An extremely low orbit medium-inclination remote sensing satellite and design method

Abstract

The application relates to the fields of aerospace engineering and remote sensing technology, in particular to a very low orbit medium-inclination remote sensing satellite and a design method, the satellite runs on a medium-inclination orbit, adopts a spherical cap windward surface and a smooth side plate transition structure to reduce aerodynamic resistance and homogenize heat flow; a detector and a refrigerator are connected with a main heat dissipation surface through a heat conduction structure to form a cold end dissipation channel; when the satellite is in a launch state, a solar panel is folded on the -Zs side of the satellite body, the satellite is unfolded in the +Ys direction after being put into the orbit, and the normal line of the solar panel and the -Zs direction form an included angle, a SAR antenna is folded on the -Ys side of the satellite body, the satellite is unfolded in the -Ys direction after being put into the orbit, the normal line of the SAR antenna array surface points to the +Zs direction, and a sun-shading and counterweight structure is formed; double Hall electric thrusters are arranged on the positive and negative Xs surfaces, the orbit sun angle is calculated through GNSS and a sun vector, and yaw switching is performed to realize forward and reverse flight switching.

Description

Technical Field

[0001] This invention relates to the fields of aerospace engineering and remote sensing technology, specifically to a very low Earth orbit medium inclination remote sensing satellite and its design method. Background Technology

[0002] Ultra Low Earth Orbit (ULEO, below 250 km altitude) offers advantages such as high imaging resolution, light system weight, and low launch cost, making it suitable for Earth remote sensing missions. It represents the lowest orbit and most extreme performance among Very Low Earth Orbit (VLEO) satellites. Middle Inclined Orbit (MIO, inclination approximately 20°–60°) can cover mid- and low-latitude regions more frequently and across multiple time phases, meeting the flexibility requirements of urban monitoring, agricultural assessment, and disaster emergency response in terms of revisit cycles and imaging time periods.

[0003] However, due to significant aerodynamic drag in ULEO orbits, the satellite's frontal area must be kept extremely small to extend its lifespan, while simultaneously facing a contradiction between solar energy acquisition and effective heat dissipation. Especially in MIO orbits, the large variation in the solar incidence angle reduces the efficiency of traditional horizontal biplane solar panels, as seen in the Qiankun-1 satellite. Figure 6 As shown, the configuration of Japan's SLATS satellite is as follows: Figure 7 As shown. In MIO orbit, on the one hand, the horizontal solar panels are inefficient when the orbital solar angle is large. Therefore, the satellite adopts a rolling maneuver to align with the sun, but this affects the camera's Earth observation. On the other hand, one side of the solar panel will have a large amount of thermal radiation, which disrupts the stability of the heat dissipation direction.

[0004] For the long-life design of ELEAP satellites, the core is to improve the total impulse and reliability of the propulsion system. At present, high specific impulse electric propulsion systems are becoming increasingly mature. Traditional electric propulsion systems all have a single cathode, and most ELEAP satellites have an electric propulsion system installed at the tail for orbital elevation.

[0005] Conventional remote sensing satellites are concentrated in sun-synchronous orbits (SSOs), and most employ three-axis stabilized Earth-oriented or roll-and-slide methods to extend their observation bandwidth. A few attempts to deploy payloads in extremely low Earth orbits (ULOs) or medium-inclination orbits have often failed to systematically address the complex coupling issues of solar panel irradiation interference on the main heat dissipation surface, limited thruster lifespan, and aerodynamic attitude interference. For medium-inclination orbits (MIOs), where the ground solar altitude varies significantly, simply maintaining a fixed attitude makes it difficult to simultaneously address solar irradiation, heat dissipation, and propulsion, thus limiting long-term operation and the engineering implementation of high-performance remote sensing systems.

[0006] In view of this, the present invention provides a very low Earth orbit medium inclination remote sensing satellite and its design method to solve the above problems. Summary of the Invention

[0007] The purpose of this invention is to provide a very low Earth orbit medium inclination remote sensing satellite and its design method, which solves the problems of achieving stable observation, efficient power supply, reliable thermal control and long-term propulsion capability of remote sensing imaging payload components under strong aerodynamic drag and high solar angle variation.

[0008] To achieve the above objectives, the present invention provides the following technical solution:

[0009] In a first aspect, the present invention provides a very low Earth orbit (LEO) medium-inclination remote sensing satellite. The satellite has Xs, Ys, and Zs directions, where the positive Xs direction is the forward direction of the satellite in orbit, the positive Ys direction is the negative normal to the orbital plane, and the positive Zs direction is the Earth observation direction. The satellite includes a satellite body, a remote sensing imaging payload assembly, solar panels, and an attitude propulsion system, wherein:

[0010] The satellite body has a spherical crown surface structure on its windward side. The spherical crown surface has a continuous curvature and a smoothing surface without a broken line transition. Its height H is adjustable according to the rectification requirements of the satellite's forebody.

