Solar energy tracking system

Abstract

The antenna system is configured for use in a low Earth orbit (LEO). The system has multiple antenna satellites coupled together to form a phased array. Each of the multiple antenna satellites is equipped with an antenna body having an antenna and a solar cell. A processing device determines the orientation of the multiple antenna satellites and, based on an analysis of the solar cells of the antenna bodies facing the sun and the antennas of the antenna bodies facing the Earth, arranges the phased array in the corresponding orientation to maintain torque equilibrium of the phased array.

Description

Technology Field

[0001] This application claims priority to U.S. provisional application No. 62 / 976,127 filed on February 13, 2020, the entire contents of which are incorporated herein by reference. Background Technology

[0002] U.S. Patent Application No. 9,973,266 and U.S. Patent Application Publication No. 2019 / 0238216 describe a system for assembling a large number of small satellite antenna assemblies in space to form a large array. The entire contents of Patent Application '266 are incorporated herein by reference. As disclosed in Patent Application '266, FIGS. 1a and FIGS. 1b illustrate a satellite communication system (100) having an array (300) of small satellites (10) and a central or control satellite (200). The small satellites (10) communicate with an end user (500) within a footprint (400) on Earth, and also communicate with the control satellite (200), which in turn communicates with a gateway (600) of a base station. Each small satellite (10) may include, for example, a processing device (e.g., a processor or a controller) and one or more antenna elements. And the control satellite (200) may include a processing device and one or more antennas or antenna elements. Accordingly, the large planar phase array (300) is composed of a plurality of satellite modules (10). The problem to be solved

[0003] U.S. Patent Application Publication No. 2016 / 0065006 by Woods illustrates a solar energy conversion and transmission system and method. Woods has an oriented solar cell capable of receiving sunlight, but does not teach how this is achieved, or even what kind of orientation it is. means of solving the problem

[0004] Accordingly, one object of the present disclosure is to provide a wide, flat space structure having solar cells on one side and antennas on the other side. Another object of the present disclosure is to provide such a structure having a control system and method for orienting the solar cell side toward the sun and the antenna side toward the earth. A further object of the present disclosure is for the control system to stabilize gravity gradient torque in a gyroscope manner. Another object of the present disclosure is for the structure and the control system to operate in low Earth orbit (LEO) where gravity gradients are large and are not easily overcome by other means, such as a propulsion system. Another object of the present disclosure is for the structure to offset gravity gradient torque using gyroscope torques, provided that the sunlight received by the solar cells is enhanced and phase array communications with users on the ground are not excessively interfered with. Brief explanation of the drawing

[0005] Figures 1a and 1b are drawings illustrating a large array. Figure 2 is a drawing illustrating a large planar structure having an orbital track in space. FIG. 3a is a drawing illustrating the first stable orientation. FIG. 3b is a drawing illustrating the second stable orientation. FIG. 3c is a diagram illustrating an unstable equilibrium orientation. FIG. 3d is a drawing illustrating a gyroscope-stabilized orientation. FIG. 3e is a drawing illustrating an orientation without gyroscope stability. FIG. 3f is a diagram illustrating torque and rotation in a gyroscope-stabilized orientation. Fig. 4a is a diagram illustrating solar movement and radio-frequency communication cone and unstable equilibrium orientation. FIG. 4b is a diagram illustrating solar movement and radio-frequency communication cone and gyroscope-stabilized orientation. FIG. 4c is a diagram illustrating solar movement and radio frequency-communication cone and the first stable orientation. FIG. 4d is a diagram illustrating solar movement and radio-frequency communication cone and second stable orientation. FIG. 5 is a drawing illustrating an antenna assembly. FIG. 6 is a block diagram of an antenna assembly. Fig. 7a is a block diagram of a control satellite. FIG. 7b is a block diagram of a gyroscope component. Figures 8 and 9 are flowcharts. Specific details for implementing the invention

[0006] In describing the exemplary and non-limiting embodiments of the present disclosure illustrated in the drawings, specific terms will be used for clarity. However, the present disclosure is not intended to be limited to the specific terms so chosen, and each specific term should be understood to include all technical equivalents that operate in a similar manner to achieve a similar purpose. Some embodiments of the present disclosure are described for illustrative purposes and it is understood that the present disclosure may be implemented in other forms not specifically illustrated in the drawings.

[0007] As illustrated in FIG. 2, the present disclosure includes a structure (100) in an orbital track (402) relative to the Earth. In one embodiment, the structure (100) may be a spacecraft such as a data center or satellite, or a large antenna or satellite array as illustrated in patent application '266. The structure (100) has a number of small structures (10), such as a satellite or antenna assembly (10) illustrated in FIG. 5, which are rigidly connected to form a single array together.

[0008] Referring to FIG. 5. The present disclosure is a system, for example, a spacecraft (10) orbiting the Earth in one embodiment. In one embodiment, the spacecraft is a structure such as a satellite or an antenna. The structure (10) has photovoltaic panels that receive solar energy from the sun and generate solar power for use in the structure. In one example, the structure (10) has an antenna body (11) having a communication layer (12) comprising antennas or antenna elements (18), an optional intermediate layer (16) comprising a rigid structure and electronic components, and a photovoltaic layer (14) comprising one or more solar cells. Each layer is a flat rectangular or square panel or tile, and the entire structure is flat and rectangular or square. The antenna body (11) has an interior or internal space.

