape made of polyvinyl fluoride (PVF) film is attached to the radiation surface. In some embodiments, Tedlar® is used for the PVF film. An example of a tape based on Tedlar® PVF film as a substrate is 3M™ Weatherable Film Tape 838.

[0066] Tedlar (registered trademark) PVF film has properties that can meet strict toxicity and flammability requirements, and it was first used in aircraft interiors. It has also been used early on as a sound barrier for highways. Tedlar (registered trademark) PVF film is excellent in chemical resistance and weather resistance. This material was not specially developed for space use, nor was it developed with specific absorption rates or emissivities in mind. The data sheet of 3M (registered trademark) Weatherable Film Tape 838 does not describe the technical specifications regarding absorption rates and emissivities. Similarly, the general properties are not described in the data sheet of Tedlar (registered trademark) PVF film. In some examples, for instance, the radiation surface made of Tedlar (registered trademark) PVF tape has a high infrared emissivity compared to bare aluminum and / or a high solar absorptance compared to bare (polished) aluminum. For this purpose, the radiation surface may have an infrared emissivity ε of at least 0.85, optionally at least 0.9, and / or a solar absorptance a of at least 0.3, optionally at least 0.35.

[0067] Table 1 shows the results of experimentally measuring the absorptance (α) and emissivity (ε) of two samples shown in Figure 13 at a Swiss research institute. The emissivity and absorptance were measured in a specialized laboratory using a Perkin Elmer Lambda 900 spectrophotometer equipped with a Spectralon integrating sphere. In the solar wavelength range, the total reflectance was measured between 300 nm and 2500 nm. Also, in the infrared wavelength range, the total reflectance was measured between 2.5 μm and 20 μm. The value of the solar reflectance a (absorbance or absorptance) was calculated from the measured values using the solar spectrum with an air mass coefficient of 0 outside the atmosphere according to ASTM E 490 and following the solar intensity spectrum of ISO 9845. The value of the emissivity £ (emittance) was calculated for blackbody radiation at room temperature. The reflectance values for wavelengths longer than 20 microns were extrapolated using a constant value.

[0068] In Sample 1, on Side 1, there are two layers of Tedlar (registered trademark) PVF tape (Fig. 13a), which faces the outside of the spacecraft, i.e., the radiation surface. Side 2 (Fig. 13b) is exposed polished aluminum with two layers of polyimide film pasted on it. Side 2 faces the inside of the spacecraft. In Sample 2, on Side 1, there is one layer of Tedlar (registered trademark) PVF tape (Fig. 13c), and Side 2 is polished exposed aluminum (Fig. 13d). As can be seen from the data, the absorption rates when using two layers (α = 0.383) and one layer (α = 0.377) of PVF film are higher than that of polished exposed aluminum. At the same time, the emissivity of the surface is also higher than that of exposed aluminum. In fact, the emissivity when pasting PVF tape is extremely high. Table 1: Absorption rate (α) and emissivity (ε) of exemplary materials TIFF2025523633000002.tif57134

[0069] In some embodiments, the radiation surface 140 may be in the form of a structural panel that forms part of the existing structure of the satellite body 110, as shown in Fig. 14. This configuration corresponds to the embodiments shown in Figs. 1 to 11. In this embodiment, the radiation surface 140 is substantially circular, and a plurality of mounting holes are provided on its outer periphery. When the radiation surface forms part of an existing structure, it can be called a radiation panel. The radiation panel may be integrated with one side of the satellite. As an example, the radiation panel can also be part of the structure attached to the release mechanism of the rocket when launching the satellite into space. The radiation panel may be made of aluminum with a relatively high conductivity. Therefore, the temperature is (relatively) uniform on the panel, and heat is conducted from one or more components that require cooling at the first position to the heat dissipation panel, and from the heat dissipation panel to one or more components that require heating at the second position. This is particularly advantageous for small spacecraft with strict mass constraints. Heat moves between the structures by emission and absorption on the back side of the radiation surface and heat conduction through the structural members connecting the radiation surface and other parts of the spacecraft.