[0011] The remote sensing imaging payload assembly includes a thermal infrared camera lens, a thermal infrared camera detector, and a cooler arranged along the flight direction. The thermal infrared camera lens and data transmission antenna are arranged in series along the Xs direction and share the Earth observation axis Zs. The thermal infrared camera detector and cooler are arranged near the main heat dissipation surface in the -Ys direction and form a cold end heat dissipation channel through thermal conductive connection.

[0012] The solar panel is installed on the +Ys side, opposite to the main heat dissipation surface, to avoid mutual shading; and together with the SAR antenna installed on the -Ys side with the array normal pointing to the +Zs direction, it forms a sunshade and counterweight system, so that the main heat dissipation surface is always in the shaded state on orbit.

[0013] The attitude propulsion system includes Hall thrusters arranged on the positive and negative Xs planes respectively. The attitude control system performs yaw maneuvers according to the orbital solar angle β. When |β| is greater than the set threshold ε, it switches between forward and backward flight attitudes to ensure that the solar panels are illuminated and the main heat dissipation surface is shaded. The orbit is maintained by the alternating operation of the two thrusters.

[0014] As a preferred technical solution of the first aspect of the present invention, the spherical cap surface structure is formed by the intersection of the circumscribed spherical surface and the side surface of the star, and the spherical surface is cut by the side surface of the star to obtain a geometric body with square sides and a spherical cap top surface;

[0015] The spherical cap curved surface structure is integrally formed with the side plate of the satellite. The side plate is smooth and without sharp edges, which is used to reduce aerodynamic interference at the edges and prevent heat from concentrating on the side of the satellite at the sharp edges. The satellite body is made of materials with high temperature resistance and good heat insulation.

[0016] As a preferred embodiment of the first aspect of the present invention, a heat-conducting connection plate is provided between the thermal infrared camera detector and the refrigerator and the main heat dissipation surface. The heat-conducting connection plate is fixed to the main heat dissipation surface through a thermal coupling component. When the satellite attitude is adjusted, the direction of heat flow at the cold end is consistent with the normal direction of the heat dissipation surface, thereby reducing the temperature drift of the refrigerator and improving thermal stability.

[0017] As a preferred technical solution of the first aspect of the present invention, the solar panel is connected to the satellite body through a hinge assembly. After unfolding, the angle between the plane of the solar panel and the top surface of the satellite is 10° to 30°, which is used to maintain stable incident light conditions when the orbital solar angle changes, and the mechanical limiting structure constrains the sway amplitude of the solar panel during attitude switching to not exceed 3°.

[0018] As a preferred embodiment of the first aspect of the present invention, the SAR antenna surface has a low-absorption reflective layer for reflecting thermal radiation from the sun and the earth; the reflective layer is arranged parallel to the un-illuminated surface of the solar panel, so that the two form a passive shading and thermal radiation balance zone to ensure the thermal stability of the main heat dissipation surface.

[0019] As a preferred technical solution of the first aspect of the present invention, the Hall electric thrusters on the positive and negative Xs planes are respectively equipped with independent cathode modules and propulsion storage tanks. The satellite body performs life assessment based on the cumulative working time and thrust deviation of the two thrusters. When the thrust performance of either thruster drops below a preset threshold, the other thruster is automatically switched to work.

[0020] As a preferred technical solution of the first aspect of the present invention, when the attitude control system performs yaw maneuvers, it uses the orbital solar angle β and the solar panel output power as a joint criterion. When the solar panel output power drops by more than 10% after attitude switching, it performs local roll correction to maintain the optimal value of the solar panel's illumination angle, thereby achieving joint optimization of energy and attitude.

[0021] As a preferred embodiment of the first aspect of this invention, during the ground assembly stage, the solar panel and SAR antenna are folded and fixed to the top surface of the satellite and the −Ys plane, respectively. After orbit insertion, they are sequentially unfolded by a release mechanism, causing the SAR antenna to first rotate 90 degrees in the −Ys direction until the array normal points to the +Zs direction. Subsequently, the solar panel unfolds to a preset angle and is locked. By controlling the unfolding sequence, the attitude disturbance during the structural unfolding process is minimized.

[0022] As a preferred technical solution of the first aspect of the present invention, the main heat dissipation surface is set on the -Ys plane, and its normal direction is opposite to that of the normal direction of the solar panel in the Ys direction. When the orbital solar angle β changes, the incident angle of the heat dissipation surface is automatically adjusted by attitude yaw so that the heat dissipation efficiency is maintained within the design range.

[0023] In a second aspect, the present invention provides a design method for a low Earth orbit medium inclination remote sensing satellite, used for designing the first aspect, comprising the following steps:

[0024] S101: Determine the orbit and coordinate reference; based on the mission coverage and resolution requirements, select a medium-inclination orbit suitable for very low Earth orbit operation and establish a structural coordinate system, wherein: the remote sensing satellite has Xs, Ys and Zs directions, the positive direction of Xs is the forward direction of the remote sensing satellite in orbit, the positive direction of Ys is the negative normal of the orbital plane of the remote sensing satellite in orbit, and the positive direction of Zs is the Earth observation direction of the remote sensing satellite in orbit.