[0009] Referring to FIG. 6, one or more of the antenna structures (10) may include a number of electronic components, including a processing device (20), a front-end module (FEM) (22), a flight computer (24), one or more sensors (26), one or more flight actuators (28), a solar collector, and a link to an external system. The solar collectors / cells generate energy from the sun. Excess power from the solar cells can and / or charge the battery, or the battery can provide power to the electronic components. Each FEM (22) is associated with each antenna of the communication layer (12) to transmit / receive signals. The processing device (20) controls the overall operation of the electronic components, including the reception / transmission of signals through the FEM (22). The flight computer (24) receives signals regarding the orientation and position of the antenna structure (10) from the sensors (26). In response to the received sensor signals, the flight computer (24) determines the current orientation and position and transmits control signals to the actuators (28) to properly position the antenna structures (10).

[0010] In one embodiment, the control satellite (200) has access to and authority over all sensor data and actuators to control the orientation of the entire phase array (100). The control satellite flight computer (224) reads all sensor data and issues commands to all actuators. Each small structure (10) only reads its own sensor data to attempt and correct its orientation within the phase array (100). Additionally, as disclosed in the jointly pending patent application filed on February 13, 2020, the entire contents of which are incorporated herein by reference, FC (24), sensors (26), and actuators (28) may be used to ensure the flatness of the array. In one embodiment, the control satellite (200) may be detached from the array (100) and connected to the array (100).

[0011] The orientation of the phase array (100) is controlled by the flight computer (224) of the control satellite (200). This flight computer (224) receives sensor data from readings of a set of its own sensors (226) and any other sensors (26) to determine the orientation of the phase array (100). Based on the determined and desired orientation of the phase array (100), the flight computer (224) commands its own actuators (228) and the actuators (28) of the small structures (10) as needed to ensure the proper orientation of the phase array (100). In one embodiment, the sensor (26) may be a GPS that provides position data. In another embodiment, the sensor (26) may be an IMU, such as an accelerometer, that provides acceleration data from which position data can be determined.

[0012] Returning to FIG. 5, the communication layer (12) has an outward-facing surface having antenna elements facing the earth (bottom) to communicate with user devices (e.g., mobile phones). The solar layer (14) has an outward-facing surface having a solar cell that generates solar power for use by electronic components shown in FIG. 6, such as a processing device (20), antennas, antenna front-end modules (22), flight computer (24), sensors (26), and actuators (28). The outward-facing surface of the solar layer (14) faces in an orientation (ceiling) opposite to that of the outward-facing surface of the communication layer (12). The outward-facing surface of the solar layer (14) forms the upper surface (15) of the antenna structure (10) (and coupling structure (100)), and the outward-facing surface of the communication layer (12) forms the lower surface (17) of the antenna structure (10) (and coupling structure (100)).

[0013] As the structure (100) orbits the Earth, the sun side of the structure will be angled toward the sun. By selecting an appropriate deflection angle (θ) as illustrated in FIG. 4b, the solar cells on the upper surface of the structure (100) will receive improved solar exposure through the orbit of the structure (100) around the Earth. In this way, the deflection angle controls the pointing of the lower surface of the structure (100) relative to the Earth range (400) and, by extension, controls the phase array (300).

[0014] The rigid connection of each small structure (10) requires that the orientation of the structure (100) determines the orientation of each individual small structure (10). The orbital track (402) is approximately a circular orbit. The velocity vector is considered as the unit vector of the direction of movement of the structure (100). The zenith vector is considered as the unit vector pointing away from the center of the Earth with the spacecraft at the center. The cross-track vector is considered as the unit vector orthogonal to both the zenith vector and the velocity vector.

[0015] One objective of the present disclosure is to orient a structure (100) in a controlled manner based on the following three factors:

[0016] 1. The first factor is that the structure (100) receives as much sunlight as possible from the upper surface (15), and the sunlight is converted into electricity using solar cells on that surface. Here, the ultimate goal depends on the desired application. If the normal vector for the upper surface (15) is parallel to the satellite-sun vector, the maximum possible sunlight is received at that moment. In other embodiments, this condition of maximum received sunlight is not required and is achieved only for transient moments. As the normal vector for the upper surface (15) moves away from being parallel to the sun, less sunlight is received. How little is received depends on the design, and how much sunlight is required also depends on the design. The criteria for determining a sufficient amount of sunlight depend on how much power is consumed when the satellite is turned on and how long the satellite is turned on or off.

[0017] One goal of this element (in the absence of other elements) is to ensure that the normal to the upper surface (15) is always parallel to the satellite-sun vector. Since any subsequent deviations are monotonic with respect to the received power, and the closer the received power is to parallel, the higher it is, the closer it is to parallel, the better. However, this element can be satisfied even when the satellite does not maximize the amount of sunlight being received, such as when the satellite's power consumption requirements are low (e.g., when it is turned off or consuming little power).

[0018] 2. The second element is to ensure that the structure (100) points toward the lower surface (17) as close as possible to the center of the Earth, which allows the antenna elements of the phase array (300) to effectively radiate radio frequency signals to the area (400) of the user (500) on the Earth's surface. One goal (in the absence of the other two elements) is to point perfectly toward the center of the Earth. The cost of deflection from this perfect orientation depends on the design-related elements. In particular, pointing 10° is worse than pointing 5°, and if pointing 180° (the antenna-side pointing toward space), it cannot operate. If pointing 90°, it can still communicate with some users on the ground, but it may not be sufficient because at least half of the users are cut off from communication.

[0019] However, there are a few exceptions; for example, since users on the ground do not need to be evenly distributed, pointing away from the center of the Earth is not always worse. If all users are located north of the satellite, it would be useful to point the antenna north instead of toward the center of the Earth. Pointing toward the center of the Earth is a general rule that can be applied.