[0070] FIG. 12 is a partial perspective view of a satellite such as satellite 100 of FIG. 1. The ADCS 302 is generally disposed on a satellite body 310 such as the satellite body 110 of FIG. 1 and is used to control the orientation of the satellite. The ADCS unit 302 is composed of a set of three reaction wheels 370a, 370b, 370c disposed within the satellite body 310. The reaction wheels are also called momentum wheels. The reaction wheels 370a, 370b, 370c are controlled by an ADCS controller 341. The reaction wheels 370a, 370b, 370c function by rotating the wheels within the satellite body 310 by an electric motor. Due to the conservation of angular momentum, simply rotating the wheel in one direction causes the satellite to rotate in the opposite direction. Using reaction wheels is a well-known method for determining the orientation of spacecraft such as satellites. The satellite in this example has three reaction wheels 370a, 370b, 370c, each of which serves to orient the satellite on each axis. The reaction wheels 370a, 370b, 370c are shown as having orthogonal axes. In another example, especially for satellites with a large moment of inertia, four or more reaction wheels may be used to more appropriately control various aspects of satellite dynamics, such as slew rate and fine positioning control.

[0071] The ADCS unit shown in FIG. 12 further includes torque rods 305a, 305b, 305c. The torque rods are also used for satellite attitude control. The torque rods 305a, 305b, 305c are typically operated to maintain the satellite 300 in a specific attitude, and that operation is controlled by the ADCS controller 341 described below.

[0072] The ADCS unit 302 further includes an ADCS controller 341. The ADCS controller 341 communicates with an on-board computing system 340. The on-board computing system 340 is composed of a processor 349, a memory 348, and a telemetry unit 345. The memory 348 can be used to store the allowable temperature range or preferred temperature range of the satellite's components and / or temperature set values. The ADCS controller 341 is further configured to receive information from one or more sensors 347. The one or more sensors 347 are configured to measure various quantitative data during the flight of the satellite, such as solar sensors, star trackers 170, temperature sensors, and / or magnetometers for measuring local magnetic fields. The ADCS controller 341 further communicates with a GPS receiver module including a GPS receiver 352 and a GPS antenna 353. The ADCS unit 302 is configured to change the orientation of the satellite with respect to the sun between a first position where the radiation surface is away from the sun and a second position where the radiation surface is exposed to the sun in order to control the satellite's heat radiation heat transfer.

[0073] Nanosatellites and microsatellites are small and sensitive, so the entire satellite can be maneuvered and oriented. This type of operation may be performed using an ADCS unit. In one example, the satellite 100 may be a nanosatellite with a mass of 100 kg. Conventional satellites with a mass of about 1000 kg are generally more expensive and less sensitive than nanosatellites. Satellites can be classified according to their mass. For example, satellites with a mass of about 1 kg to about 10 kg are classified as CubeSats, satellites with a mass of about 50 kg to about 250 kg are classified as microsatellites, satellites with a mass of about 500 kg are classified as small satellites, and satellites with a mass of about 800 kg to about 1200 kg can be classified as conventional satellites.

[0074] In large satellites, it may be possible to incorporate active thermal control such as electric resistance heaters, cryocoolers, thermoelectric coolers, and fluid loops. However, in small or microsatellites, it may not be possible to implement such active thermal control (without increasing weight or volume).

[0075] In some embodiments, the SAR antenna faces the Earth at the first and second positions. In this embodiment, as the satellite rotates about one axis, the orientation from the satellite between the first position and the second position is changed. The SAR antenna has strict temperature requirements, and it is necessary to minimize the thermal cycle. By keeping the SAR antenna in the attitude facing the Earth, the SAR antenna is in the optimal attitude for starting imaging, and temperature changes can be suppressed as much as possible.

[0076] FIG. 15 shows the temperature profile of the radiation surface. The temperature profile is simulated for a heat dissipation panel that forms part of the existing structure of the satellite body 110, and the radiation surface 140 is made of an adhesive tape of polyvinyl fluoride film. Between about 16:30 and 16:45, and between about 19:45 and 20:00, the orientation of the satellite is changed to the second position where the radiation surface is exposed to the sun. During these time periods, in the simulation, the temperature T of the radiation surface rises almost linearly with time. During these time periods, the temperature rises due to the absorption of incident heat from solar radiation received through the radiation surface. These time periods can be called the heating stage. During the heating period, the satellite is oriented to be maintained at the first position. At this position, since the satellite heat is radiated into space through the radiation surface, the satellite temperature drops. These time periods can be called the cooling stage. As can be seen from FIG. 15, the temperature of the radiation surface remains between 8° C. and 18° C. for about 5 hours. Therefore, by alternately controlling the orientation of the satellite between the first position and the second position, the radiation surface selectively heats and cools the satellite. This enables active thermal control of the satellite controlled by the ADCS system.