[0025] S102: Determine the aerodynamic shape of the satellite; in the coordinate system, the windward surface of the satellite is a spherical cap surface structure with continuous curvature, and the spherical cap surface structure is smoothly connected to the side plate of the satellite to form a rectifying surface;

[0026] S103: Configure imaging and heat dissipation system; arrange thermal infrared camera lens and data transmission antenna along Xs direction and share the ground axis Zs; place thermal infrared camera detector and cooler on the main heat dissipation surface side close to -Ys; form a dissipation path from cold end to heat dissipation surface through thermal conduction connection to achieve thermal control and structural coordination.

[0027] S104: Design energy and sunshade structure; solar panels are arranged on the +Ys side, and SAR antennas are arranged on the -Ys side offset in the +Zs direction, which have the functions of imaging, sunshade and counterweight, to ensure that the main heat dissipation surface operates in the shadow of the sun in orbit.

[0028] S105: Set the control logic of the attitude propulsion system; install Hall electric thrusters on the positive and negative Xs planes, calculate the orbital solar angle β based on GNSS and solar vector, and perform yaw switching when |β| exceeds the threshold ε to realize positive-to-negative flight conversion, ensuring that the solar panels are illuminated, the main heat dissipation surface is shaded, and maintaining propulsion redundancy;

[0029] S106: Perform structural counterweighting and performance verification; perform lateral counterweighting according to the mass distribution of solar panels and SAR antennas, and comprehensively verify aerodynamics, thermal control and attitude stability to determine the overall satellite coupling balance and complete the design of the ultra-low orbit medium inclination remote sensing satellite.

[0030] Compared with the prior art, the beneficial effects of the present invention are:

[0031] This invention employs a very low orbit with a medium inclination, resulting in two significant benefits. Firstly, the spatial resolution of remote sensing observations increases linearly with decreasing orbital altitude; for example, the spatial resolution of ground features at an orbital altitude of 250 km is twice that of a satellite at an altitude of 500 km using the same remote sensing camera. Secondly, the revisit time resolution of observations in a medium-inclination orbit is greatly improved, and multi-phase day-night observations can be achieved within a single day. Taking the revisit coverage of Beijing by remote sensing satellites as an example, the average revisit coverage of Beijing by a traditional sun-synchronous orbit (SSO) and a 42° medium-inclination orbit is 1 revisit / day and 3 revisits / day, respectively.

[0032] The change in orbit also brings many engineering problems, such as aerodynamic drag and large variations in the orbital solar angle. This invention presents system solutions including solar panels, a main heat dissipation surface, attitude pointing mode, dual-cathode electric propulsion, and a backup electric propulsion layout. Using the satellite design configuration and operating mode of this invention, long-life and reliable operation of the electric propulsion system can be achieved, enabling efficient solar irradiation without the need for satellite attitude roll-to-solar alignment at small orbital solar angles, and providing a stable main heat dissipation surface required for cooling the thermal infrared remote sensing imaging payload components.

[0033] In summary, while adapting to the inclination characteristics of very low Earth orbit, this invention provides a better operating environment and condition. Satellites employing this invention are particularly suitable for thermal infrared remote sensing imaging payloads requiring cryogenic cooling, continuously providing remote sensing data such as ground light, mid-wave thermal infrared, and multi-spectral water color to support the monitoring and assessment of sustainable development goals. Attached Figure Description

[0034] Figure 1 This is a satellite X-axis serial payload arrangement diagram of the present invention;

[0035] Figure 2 This is a schematic diagram of the structure of the spherical crown-shaped windward surface of the present invention;

[0036] Figure 3 This is a schematic diagram of the SAR antenna and solar panel deployment structure of the present invention;

[0037] Figure 4 This is a schematic diagram of the layout of a single unit inside the satellite of this invention.

[0038] Figure 5 This is the logic diagram of attitude yaw maneuvering and sun angle switching in this invention;

[0039] Figure 6 This is a diagram illustrating the configuration of the Qiankun-1 satellite in the background technology.

[0040] Figure 7 This is a diagram illustrating the configuration of the Japanese SLATS satellite in the background technology.

[0041] Figure 8This is a flowchart of the design method for the low Earth orbit inclination remote sensing satellite of the present invention;

[0042] The following are the labels in the diagram: 1. Satellite body; 101. Spherical cap curved surface structure; 2. Thermal infrared camera lens; 3. Hall electric thruster; 4. SAR antenna; 5. Main heat dissipation surface; 6. Solar panel; 7. Data transmission antenna; 8. Propulsion separate storage tank; 9. Thermal infrared camera detector and refrigerator. Detailed Implementation

[0043] The technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of the present invention, and not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0044] In the description of this invention, it should be noted that the terms "vertical," "upper," "lower," "horizontal," etc., indicate the orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings. They are only for the convenience of describing this invention and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation. Therefore, they should not be construed as limitations on this invention.