[0020] 3. The third element is to have minimal active control over the structure (100). Mechanically controlling the attitude of the satellite requires a torque-applying actuation system, and providing such active control in space is cumbersome. One goal is to keep the satellite in a torque-balanced state, which requires minimal active control. This, in turn, reduces the cost of actuators and power in orbit. Torque balance can be achieved, for example, by gyroscope stabilization.

[0021] To explore Element 3 in greater depth, it is best to look for analogies regarding mechanical control. A broom standing upright on the floor (inverted pendulum) requires minimal input control when maintained very accurately vertically, whereas deflections mean that a lot of work must be done to keep it upright. A horizontal broom fixed at one corner requires a very large counter-torque to prevent it from falling, and gravity continues to exert force to try and drop it. This very large torque is necessary to control the broom accurately or inaccurately. A broom suspended from above requires negligible control because gravity naturally causes it to fall to this position.

[0022] Because available technologies, such as star trackers, allow for highly accurate attitude knowledge of the structure (100), stable orientations (such as a broom suspended from above) (Fig. 3a) and unstable equilibrium orientations (such as an upright broom held from below) (Fig. 3b) are considered to satisfy factor 3. Any non-equilibrium orientation (e.g., Fig. 3c) requires a constant determinate torque (such as a horizontal broom held from one corner), and generally, constant torques are cumbersome to provide in the context of spatial operation. A notable exception to this description is gyroscope torque, but these are inherently limited. Gyroscope torques are generated by vertical rotation axes and are therefore not universally applicable in all cases. For example, inertially stable objects (those not rotating about the sun and stars) cannot generate gyroscope torques because they do not rotate.

[0023] Equilibrium orientation exists when there is no net torque acting on the structure (100). However, this is achieved by using a gyroscope torque that allows certain non-equilibrium orientations (orientations with net torque acting on the structure) to be equilibrated. When the gyroscope torque corresponds exactly to an external torque (e.g., a gravity gradient), it causes the net torque on the structure to become zero. The category of orientations that can be equilibrated (zero net torque on the structure) using gyroscope torques is referred to as gyroscope stabilization orientations.

[0024] Figure 3e illustrates an orientation where the gravitational gradient torque cannot be countered by gyroscope torques. The orientation in Figure 3e is similar to that depicted in Figure 3a, but the satellite has rotated around a vertical (cross-track) vector. Since Figure 3a represents a stable orientation (like a broom hanging from above), Figure 3e represents an unstable orientation (like a broom held from one edge). For this unstable orientation to be maintained, an external torque (gravitational gradient) must be continuously countered. However, since both the gravitational gradient torque and the orbital motion relate to the vertical vector, gyroscope torques cannot be used for this purpose (gyroscope torques are always perpendicular to the axis of rotation).

[0025] FIG. 3f illustrates the orientation of gravity gradient torque and, in the case of flat bodies, the natural gyroscope torque that inevitably exists around the velocity vector. Without intentional control of the gyroscope torque, the structure (100) will deviate from its position (like a broom held by one edge). Intentional manipulation of the gyroscope torque allows it to accurately respond to the gravity gradient torque, thereby bringing the structure (100) into a state of equilibrium. There are two methods for intentionally manipulating the gyroscope torque.

[0026] 1) Rotate the structure (100) to generate additional rotational movement and change the net gyroscope torque to correspond to the gravity gradient torque.

[0027] 2) Rotate internal components to generate additional rotational movement and change net gyroscope torque to correspond to gravity gradient torque.

[0028] In each case, bringing the structure (100) into a state of equilibrium allows the structure to be easily controlled when the orientation of the structure (100) is accurately known. This equilibrium can be achieved regardless of the angle (θ) shown in FIG. 3f, thereby enabling an infinite number of orientations, each having a different angle (θ) between 0° and 360°. These gyroscopically stabilized orientations can satisfy elements 1-3 better than the stable orientations (Fig. 3a-Fig. 3b) and unstable (Fig. 3(c)) equilibrium orientations that exist without gyroscopic stabilization, depending on the orientation of the sun relative to the orbit of the structure (100).

[0029] Orientations of large flat structures in space

[0030] The application of the structure (100) requires radio-frequency communication with the surface of the Earth. This is the basis for the definition of element 2, and implies that when the structure (100) orbits the Earth through the orbital track (402) (Fig. 2), it must rotate with the orbital movement and change its orientation to ensure that the lower surface (17) points toward the surface of the Earth (400). This creates a rotational reference frame that maintains the communication range on the Earth (400) at the same location for the structure (100) in a manner most suitable for describing the orientation of the structure (100). The rotational reference frame is based on a velocity vector (direction of movement of the structure (100)), a zenith vector (a vector pointing from the center of the Earth through the center of the structure (100)), and a vertical direction vector (orthogonal to the velocity and zenith vectors) and is used in Figs. 3a through 3f, Figs. 4a and 4b. In the rotational reference frame, orientations that allow elements 1 through 3 to be satisfied can be determined.

[0031] In the first stable orientation shown in FIG. 3a and FIG. 4c, the normal to the lower surface (17) of the structure (100) is the direction of the velocity vector (or the vector opposite to the velocity in the equivalent mirrored orientation). This implies that the lower surface is 90° from the range (400) on the earth (i.e., the lower surface (17) is facing the page), and thus element 2 (the lower communication side surface (17) facing the earth) is not satisfied.

[0032] In the zenith, velocity, and vertical reference frames, the Sun follows a conical pattern once per orbit. The half-angle (α) of this cone is determined by the beta angle. The beta angle is determined by the orbit and the position of the Sun relative to the Earth, where the beta angle is the angle of closest approach between the zenith vector and the Sun vector (the vector pointing from the satellite to the center of the Sun). The beta angle is not considered a controllable variable because the orbital choice is determined by other factors (such as which region of the Earth the satellite will orbit).