[0077] Figure 16 shows the multiple temperature profiles of various components mounted on the satellite. Similar to Figure 15, the temperature control includes two heating stages, each lasting approximately 15 minutes. As can be seen in Figure 16, during a simulation period of about 5 hours, the temperature of each component remains between 0°C and 20°C. Therefore, the power generated from sunlight received through the radiation surface is sufficient to meet the temperature requirements of multiple devices mounted on the satellite. For example, the battery is a component with relatively narrow temperature limits as its operating range is generally narrow, from 0°C to 20°C. The method of controlling heat radiation heat transfer of the present invention can be used to maintain the temperature range of on-orbit electronic devices and batteries. By adjusting the incident angle between the radiation surface and the sun throughout the satellite's orbit, heat exchange can be proactively controlled and the components can be thermally stabilized. By adjusting the time when the satellite is in the second position, the absorbed heat can be controlled with an accuracy of up to 1 watt.

[0078] The method of controlling heat radiation heat transfer centered on the ADCS unit can provide redundancy to the standard heat control system. Therefore, this method can replace the redundant lines of the heat control system. The standard heat control system usually consists of a redundant system with two electric heaters, two thermostats, and two cables respectively. One is the main line and the other is the redundant line. By replacing the redundant line with a heat radiation heat transfer control method based on the ADCS unit, the number of components incorporated into the satellite can be reduced and the cost can be cut. Therefore, the satellite can be designed with one electric heating system instead of two redundant electric heating systems, and the power consumption can also be reduced.

[0079] In the case of an imaging satellite, the method of controlling the heat radiation heat transfer of the satellite is generally executed when the satellite is not imaging. However, it can also be used to complement imaging. The method of controlling the heat radiation heat transfer can be executed between imaging tasks, for example, to equalize the temperature conditions at the start of each imaging task. Furthermore, this method can also be executed during the propulsion stage to cool the satellite. During the propulsion stage, while the attitude of the spacecraft is fixed, within the spacecraft, electric power input is generated by the propulsion force. This electric power input may include sunlight incident on the radiation surface (when the satellite is in the second position), which leads to an undesirable increase in the temperature of the satellite. Between two propulsion stages, the radiation surface is directed towards space (the first position) for a predetermined time to discharge the heat generated previously. This eliminates the concern about temperature, so the duration of the propulsion stage can be shortened. The method of controlling the heat radiation heat transfer of the satellite can also be used for other spacecraft such as communication satellites.

[0080] Additionally or alternatively, this method can be repeated at a predetermined position and time in orbit and / or in response to the measured temperature. For example, the satellite can be set to change between the first position and the second position while orbiting a certain part of the orbit. The certain part of the orbit may be the time zone when the satellite passes over the earth's sky where it does not need to image. Also, the certain part of the orbit can be set to change from the first position to the second position only when the satellite is not in the earth's shadow from the sun. This method can also be executed based on the measured temperature and one or more temperature set values of the components or the radiation surface. Similar to the embodiments shown in FIGS. 15 and 16, the satellite can be controlled to stay in the first position until the temperature drops below, for example, 8 °C and / or to maintain in the second position until the temperature exceeds, for example, 18 °C. Therefore, the satellite is maintained in the second position until it reaches the (predetermined) temperature set point.

[0081] The above embodiments are fully automatic. In some examples, the user or operator of the system may manually execute some steps of the method to be executed.

[0082] In the embodiments described in the present invention, the system may be implemented as any form of computing and / or electronic device. Such a device may include one or more processors that are microprocessors, controllers, or other suitable types of processors that process computer-executable instructions for controlling the operation of the device to collect and record routing information. In some examples, such as when a system-on-chip architecture is used, the processor may include one or more fixed-function blocks (also referred to as accelerators) that implement part of the method in hardware (rather than software or firmware). Platform software including an operating system or other suitable platform software may be provided to a computing-based device to configure it to execute application software on the device.

[0083] The various functions described herein may be implemented in hardware, software, or any combination thereof. When implemented in software, the functions may be stored or transmitted as one or more instructions or codes on a computer-readable medium. The computer-readable medium may include, for example, a computer-readable storage medium. The computer-readable storage medium may include volatile or non-volatile, removable or non-removable media implemented in any method or technology for storing information such as computer-readable instructions, data structures, program modules, or other data. Any commercially available storage medium accessible by a computer can be used as the computer-readable storage medium. By way of non-limiting example, such computer-readable storage medium includes RAM, ROM, EEPROM, flash memory or other memory devices, CD-ROM or other optical disk storage, magnetic disk storage or other magnetic storage devices, or other media accessible by a computer that can be used to transfer or store the program code of interest in the form of instructions or data structures.