[0045] In the description of this invention, it should also be noted that, unless otherwise explicitly specified and limited, the terms "set," "install," "connect," and "link" should be interpreted broadly. For example, they can refer to a fixed connection, a detachable connection, or an integral connection; they can refer to a mechanical connection or an electrical connection; they can refer to a direct connection or an indirect connection through an intermediate medium; and they can refer to the internal connection of two components. Those skilled in the art can understand the specific meaning of the above terms in this invention according to the specific circumstances.

[0046] Example 1

[0047] Please see Figure 1This invention provides a technical solution: a low Earth orbit (LEO) medium-inclination remote sensing satellite, addressing the challenges of stable observation, efficient power supply, reliable thermal control, and long-term propulsion capabilities for remote sensing imaging payload components under conditions of strong aerodynamic drag and high solar angle variation. To illustrate the relative positional relationships between the components of the remote sensing satellite, the directions of the satellite during flight are defined. The remote sensing satellite has x, y, and z directions. The positive x-direction is defined by the arrow on the Xs axis in the coordinate system shown in the figure; the positive y-direction is defined by the arrow on the Ys axis; and the positive z-direction is defined by the arrow on the Zs axis. Specifically, the positive Xs direction is the forward direction of the remote sensing satellite during its on-orbit flight; the positive Ys direction is the negative normal to the orbital plane; and the positive Zs direction is the Earth observation direction of the remote sensing satellite during its on-orbit flight.

[0048] It is understandable that, corresponding to the above definitions, the negative x-axis is the direction of the wake of the remote sensing satellite in orbit, the negative y-axis is the positive normal of the orbital plane of the remote sensing satellite in orbit, and the positive z-axis is the zenith direction of the remote sensing satellite in orbit. After defining the x and z axes, the y-axis is defined using a right-handed coordinate system.

[0049] The remote sensing satellite includes the satellite body 1, thermal infrared camera lens 2, thermal infrared camera detector and cooler 9, SAR antenna 4, main heat dissipation surface 5, Hall thrusters on positive and negative Xs planes 3, propulsion tanks 8, data transmission antenna 7, and solar panels 6, among which:

[0050] Satellite body 1 provides a structural support platform, serving as the installation foundation for various payloads, propulsion, and energy systems. Specifically, satellite body 1 has a cuboid-like structure, and its windward surface is a spherical cap curved surface structure 101. That is, the sides of satellite body 1 in the positive and negative X directions feature a spherical cap curved surface structure 101 with truncated edges to rectify wind currents. The sides of the satellite are smooth and without sharp edges, and can be less than half a sphere.

[0051] like Figure 2 As shown, the spherical cap structure 101 adopts a structural form obtained by the intersection of a sphere and the side surface of the celestial body. Specifically, the four side surfaces of the celestial body serve as cutting planes, cutting the sphere to form a square-sided structure with two windward top surfaces forming the spherical cap structure 101. The height H of the spherical cap is an adjustable design parameter used to determine the optimal shape while satisfying the balance of aerodynamic drag, attitude moment, and structural weight. The smooth, edgeless side surfaces of the celestial body prevent heat concentration at the corners during flight, which could affect the accuracy of instruments / equipment within the spacecraft's compartments. Furthermore, the side surfaces of the celestial body are made of materials with high temperature resistance and good thermal insulation properties. For example, titanium alloy or cobalt alloy can be used.

[0052] All remote sensing imaging payload components, such as the thermal infrared camera lens 2 and the data transmission antenna 7, are positioned on the Zs plane (the plane facing the ground) during in-orbit flight and arranged in a line along the Xs axis. The largest remote sensing imaging payload component is selected, and its installation is chosen to meet the Earth observation requirements (i.e., the Z-direction of the lens facing the ground and the push-broom direction). The smallest face of this payload component is then installed facing the wind, with the plane normal aligned along the X-direction. The remaining remote sensing imaging payload components are arranged in a line along the X-direction according to their Earth observation needs, with the primary goal of not increasing the cross-section in the windward direction. Other individual devices are arranged in a line along the windward direction. Multiple partitions extend along the X-direction inside the satellite body 1 to support the satellite body 1 and fix the payloads.

[0053] The thermal infrared camera lens 2 is installed facing the ground during satellite flight. The thermal infrared camera detector and the cooler 9 are close to the side of the satellite body 1 for easy connection to the main heat dissipation surface 5. The SAR antenna 4 is installed at the intersection of the main heat dissipation surface 5 and the edge facing the ground of the satellite body 1. The main heat dissipation surface 5 is installed on the side of the satellite body 1. The Hall electric thruster 3 of the +Xs plane is installed on the top of the satellite body 1. The Hall electric thruster 3 of the -Xs plane is installed on the bottom of the satellite body 1. The data transmission antenna 7 is installed facing the ground of the satellite body 1. The solar panel 6 is installed on the side of the satellite body 1, opposite to the main heat dissipation surface 5. By interchangeably maintaining the ascent and descent orbits of the Hall electric thruster 3 of the +Xs plane and the Hall electric thruster 3 (7) of the -Xs plane, direct jet maneuvering is possible during ascent and descent orbit maneuvers. Capability backup can also be achieved after the satellite turns around.