[0033]

[0034] At small beta angles, the noticeable movement of the sun is a broad cone passing near the velocity, bottom, half-velocity, and zenith vectors in each orbit of the repeating pattern. For the first stable orientation illustrated in FIG. 3a, the following conclusion can be drawn: when the sun is near the half-velocity vector (or the velocity vector of the equivalent mirrored orientation), the upper surface (15) will receive a large amount of sunlight once per orbit. At large beta angles, the noticeable movement of the sun follows a narrow cone pattern, so that it remains close to the cross-track vector perpendicular to the upper surface (15). In this case, almost no sunlight will reach the upper surface. For orbits observing large beta angles, Factor 1 is not satisfied. Factor 3 is clearly satisfied because it is a stable orientation (like a broom hanging from above) that requires negligible control.

[0035] In the second stable orientation shown in FIG. 3b, the normal to the lower surface is aligned with the vertical vector (or the semi-vertical vector in the equivalent mirrored orientation). This means that the lower surface (17) points 90° from the Earth range (400) (in a manner similar to the first stable orientation), and Element 2 (the lower communication side surface (17) facing the Earth) is not satisfied. As shown in FIG. 4d, the normal to the upper surface (15) is aligned with the semi-vertical vector (or the vertical vector in the equivalent mirrored orientation). This means that at small beta angles, the sun is perpendicular to the upper surface (15) and sunlight barely reaches the solar cells. At large beta angles, the sun is aligned with the normal to the upper surface, and sunlight is received excellently. For orbits observing small beta angles, Element 1 is not satisfied. Element 3 is clearly satisfied because it is a stable orientation (like a broom hanging from above) that requires negligible control.

[0036] In the unstable equilibrium orientation illustrated in FIG. 3c, the normal to the lower surface (17) points directly toward the Earth range (400), satisfying element 2 (i.e., the lower communication side surface (17) faces the Earth). In this precise orientation, no gravitational gradient torque exists, but a deflection from this point will generate a gravitational gradient torque that pulls the structure (100) away from this position. This is because when the structure (100) is in the illustrated precise orientation, each part of the structure is approximately equidistant from the center of the Earth, and gravity acts equally across the surface. The time when part of the structure is closer to the Earth and another part is further away is when the structure (100) is deflected from this position. Since gravity acts at the square of the distance from the center of the Earth, the closer part of the structure will experience greater gravity than the farther part of the structure. This differential application of gravity generates a torque that rotates the structure. Such behavior makes this orientation an unstable equilibrium point.

[0037] With precise active control of the structure (100), this orientation can be maintained without the constant application of torque satisfying element 3 (i.e., minimizing active control). As illustrated in FIG. 4a, noticeable solar movement and, by extension, sunlight received by the upper surface of the structure (100) is a function of the beta angle (β). When the beta angle is small, element 1 (i.e., the upper sun-side surface (15) is facing the sun) can be satisfied, but for a large beta angle approaching 90°, element 1 is not satisfied.

[0038] In the gyroscopically stabilized orientation shown in FIG. 3d, the structure (100) is deflected by a controlled angle (θ) from the orientation of FIG. 3c. Depending on the specific limitations of the phase array (300) forming the radio frequency signals, this deflection can still satisfy element 2 (i.e., the lower communication side surface (17) facing the Earth) when the deflection angle (θ) is small.

[0039] In FIG. 4b, it can be seen that for large beta angles, this deflection improves the exposure of the upper surface (15) of the structure (100) to sunlight, thereby better satisfying Factor 1 (of course, the best relationship of the deflection angle with respect to beta is when theta is equal to beta). However, as shown in FIG. 3f, the gravity gradient torque ( ) is generated due to the orientation of the structure (100). This gravitational gradient torque is applied to the velocity vector, causing the structure (100) to start rotating and inducing the orientation shown in FIG. 3b. Because the velocity vector is perpendicular to the orbital rotation of the satellite (100), the gyroscope torque ( ) can be used to respond to gravity gradient torque. By using gyroscope torque, gravity gradients can be responded to without cumbersome requirements for the attitude control system of the structure (100), thus satisfying element 3 (i.e., minimizing active control).

[0040] Such gyroscopic torques, utilizing the orbital rotation of a structure, can be generated permanently in an idealized manner without any arbitrary input (power, control, etc.). In practice, gyroscopic torques require some maintenance to ensure that rotational speed is maintained despite internal friction or similar effects. This maintenance is minimized compared to the cost of using existing alternative approaches to generate constant torques (electric propulsion, chemical propulsion, electromagnetic torque, etc.) in space, which are unsuitable for large structures in LEO orbits.

[0041] In the orientation without gyroscope stabilization shown in FIG. 3e, the gravity gradient is generated around a vertical vector that is also the axis of rotation. Since gyroscope torques can only be generated perpendicular to rotation, rotation around the vertical vector cannot generate gyroscope torque around the vertical vector. In this case, the gyroscope torques cannot stabilize the system, and Factor 3 (i.e., minimization of active control) is not satisfied.