[0084] Although a local device is illustrated, it should be understood that the computing device may be located remotely and accessed via a network or other communication link (e.g., using a communication interface).

[0085] It is understood that the above advantages and benefits may relate to one embodiment or multiple embodiments. Multiple embodiments are not limited to those that solve any or all of the problems mentioned or have the advantages and benefits mentioned. Variations should be considered to be within the scope of the present invention.

[0086] References to "one" item refer to one or more of that item. As used herein, the term "comprising" means including the steps or elements of a particular method, but such steps or elements do not constitute an exclusive list and the method or apparatus may include additional steps or elements.

[0087] As used herein, the terms "component" and "system" are intended to include a computer-readable data store composed of computer-executable instructions that, when executed by a processor, perform a specific function. The computer-executable instructions may include routines, functions, etc. It should also be understood that a component or system may be localized on a single device or distributed across multiple devices.

[0088] Furthermore, as used herein, the term "exemplary" is intended to mean "an example or illustration of something".

[0089] Furthermore, to the extent that the term "comprising" is used within the detailed description or claims, since the term "comprising" is construed as a transitional term within the claims, this term is intended to have the same inclusiveness as the term "including".

[0090] The accompanying drawings illustrate exemplary ways. The ways are shown and described as a series of operations performed in a particular order, but it should be understood and appreciated that the ways are not limited by the order. For example, some operations can occur in an order different from that described herein. Further, an act may occur simultaneously with another act. Additionally, in some cases, not all acts may be necessary to perform the ways described herein.

[0091] Furthermore, the acts described herein may be implemented by one or more processors and / or may include computer-executable instructions stored on a computer-readable medium. The computer-executable instructions may include routines, subroutines, programs, execution threads, and the like. Additionally, the results of the operations of the ways can be stored on a computer-readable medium or displayed on a display device.

[0092] The order of the steps of the ways described herein is exemplary, but these steps may be performed in any suitable order or, where appropriate, simultaneously. Further, steps may be added or substituted in any of the ways, or individual steps may be deleted, without departing from the scope of the subject matter described herein. Combinations of aspects of any of the above embodiments with aspects of any of the other above embodiments may further form embodiments without losing the desired effects.

[0093] The description of the above preferred embodiments is shown by way of example only, and it should be understood that those skilled in the art can make various modifications. The above content includes an example of one or more embodiments. Of course, for the purpose of explaining the foregoing aspects, it is impossible to explain all possible modifications and changes to the above device or method, but those skilled in the art can recognize that many more modifications and combinations of various aspects are possible. Therefore, the described aspects are intended to include all such modifications, revisions, and amendments that fall within the scope of the appended claims.

Claims

1. Synthetic Aperture Radar (SAR) satellite, A satellite body including at least one radiating surface configured to release heat from the satellite into space, The SAR antenna attached to the satellite body, The system comprises an attitude determination and control (ADCS) system for controlling the orientation of the satellite orbiting the Earth, wherein the ADCS system is configured to change the orientation of the satellite relative to the Sun between a first position where the radiating surface is away from the Sun and a second position where the radiating surface is exposed to the Sun, and the ADCS system Rotating the satellite around a rotation axis (Z) to change the orientation of the satellite, wherein the normal vector to the radiating plane extends away from the satellite body in a first direction (Y) substantially perpendicular to the rotation axis (Z), To control the satellite temperature by absorbing incident heat from solar radiation on the radiating surface, the satellite is maintained in the second position, A satellite configured to perform the following actions.

2. The satellite according to claim 1, wherein the axis of rotation (Z) extends substantially perpendicular to the direction of travel of the satellite, and at the second position, the angle between the first direction (Y) and the solar radiation incident on the radiating surface is 90° or less, and is arbitrarily less than 60°, arbitrarily less than 30°, or arbitrarily less than 10°.

3. The satellite according to claim 1 or 2, further comprising one or more temperature sensors for determining the temperature of the satellite and / or the temperature of one or more components of the satellite, wherein the ADCS system is configured to change the orientation of the satellite in response to signals from the one or more temperature sensors.