[0054] like Figure 3-5 As shown, both the plane of solar panel 6 and the plane of SAR antenna 4 are parallel to the flight direction. This side-facing orientation helps reduce aerodynamic drag caused by the frontal area of ​​solar panel 6 and SAR antenna 4 during flight. Furthermore, the weight of SAR antenna 4 can balance the weight of solar panel 6 on the opposite side, ensuring that the satellite's overall lateral center of mass is relatively close to the aerodynamic pressure center. Moreover, since both SAR antenna 4 and solar panel 6 face the flight direction from the side and are both subject to aerodynamic drag, their opposing moments can balance each other to some extent, reducing the impact of aerodynamic attitude interference during extremely low Earth orbit flight.

[0055] It should be noted that the above-mentioned three-axis orientation definition and functional layout design achieve organic coordination between remote sensing imaging requirements, heat dissipation balance, aerodynamic control and attitude stability. It is especially suitable for low-Earth orbit remote sensing mission platforms that integrate thermal infrared and SAR dual payloads and have high revisit cycle requirements. It can be widely used in the next generation of high-performance Earth observation satellite systems and has good engineering feasibility and commercial promotion prospects.

[0056] Specifically, the remote sensing imaging payload component uses a thermal infrared camera lens 2 to observe thermal infrared radiation from the Earth, with the lens facing the Zs axis. The thermal infrared camera detector and cooler 9 are also included. The detector receives infrared signals, and the cooler ensures its low-temperature operation and connects to the main heat dissipation surface 5 for heat conduction, thermally coupling with the main heat dissipation surface 5. A large-area main heat dissipation surface 5 is located on the -Ys plane, and its orientation is controlled by the satellite to ensure it is always shaded from the sun, ensuring effective heat dissipation at the cold end. The SAR antenna 4, used for synthetic aperture radar imaging, also serves as a counterweight and sunshade structure, and is positioned in the -Ys side-biased +Zs direction. The main heat dissipation surface 5, installed on the -Ys plane, provides long-term shaded heat dissipation, serving thermally sensitive equipment such as the thermal infrared cooling system. The Hall thruster 3 on the +Xs plane is located at the flight direction end, used for orbital ascent propulsion during normal flight and redundantly used during inverted flight. The Hall thruster 3 on the -Xs plane is located at the leeward end, redundant during normal flight and used for orbit correction during inverted flight. The propulsion tanks 8 are located within the satellite body 1, close to their respective thrusters, facilitating optimized liquid storage and center of gravity distribution. The data transmission antenna 7 is used for downlinking remote sensing data and maintains a ground-pointing orientation with the ground communication link. The solar panels 6 are installed on the +Ys side and tilted to provide energy while avoiding interference with the main heat dissipation surface 5.

[0057] To address the contradiction between efficient solar panel 6 receiving sunlight and minimizing aerodynamic drag under inclination conditions in a very low Earth orbit (LEO) orbit, this embodiment employs a thermal control coordination between the solar panel 6 and the SAR antenna 4. The SAR antenna 4 is positioned at the -Ys side offset +Zs direction for oblique Earth-view imaging while simultaneously blocking thermal radiation from the Earth. The solar panel 6 is only positioned on the +Ys side of the satellite, with its single-wing offset side facing the flight direction, avoiding direct wind contact and significantly reducing aerodynamic drag. A symmetrical weight design between the solar panel 6 and the SAR antenna 4 achieves weight balance. After deployment, the solar panel 6 is tilted relative to the horizontal plane at a specific angle, calculated based on the ratio of the SAR antenna 4's width to the satellite's side height. This tilt ensures that, regardless of the solar angle, as long as the solar panel faces the sun, the SAR antenna 4 remains unilluminated, thus maintaining a relatively low temperature.

[0058] Furthermore, by calculating the solar angle β based on the GNSS orbit position and the solar vector, and implementing periodic yaw maneuvers of 0° or 180°, the solar panel 6 is always in the sunlit side, while the main heat dissipation surface 5 is always in the shade.

[0059] The SAR antenna 4 and the solar panel 6 are located on the −Y / +Y sides respectively, with equivalent mass, achieving lateral counterweight balance. At the same time, the SAR antenna 4 can shield the Earth's thermal radiation, helping to protect the main heat dissipation surface 5 from being affected. This allows the solar panel 6 to achieve efficient solar irradiation, stable heat dissipation of the thermal infrared load, and minimal wind resistance for the entire satellite, adapting to the on-orbit operating environment of the MIO orbit where the solar angle changes significantly.