[0042] Control of large flat structures in space

[0043] Control of large flat structures (100) in space requires attitude determination, which is a process in which the structure (100) measures its own orientation. As illustrated in FIG. 7a, the control satellite (200) may have a single sensor or a set of sensors (226) (e.g., star trackers, solar sensors, inertial measurement units) used to determine the orientation of the structure (100). This attitude information is interpreted by a flight computer (224), and the flight computer outputs command signals to an actuator or a set of actuators (228) (e.g., torque rods, reaction wheels, momentum wheels) to increase, decrease, or eliminate rotational movement as needed. In one embodiment, the flight controller (224) is a processing device such as a processor or controller. Although the flight controller (224) is depicted as a separate component such as the processing device (220), the flight controller (224) may be combined with the processing device (220).

[0044] As previously described, maintaining a stable orientation requires minimal control because the structure (100) will naturally be induced toward these orientations. However, no stable orientation of the structure (100) can satisfy element 2 (i.e., the lower communication side surface (17) faces the Earth). Unstable orientations without gyroscope stabilization require constant torque, which overdemands the actuator system and cannot satisfy element 3 (minimize active control). This could be the unstable equilibrium orientation shown in FIG. 3c, or the gyroscope-stabilized orientation shown in FIG. 3d, where all elements 1-3 can be satisfied. In the case of a gyroscope-stabilized orientation, it is essential that the gradient torque and the required corresponding equilibrium gyroscope torques are limited.

[0045] The gravitational gradient torque is the result of unbalanced gravity across a structure and is calculated by the following formula. α is the semi-major axis of the orbit, μ is the Earth's standard gravity parameter, and is the unit vector of the center of mass of the structure toward the center of the earth, and I is the matrix of inertia of the structure (100).

[0046]

[0047] As can be seen from the formula for gravitational gradient torque, the magnitude of the gravitational gradient is determined by r to the cube of the semi-major axis. This means that a structure in low Earth orbit (LEO) at a semi-major axis of about 7,000 km (from the center of the Earth) generates gravitational gradient torques that are about 200 times greater than an equivalent structure in geosynchronous orbit (GEO) at a semi-major axis of about 42,164 km (from the center of the Earth).

[0048] Gyroscope torques are generated by two simultaneous rotation axes on the main body. In the absence of gyroscope stabilization, large, flat structures naturally exhibit some gyroscope torque when arranged in the orientation shown in Fig. 3f. However, this gyroscope torque naturally corresponds to the gravity gradient torque only for very specific structural shapes. Generally, active control of the second rotation axis is required to develop gyroscope torque corresponding to the gravity gradient.

[0049] If the structure (100) is symmetric with respect to the normal vector, it can rotate around the normal vector while maintaining the gyroscopically stabilized orientation shown in FIG. 3d and FIG. 7b. This type of rotation (ω x )(rotating like a frisbee) is a first means of generating a gyroscope torque capable of responding to a gravity gradient. A second means of the structure (100) generating the gyroscope torque is the rotation (ω) of an internal component (i.e., momentum wheel) with respect to a normal vector. wheel Through ), it allows gyroscope stabilization of the structure (100) regardless of whether it is symmetric with respect to the normal vector. Rotation (ω x )(ω wheel ) can be thousands of rotations per minute.

[0050] As illustrated in FIG. 7b and FIG. 3d through 3f, in two examples (where the entire phase array rotates (without a gyroscope wheel), and where the gyroscope component rotates (where the phase array does not rotate)), the structure has rotation components (illustrated) about the x and y axes, which are defined in the formula below. The wheel (in the second example) rotates about the x-axis, which is also illustrated, but the entire phase array (300) and the control satellite (200) remain placed (fixed) within the rotation frame (i.e., while orbiting the Earth, they rotate (once per orbit) to maintain the orientation illustrated in FIG. 4b throughout the entire orbit). It is the gyroscope component that rotates, but it exists only within the control satellite.

[0051] A first means of gyroscope stabilization of the structure (100) is through the rotational speed of a specific combination of structures (100). As illustrated in FIG. 3f, in a rotational reference frame defined by a velocity vector, a zenith vector, and a vertical direction vector, the structure (100) can rotate around a normal vector (like a frisbee) without affecting the tilt (θ) with respect to the axis of inertia rotation. This entire reference frame also rotates with the structure (100) around the Earth once per orbit, driven by the orbital motion of the structure (100). It is the combination of these two rotations that generates the controlled gyroscope torque. While the rotation of the reference frame is driven by orbital motion and cannot be controlled, the rotation of the structure about the normal vector can be directly controlled so that the net gyroscope torque balances the gravity gradient torque, creating equilibrium for the structure (100) in the rotational reference frame. It should be noted that the gravity gradient torque and the gyroscope torque are not balanced in the inertia reference frame.

[0052] The angle between the ceiling vector and the normal vector is defined as θ. The x-axis is defined as collinear with the normal vector with respect to the upper surface. The z-axis is defined as being aligned with the velocity vector. The y-axis is defined as being orthogonal to the x and z axes using the right-hand rule.

[0053] In the first example of control in which the rotation of a structure (100) generates a gyroscope torque that balances the gravity gradient torque, the rotation of the structure with respect to the normal vector is according to the following formula (ω x It is defined as ). I x , I y and I z ε₀ and y₀ are the moments of inertia of the structure in the x, y, and z directions, respectively, and ω orbit is the orbital rotation speed, and ω x , ω y , and ω z is the rotational velocity of the spacecraft around the x, y, and z axes, respectively. This formula is I, a property of a large flat structure symmetric with respect to the z-axis. y = I z = It is valid only in this case.

[0054]

[0055] Through the rotation of the entire structure (100), it will remain stationary in the rotation reference frame shown in FIG. 3d, except for a frisbee-like spin about the normal vector. The simplicity of these formulas should not be considered trivial, because this is all based on the Earth standard gravity parameter (μ), the semi-major axis of the orbit (α), and the moment of inertia of the structure (I x , I y , I z This is because it is an elegant result of gravitational gradient, orbital dynamics, and gyroscope precession, which are functions of ).