4. The satellite according to claim 1 or 2, wherein the ADCS system is configured to maintain the satellite in the second position for at least 5 minutes, optionally at least 10 minutes, or optionally at least 15 minutes.

5. The satellite according to claim 1 or 2, wherein the ADCS system is configured to maintain the satellite in the second position until a temperature setpoint is reached.

6. The satellite according to claim 1 or 2, wherein the ADCS system is configured to maintain the satellite in a first position in order to control the temperature of the satellite by releasing the heat of the satellite into space through the radiating surface.

7. The satellite according to claim 1 or 2, wherein the solar absorption rate of the radiating surface is between 10% and 60%, arbitrarily between 20% and 50%, or arbitrarily between 30% and 40%.

8. The satellite according to claim 1 or 2, wherein the radiating surface has an infrared "IR" emissivity of at least 0.85, optionally at least 0.9, and a solar absorptivity of at least 0.3, optionally at least 0.

35.

9. The radiating surface is at least 0.05 m 2 , optionally at least 0.1 m 2 , or optionally at least 0.15 m 2 The satellite according to claim 1 or 2, having an area of ​​the specified size.

10. The satellite according to claim 1 or 2, wherein the radiating surface includes a portion of a radiating panel that forms part of the existing structure of the satellite body.

11. The satellite according to claim 1 or 2, wherein the radiating surface is provided with an adhesive tape made of polyvinyl fluoride film.

12. The satellite according to claim 1 or 2, further comprising one or more solar panels attached to the satellite body, wherein the sun-perpendicular surface of the solar panel extends away from the satellite body in a direction substantially opposite to the first direction (Y).

13. The satellite according to claim 1 or 2, wherein the ADCS system is configured to alternately control the orientation of the satellite between the first position and the second position in order to selectively heat and cool the satellite.

14. The satellite according to claim 1 or 2, wherein the SAR antenna is oriented toward the Earth at the first and second positions.

15. The satellite according to claim 1 or 2, wherein the satellite is a microsatellite or a small satellite.

16. A method for controlling thermal radiative heat transfer of a synthetic aperture radar (SAR) satellite, which is optionally the satellite described in claim 1 or 2, wherein the satellite comprises a satellite body including at least one radiating surface configured to release heat from the satellite into space, a SAR antenna attached to the satellite body, and an attitude determination and control (ADCS) system for controlling the orientation of the satellite as it orbits the Earth, wherein the method is performed by the ADCS system. This includes changing the orientation of the satellite relative to the sun between a first position where the radiating surface is away from the sun and a second position where the radiating surface is exposed to the sun, A method for changing the orientation of the satellite, comprising rotating the satellite about a rotation axis (Z), wherein the normal vector to the radiating plane extends away from the satellite body in a first direction (Y) substantially perpendicular to the rotation axis (Z).

17. The method according to claim 16, wherein the axis of rotation (Z) extends substantially perpendicular to the direction of motion of the satellite.

18. A method for controlling thermal radiation heat transfer of a synthetic aperture radar (SAR) satellite, which is optionally the satellite described in claim 1 or 2, wherein the satellite comprises a satellite body including at least one radiating surface configured to emit heat from the satellite into space, a SAR antenna attached to the satellite body, and an attitude determination and control (ADCS) system for controlling the orientation of the satellite as it orbits the Earth, wherein the method is performed by the ADCS system. The orientation of the satellite relative to the sun is changed between a first position where the radiating surface is away from the sun and a second position where the radiating surface is exposed to the sun. To control the satellite temperature by absorbing incident heat from solar radiation on the radiating surface, the satellite is maintained in the second position, Equipped with, The method replaces redundant lines in the thermal control system, i.e., the satellite is designed with one electric heating system instead of two redundant electric heating systems.

19. Maintaining the satellite in the second position for at least 5 minutes, optionally at least 10 minutes, or optionally at least 15 minutes, or Maintaining the satellite in the second position until the temperature set point is reached, To control the temperature of the satellite by releasing its heat into space through the radiating surface, the satellite is maintained in the first position, In order to selectively enable heating and cooling of the satellite, the orientation of the satellite is alternated between the first position and the second position, The method according to claim 18, further comprising:

20. The method according to claim 16, The method replaces redundant lines in the thermal control system, i.e., the satellite is designed with one electric heating system instead of two redundant electric heating systems. The above method is performed during the imaging mission. A method wherein the method is performed in response to a predetermined position, a predetermined time, and / or a measured temperature on the orbit.

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