[0060] The tilt angle of the solar panel 6, in conjunction with the SAR antenna 4, ensures heat dissipation. The SAR flat panel antenna acts as a shield against Earth's thermal radiation, ensuring the operation of the main heat dissipation surface 5. The installation angle design of the single-wing solar panel 6 ensures a stable main heat dissipation surface 5 even when the sun moves along the sides of the orbit. During the active phase of satellite launch, the solar panel 6 is in a retracted state; after the satellite is launched and placed into orbit, it is in an extended state, forming an angle with the horizontal plane during flight. This angle is calculated by the ratio of the width B of the SAR antenna 4 to the side height L of the satellite, ensuring that the SAR antenna 4 and the main heat dissipation surface 5 are not exposed to sunlight during the satellite's rolling alignment with the sun throughout its entire lifespan.

[0061] The solar panel 6 and the main heat dissipation surface 5 are positioned on opposite sides of the satellite. Through periodic yaw maneuvers of the satellite's attitude, the solar panel 6 remains in direct sunlight during in-orbit flight, ensuring energy acquisition. The main heat dissipation surface 5 remains in the shaded area and is unaffected by the thermal radiation from the solar panel 6, ensuring efficient heat dissipation from the cold air. Even with changes in the solar angle, the SAR antenna 4 and the main heat dissipation surface 5 will not be exposed to direct sunlight, and the windward area will not be increased.

[0062] The satellite employs two independent and complete Hall thrusters, located at the front and rear of the satellite respectively, corresponding to the forward and backward directions during on-orbit flight. Specifically, Hall thrusters are installed at both ends of +Xs and −Xs. The rear thruster is used to ascend the orbit during forward flight, and the front thruster is used to descend the orbit during inverted flight. Both thrusters have the same thrust capacity and are equipped with corresponding storage tanks and redundant power supply circuits.

[0063] Example 2

[0064] like Figure 8 As shown, this embodiment, based on the overall configuration of a medium-inclination remote sensing satellite in Extremely Low Earth Orbit (EEO) as described in Embodiment 1, provides a design method for an EEO medium-inclination remote sensing satellite, including the following steps:

[0065] Establish the satellite's structural coordinate system (OsXsYsZs), defined as follows:

[0066] Os, the origin of the coordinate system: the installation geometric center corresponding to the four separation nuts on the side of the satellite structure body 1;

[0067] Zs axis: parallel to the side of the structure body 1, pointing towards the camera lens's line of sight;

[0068] Ys axis: perpendicular to the side of the main body 1 of the structure, pointing outward from the satellite body;

[0069] Xs axis: forms a right-handed relationship with the Ys and Zs axes;

[0070] Establish the satellite's center-of-mass orbital coordinate system (OoXoYoZo), defined as follows:

[0071] Origin Oo: The center of mass of the satellite;

[0072] Zo axis: pointing from the satellite's center of mass towards the Earth's center;

[0073] Xo axis: The direction of the satellite's flight;

[0074] Yo axis: forms a right-handed relationship with the Xo and Zo axes;

[0075] When the satellite is flying normally, the +Xs side is the windward side and the -Xs side is the leeward side; when the satellite is flying backward, the -Xs side is the windward side and the +Xs side is the leeward side. The term "normal flight" refers to the satellite's attitude with a 0° yaw, and the term "backward flight" refers to the satellite's attitude with a 180° yaw.

[0076] Before the rocket is assembled on the ground, the solar panel 6 is in a retracted state. The solar panel 6 is folded to the top surface of the satellite -Zs and fixed, and the SAR flat panel antenna is folded to the -Ys surface and fixed.

[0077] After the satellite enters orbit, solar panel 6 deploys to the +Ys side of the satellite and locks in place. The angle φ0 between the plane of solar panel 6 and its top surface is selected as follows: This value is generally 10° to 30°. The SAR antenna 4 rotates 90 degrees in the -Ys direction to form a sunshade and counterweight structure. After entering the working orbit, it begins to yaw and adjust according to the orbital solar angle. The main heat dissipation surface 5 always runs in the shade.

[0078] Specifically, the operating mode in orbit to obtain sufficient energy is as follows:

[0079] When the orbital solar angle |β|≤30°, the satellite attitude is kept oriented toward the Earth by the preset tilt of solar panel 6, and no rolling maneuver is required to align with the sun;

[0080] When the orbital solar angle |β| > 30°, in order to obtain better energy, the satellite can roll its maneuvering solar panels to track the projection of the solar vector onto the YoOoZo plane of the orbital coordinate system.

[0081] After the SAR flat panel antenna is extended 90° and parallel to the XsOsYs plane, it is locked in place. The exposed -Ys surface is the main heat dissipation surface 5, which is connected to the thermal infrared cooler or detector.

[0082] The satellite calculates the orbital solar angle β based on GNSS data and a solar vector model, controls the attitude yaw angle to switch flight modes, and adjusts the satellite's on-orbit attitude yaw maneuver cycle to ensure that the main heat dissipation surface 5 is aligned with the sun and the solar panel 6 is aligned with the sun. The specific steps are as follows:

[0083] To calculate the orbital solar angle, conventional satellites obtain orbital position information via a GNSS receiver. Using a solar model, the unit solar vector S = [Sox, Soy, Sooz] in the current centroid orbital coordinate system OoXoYoZo can be calculated. The angle between the solar vector and the orbital plane is the orbital solar angle β = asin(Soy). Figure 5 (as shown);

[0084] When β>0, it means that the sun is on the side of the orbital plane normal, and the satellite attitude needs to perform a 180° yaw maneuver to fly backwards;

[0085] When β < 0, it means that the sun is already on the negative normal side of the orbital plane, and the satellite can fly in a positive attitude.