[0056] Therefore, in this first example of gyroscope stabilization, the gyroscope component (229) is not present. Instead, magnetic torque rods (part of 228) apply torque to the entire phase array to rotate the entire phase array.

[0057] A second example of gyroscope stabilization uses a momentum wheel (229) inside the structure (100) instead of the rotation of the structure (100) itself. In this way, the structure (100) is kept completely stationary in the rotating reference frame shown in FIG. 3d. The combination of the rotation of the momentum wheel and the orbital rotation of the structure (100) generates a gyroscope torque that can be actively controlled to balance the gravity gradient torque. It should be noted that although fixed to the rotating reference frame, the structure (100) can naturally generate gyroscope torque on its own according to its shape.

[0058] In the second example of control, the rotation of the entire structure (100) about the normal vector is replaced by the rotation of the internal components about the same axis. The momentum wheel (229) inside the rotated satellite structure will generate a functionally equivalent gyroscope torque. The momentum wheel (L wheel The angular momentum required by ) and the rotational speed used are determined by the following formulas.

[0059]

[0060] This second example of gyroscope stabilization allows the entire structure to be maintained in a completely fixed state (without frisbee-type rotation) relative to rotation, so that I z ≠ I y It allows the maintenance of a gyroscopically stabilized posture for an asymmetrical structure (100).

[0061] As illustrated in FIGS. 4a and 4b, gyroscope-stabilized orientations can provide better solar exposure to the upper surface of the structure (100), but the deviation angle (θ) is limited by the requirement that the lower surface of the structure (100) be oriented toward the Earth range (400). For non-zero beta angles (β), the angle offsets increase the power generated by the solar cells on the upper surface of the structure (100) in a very predictable manner, but degrade radio-frequency communication between the Earth range (400) and the phase array (300) in a non-linear design-specific manner.

[0062] In one embodiment of this second example, the control satellite (200) includes a gyroscope component (229) used to generate a gyroscope force that establishes and maintains equilibrium with respect to the structure (100). The gyroscope component (229) is connected to a flight computer (224) and / or a processor (220), which controls the operation of the gyroscope component (229). In one embodiment, the gyroscope component (229) may be a momentum wheel, and the flight computer (224) and / or the processor (220) control the gyroscope component (229) to turn it on or off and to rotate it at a desired speed. The gyroscope component (229) may be located inside the control satellite (200). In one embodiment, the control satellite (200) is located approximately in the center of the phase array (300) and is fixedly mounted to the phase array (300). The momentum wheel rotates within the control satellite (200) relative to the control satellite housing.

[0063] For example, the momentum wheel may rotate about a rotation axis rotatably mounted by a gear in the housing of the control satellite (200). Thus, the housing of the control satellite (200) does not rotate about the phase array (300), but the momentum wheel rotates about the housing. However, the entire phase array (300) remains stationary about the Earth during orbit (except to rotate to maintain the same orientation (Fig. 3d). The gyroscope component (229) rotates slowly (from hours to days) about the corresponding torque provided by the magnetic torque rod (part of 228). In another embodiment, the momentum wheel does not need to be inside the control satellite housing, but may have its own housing fixedly mounted to the phase array (300), such as at the center of the phase array (300), or may be within one or more common satellite structures (10).

[0064] Application

[0065] The orientation of a spacecraft (especially large-scale) is frequently controlled using a Reaction Wheel System (RWS) and / or a Control Moment Gyroscope (CMG), but it can also be controlled by electric propulsion, chemical propulsion, or electromagnetic torquers. While an RWS can be used as an actuator for orientation control, it will become saturated if commanded to provide a constant torque indefinitely. Saturation is the state in which the actuator has generated as much torque as possible in a given orientation; beyond this point, it loses the ability to provide torque in that direction. CMGs suffer from the same saturation-based limitations as RWS, and the propulsion system is limited by a finite fuel source. Electromagnetic torquers are not inherently limited, but they require large devices to generate significant torque and cannot apply torque except where the local direction of the Earth's magnetic field permits.

[0066] Large structures in LEO require cumbersome propulsion or electromagnetic torque systems to balance constant gravitational gradient torques, or may use gyroscope stabilization as described above. In the case of GEO satellites, the burden imposed by gravitational gradient torques is significantly reduced compared to LEO satellites due to the difference in base radii between the two orbital regimes (200×). For this reason, GEO satellites may seek alternative means (e.g., propulsion systems) to allow for the constant application of torque to balance gravitational gradient torques. However, these alternative means are not suitable for use in LEO because they cannot provide sufficient force to counteract the increased gravitational gradient torques affecting structures in LEO.

[0067] In one embodiment of the present disclosure, a set of actuators and a set of sensors exist on a control satellite (200) (Fig. 1a). As illustrated in Fig. 7, the control satellite (200) includes a processing device (220), a flight computer (224), one or more sensors (226), and one or more actuators (228). The control actuators (228) may include, for example, reaction wheels, momentum wheels, and electromagnetic torque rods. The control sensors (226) may include, for example, star trackers, solar sensors, inertial measurement units, and global positioning system units.

[0068] In the satellite structures (10) (Figs. 5 and 6), one or more structure actuators (28) include, for example, electromagnetic torque rods, and one or more structure sensors (26) include, for example, inertial measurement units and global positioning system units.