[0086] When the temperature is near 0°, in order to avoid frequent switching of software logic, a small amount ε can be set to replace the zero-crossing judgment. Generally, ε is 3° to 5°. That is, if the previous flight was inverted, when β < -ε, the attitude is adjusted to fly normally; or if the previous flight was normal, when β > ε, the attitude is adjusted to fly inverted.

[0087] When maintaining the orbital altitude of a satellite in orbit, only one electric propulsion system located at the rear of the satellite body 1 is operational, and only one cathode of the electric propulsion system is operational to ensure the reliability and lifespan of the electric propulsion.

[0088] Those skilled in the art will recognize that the units and algorithm steps of the various examples described in conjunction with the embodiments disclosed in this invention can be implemented in electronic hardware, or a combination of computer software and electronic hardware. Whether these functions are implemented in hardware or software depends on the specific application and design constraints of the technical solution. Those skilled in the art can use different methods to implement the described functions for each specific application, but such implementations should not be considered beyond the scope of this invention.

[0089] Those skilled in the art will understand that, for the sake of convenience and brevity, the specific working processes of the systems, devices, and units described above can be referred to the corresponding processes in the foregoing method embodiments, and will not be repeated here.

[0090] Although embodiments of the invention have been shown and described, it will be understood by those skilled in the art that various changes, modifications, substitutions and alterations can be made to these embodiments without departing from the principles and spirit of the invention, the scope of which is defined by the appended claims and their equivalents.

Claims

1. A low Earth orbit (LEO) medium inclination remote sensing satellite, the satellite having Xs, Ys, and Zs directions, wherein the positive Xs direction is the forward direction of the satellite in orbit, the positive Ys direction is the negative normal to the orbital plane of the satellite in orbit, and the positive Zs direction is the Earth observation direction of the satellite in orbit; characterized in that: It includes the satellite body (1), remote sensing imaging payload assembly, solar panels (6), and attitude propulsion system: wherein: The satellite body (1) has a spherical cap surface structure (101) on its windward side. The spherical cap surface structure is a rectifier surface with continuous curvature and no broken line transition. Its height H is adjustable according to the rectification requirements of the satellite forebody. The remote sensing imaging payload assembly includes a thermal infrared camera lens (2) and a thermal infrared camera detector and a cooler (9) arranged along the flight direction. The thermal infrared camera lens (2) and the data transmission antenna (7) of the remote sensing imaging payload are arranged in series along the Xs direction and share the Zs axis of the Earth observation axis. The thermal infrared camera detector and the cooler (9) are arranged at the main heat dissipation surface (5) close to the -Ys direction and form a cold end heat dissipation channel through thermal connection. The solar panel (6) is positioned on the +Ys side of the satellite body after being deployed in orbit, opposite the main heat dissipation surface (5) to avoid mutual shading; and together with the SAR antenna (4) installed on the -Ys side after deployment with the array normal pointing in the +Zs direction, it forms a sunshade and counterweight system, so that the main heat dissipation surface is always in the shaded state in orbit. The attitude propulsion system includes Hall electric thrusters (3) arranged on the positive and negative Xs planes respectively. The attitude control system performs yaw maneuvers according to the orbital solar angle β. When |β| is greater than the set threshold ε, it switches between forward and backward flight attitudes to ensure that the solar panels are illuminated and the main heat dissipation surface is shaded. The orbit is maintained by the alternating operation of the two thrusters.

2. The ultra-low Earth orbit medium inclination remote sensing satellite according to claim 1, characterized in that: The spherical cap surface structure (101) is formed by the intersection of the circumscribed spherical surface and the side surface of the star. The spherical cap surface structure is cut by the side surface of the star to obtain a geometric body with square sides and a spherical cap top surface. The spherical cap curved surface structure (101) is integrally formed with the side plate of the star. The side plate is smooth and without sharp edges, which is used to reduce aerodynamic interference at the edges and prevent heat from concentrating on the side of the star at the sharp edges. The satellite body (1) is made of a material with high temperature resistance and good heat insulation.

3. The ultra-low Earth orbit medium inclination remote sensing satellite according to claim 1, characterized in that: A heat-conducting connection plate is provided between the thermal infrared camera detector and the refrigerator (9) and the main heat dissipation surface (5). The heat-conducting connection plate is fixed to the main heat dissipation surface (5) through a thermal coupling component. When the satellite attitude is adjusted, the direction of the cold end heat flow is consistent with the normal direction of the heat dissipation surface, thereby reducing the temperature drift of the refrigerator and improving thermal stability.