[0069] The total integrated control authority of the control satellite (200) and the small satellites (10) is used to control the orientation and rotational speeds of the entire structure (100). The momentum wheel of the control satellite (200) rotates around the normal vector of the structure (100) at a specific speed that brings the structure (100) into equilibrium in combination with the orbital rotation of the structure (100). Gyroscope torque is generated by the momentum state of the momentum wheel and the orbital rotation. This generated gyroscope torque must offset the net torque from the gravity gradient and the natural gyroscope torque of the structure.

[0070]

[0071] ω is the rotation of the structure, and I wheel is the moment of inertia of the wheel (how large the actuator is), and ω wheel is the rotational speed of the wheel (maintained at a specific rotational speed in our case), and L wheel is the angular momentum of the wheel (moment of inertia multiplied by the wheel's rotational speed).

[0072] The natural gyroscope torques of a structure are defined by Euler's rigid body formulas.

[0073]

[0074]

[0075] Gravity gradient torque was previously defined.

[0076] The reaction wheels and torque rods of 200 are used for fine control of the structure. Small deviations from the desired orientation are corrected using these systems. A set of sensors of 200 (star trackers, solar sensors, IMUs (accelerometers), and GPS) is used to determine the orientation and position of the structure. These sensors determine whether there are deviations from the desired orientation of the structure (100). A set of sensors of the satellites (10) determines the shape and orientation of the phase array for the control satellite (200).

[0077] In one embodiment, each satellite (10) has two sensors (one IMU and one GPS) and 0 to 2 torque rods. The sensors (26) and actuators (18) can be distributed to best correspond to the vibration modes of the structure based on mode analysis.

[0078] The flight computer (FC) (224) serves to control the orientation of the entire structure. It can directly receive all sensor data from the sensors (226) and command the actuators (228). Additionally, (224) can receive processed sensor data from each common satellite (10) and command the actuators (28) indirectly (the FC (224) transmits commands to the FC (24), and the FC transmits commands to the actuators (28)). In this way, each actuator (28) becomes an extension of the actuators (228). The FC (224) can determine whether the orientation of the structure is deviated from the desired orientation and command any or all actuators to correct the orientation.

[0079] The wheels (reaction and momentum) within the actuator set (228) control the structure by changing the rotational speed. Ideally, the wheels are maintained at small rotational speeds to minimize drag and improve controllability, but they can rotate at high speeds to maintain a large amount of momentum, thereby generating gyroscope torques (the principle used by the gyroscope component (229)). Wheels designed for this purpose are appropriately called momentum wheels because they maintain a high momentum state by rotating rapidly. Since the 200 is firmly attached to the structure (100), the rotational momentum wheel, in combination with orbital movement, generates torque within the 200 (due to the rigid connection) that corresponds to the gravity gradient on the entire structure (100). A flight computer (224) is required to determine the desired angle (θ) to be maintained, determine the rotational speed required for that angle (given gyroscope and gravity gradient torques), and then command the momentum wheel(s) acting as the gyroscope component (229) to rotate at that speed.

[0080] Therefore, the passive objective is to respond to the gravity gradient with a constant gyroscope torque. It rotates the satellite or the momentum wheel and maintains that rotational state until the structural angle (θ) is changed. The active objective is to maintain the desired orientation, which involves constantly detecting and adjusting the orientation provided by the actuator.

[0081] FIG. 8 is a flowchart illustrating a method ("active" operation) for maintaining orientation (50). Data from all sensors is supplied to FC (224) (step 52), and FC determines whether the orientation is out of alignment (step 53). If it is out of alignment (step 53), commands are generated for any or all actuators (step 54), and actuators (228) execute these commands (step 56); FC (24) relays the commands so that actuators (28) can execute them (step 57).

[0082] FIG. 9 is a flowchart illustrating how gyroscope torque (70) is generated. FC (224) determines the desired angle (θ) (based on the beta angle and input commands transmitted from ground control) (step 72), and determines the gravity gradient and the gyroscope torque required therefrom (step 74). In step (76), the actuator is then commanded to generate rotational movement (the entire structure or momentum wheel(s)) in the correct direction (step 78).

[0083] The remaining actuators and sensors are used to ensure the accurate orientation of the structure in the gyroscope-stabilized orientation shown in FIG. 3d. Based on information measured by control sensors (226, 26) on the control satellite (200) and / or small satellite (10), flight computers (224, 24) may provide commands to the relevant actuators (228, 28) given the currently measured orientation of the structure (100). These commands are executed by the actuators (228, 28) to generate realized torques on the structure (100) that can guide the structure (100) from its current orientation to a desired orientation. This situation allows the LEO satellite to permanently maintain the gyroscope-stabilized orientation of the entire structure (100).

[0084] In the illustrated embodiments, the small satellite (10) and / or central satellite (200) may include a processing device for performing various functions and operations according to the system and method of the present disclosure; for example, including processing devices (20, 220) and flight computers (24, 224). The processing device may be, for example, a computing device, a processor, an application-specific integrated circuit (ASIC), or a controller. The processing device may be provided with one or more of various components or subsystems, such as wired or wireless communication links, and / or storage device(s), such as analog or digital memory or a database. All or part of the systems, processes, and / or data utilized in the system and method may be stored in or read from the storage device. The processing device may execute software that may be stored in the storage device. Unless otherwise indicated, it is preferable that the process be automatically implemented by the processor in substantial real-time without delay.

[0085] Therefore, the system evaluates the detected data, generates a flight pattern, and operates an actuator to implement the flight pattern. There are three basic orientations for positioning a flat phase antenna array (300). The first two, illustrated in FIG. 3a and FIG. 3b, are stable but are not good options because they orient the bottom surface (antenna) 90° away from the Earth. The third method is unstable as illustrated in FIG. 3c and is the basic position for conventional bottom flight satellites such as the phase array (300).