4. The ultra-low Earth orbit medium inclination remote sensing satellite according to claim 1, characterized in that: The solar panel (6) is connected to the satellite body (1) via a hinge assembly. After unfolding, the angle between the plane of the solar panel (6) and the top surface of the satellite body is 10° to 30°. This is used to maintain stable incident light conditions when the orbital solar angle changes, and the mechanical limiting structure constrains the sway amplitude of the solar panel (6) during attitude switching to not exceed 3°.

5. The ultra-low Earth orbit medium inclination remote sensing satellite according to claim 1, characterized in that: The SAR antenna (4) has a low-absorption reflective layer on its surface to reflect thermal radiation from the sun and the earth; the reflective layer is arranged parallel to the +Zs surface to ensure the thermal stability of the main heat dissipation surface (5).

6. The ultra-low Earth orbit medium inclination remote sensing satellite according to claim 1, characterized in that: The Hall electric thrusters (3) on the positive and negative Xs planes are equipped with independent cathode modules and propulsion storage tanks (8). The satellite body (1) performs a life assessment based on the cumulative working time and thrust deviation of the two thrusters. When the thrust performance of either thruster drops below a preset threshold, the other thruster is automatically switched to work.

7. The ultra-low Earth orbit medium inclination remote sensing satellite according to claim 1, characterized in that: When the attitude control system performs yaw maneuvers, it uses the orbital solar angle β and the output power of the solar panel (6) as a joint criterion. When the output power of the solar panel drops by more than 10% after the attitude switch, it performs local roll correction to maintain the optimal value of the solar panel's illumination angle, thereby achieving joint optimization of energy and attitude.

8. The ultra-low Earth orbit medium inclination remote sensing satellite according to claim 1, characterized in that: During the ground assembly stage, the solar panel (6) and SAR antenna (4) are folded and fixed to the satellite's -Zs and -Ys surfaces, respectively. After entering orbit, they are sequentially unfolded by the release mechanism, so that the SAR antenna (4) first rotates 90 degrees around -Ys and then becomes parallel to the XsOsYs surface. Subsequently, the solar panel unfolds to the preset angle and is locked. The attitude disturbance during the unfolding process is minimized by controlling the unfolding sequence.

9. The ultra-low Earth orbit medium inclination remote sensing satellite according to claim 1, characterized in that: The main heat dissipation surface (5) is set on the -Ys plane, and its normal direction is opposite to that of the normal direction of the solar panel (6) in the Ys direction. When the orbital solar angle β changes, the incident angle of the heat dissipation surface is automatically adjusted by attitude yaw so that the heat dissipation efficiency is maintained within the design range.

10. A design method for a very low Earth orbit (ULE) medium inclination remote sensing satellite, used to design the ULE medium inclination remote sensing satellite according to any one of claims 1-9, characterized in that: Includes the following steps: S101: Determine the track and coordinate reference; Based on the mission coverage and resolution requirements, a medium-inclination orbit suitable for very low Earth orbit operation is selected, and a structural coordinate system is established, in which: the remote sensing satellite has Xs direction, Ys direction and Zs direction, the positive direction of Xs direction is the forward direction of the remote sensing satellite in orbit flight state, the positive direction of Ys direction is the negative normal of the orbital plane of the remote sensing satellite in orbit flight state, and the positive direction of Zs direction is the Earth observation direction of the remote sensing satellite in orbit flight state. S102: Determine the aerodynamic shape of the satellite; In the coordinate system, the windward surface of the satellite is a spherical cap surface structure (101) with continuous curvature, and the spherical cap surface structure (101) is smoothly connected to the side plate of the satellite to form a straightening surface; S103: Configure imaging and heat dissipation system; arrange thermal infrared camera lens (2) and data transmission antenna (7) along the Xs direction and share the ground axis Zs, set thermal infrared camera detector and refrigerator (9) on the side of the main heat dissipation surface (5) close to -Ys, and form a heat dissipation path from the cold end to the heat dissipation surface through thermal connection to achieve thermal control and structural coordination. S104: Design energy and sunshade structure; arrange solar panels (6) on the +Ys side and arrange SAR antennas (4) on the -Ys side at the +Zs direction, which have imaging, sunshade and counterweight functions to ensure that the main heat dissipation surface (5) operates in the sunless position on the track. S105: Set the control logic of the attitude propulsion system; install Hall electric thrusters (3) on the positive and negative Xs surfaces, calculate the orbital solar angle β based on GNSS and solar vector, and perform yaw switching when |β| exceeds the threshold ε to achieve positive-to-negative flight conversion, ensuring that the solar panels are illuminated, the main heat dissipation surface (5) is in the shade and maintains propulsion redundancy; S106: Perform structural counterweighting and performance verification; perform lateral counterweighting according to the mass distribution of solar panels (6) and SAR antennas (4), and comprehensively verify aerodynamics, thermal control and attitude stability to determine the overall satellite coupling balance and complete the design of the ultra-low orbit inclination remote sensing satellite.

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