[0086] The present disclosure maintains a phase array (300) in the orientation shown in FIGS. 3d and FIGS. 4b. This orientation can increase the power generated on the surface of the solar layer (14) in all cases except when β = 0°. As long as the offset angle (θ) is relatively small, it is acceptable from an RF perspective. That is, it provides stable communication with the Earth from the antenna in the communication layer (12) of the phase array. Since any offset has a negative effect on RF, the angle selection is a trade-off between power and RF.

[0087] However, since this orientation of FIG. 3d is not inherently in an equilibrium state, it needs to be stabilized. By inducing it into an equilibrium state, it can be maintained permanently. However, since gravity gradients and natural gyroscope torques depend on the moment of inertia and orbital altitude, it can be difficult to induce it into an equilibrium state. The structure (300) is large and flat, which means it has a significant moment of inertia. And it operates in an LEO orbit with high gravity gradients and gyroscope torques (thanks to the fast orbital period). Therefore, gyroscope torques are generated to keep the phase array (300) in an equilibrium state. In particular, the entire phase array structure (300) is rotated about its center (i.e., about the normal vector, for example, like a pizza). The flight computer (224) sends a command to the flight actuator (228) to rotate the entire structure to perform a rotation / spin. This is achieved over several hours or days on the magnetic torque rods (in the actuator (228)).

[0088] In another embodiment, an internal gyroscope component (229) is provided such that the momentum wheel (e.g., a cylindrical symmetric component such as a flat disk located inside the control satellite (200) and controlled by the flight computer (224) or processor (220)) rotates around a normal vector.

[0089] Gyroscope torques are used because they generate a constant torque. When rotational motion is generated, gyroscope torque is continuously generated to keep the array (300) in a stabilized state with minimal power. This is implemented by observing the angle of the sun (β) and performing a system-specific trade-off between RF and solar power to determine which offset angle (θ) should be used.

[0090] In another embodiment of the present disclosure, data (such as position and attitude) may be transmitted from a satellite (10 and / or 200) to a ground station (e.g., by processing devices (20 and / or 220)). The processing device of the ground station may determine the necessary gyroscope torque and / or other flight information and transmit control signals to the satellites (10 and / or 200) (e.g., processing devices (20 and / or 220)) via the ground station antenna to control the flight pattern (e.g., using actuators (28 and / or 228)).

[0091] In one example, the system of the present disclosure may be used in an array of antenna assemblies, such as forming an array (300) of small satellites (10). When the structure is configured as an antenna array, the array (e.g., antennas or antenna elements) communicates with processing devices on Earth, such as user devices (e.g., mobile phones, tablets, computers) and / or ground stations. The present disclosure also includes a method of using antennas to communicate with processing devices on Earth. The present disclosure also includes a method of processing devices on Earth that communicate with antennas. Additionally, antennas may be used in Low Earth Orbit (LEO) or other orbits, or for other applications. Furthermore, although the system and method have been described for an array of antenna assemblies, the system and method may be used for other applications, such as data centers, reflectors, and other structures implemented in space or on Earth.

[0092] It should be further noted that the description and claims use various geometric or relational terms, such as plane, top, bottom, side, parallel, normal, surface, and flat, as well as various directional or positional terms. These terms are used merely for convenience to facilitate description based on the embodiments illustrated in the drawings. These terms are not intended to limit the present disclosure. Therefore, it should be recognized that the system may be described in other ways without geometric, relational, directional, or positional terms. Furthermore, geometric or relational terms may not be precise. For example, walls may not be exactly perpendicular or parallel to each other, but are still considered to be substantially perpendicular or parallel. And, other suitable geometric structures and relationships may be provided without departing from the spirit and scope of the present disclosure.

[0093] The foregoing description and drawings should be regarded as merely illustrative of the principles of the present disclosure, which may be configured in various ways and are not intended to be limited by the embodiments described herein. Many applications of the present disclosure will readily occur to those skilled in the art. Accordingly, it is not desirable to limit the present disclosure to the specific examples disclosed or to the exact configurations and operations illustrated and described. Rather, all appropriate modifications and equivalents may be considered to fall within the scope of the present disclosure.

Claims

Claim 1 An antenna system configured for use in a low Earth orbit (LEO), comprising a plurality of antenna assemblies joined together to form a phase array, wherein a first surface of the phase array comprises a plurality of solar cells, a second surface of the phase array comprises a plurality of antennas, and the second surface is different from the first surface; and, comprising one or more processors, wherein the processor determines a desired orientation of the phase array based on the position of the first surface of the phase array relative to the sun or the position of the second surface of the phase array relative to the Earth, wherein the desired orientation comprises an orientation of the phase array such that a gyroscope torque acting on the phase array is offset by a gravitational gradient torque acting on the phase array, and controls one or more actuators to apply a gyroscope torque to the phase array to position the phase array in the desired orientation. Claim 2 delete Claim 3 An antenna system according to claim 1, wherein the one or more actuators comprise at least one gyroscope component configured to generate rotational movement. Claim 4 In claim 3, the antenna system wherein the gyroscope component includes a momentum wheel. Claim 5 delete Claim 6 An antenna system according to claim 3, wherein the gyroscope component is configured to generate rotational movement to be stabilized against gravity gradient torque using gyroscope force. Claim 7 delete Claim 8 delete Claim 9 delete Claim 10 An antenna system according to claim 1, wherein the one or more actuators comprise a torque rod. Claim 11 delete Claim 12 An antenna system according to claim 1, further comprising a control satellite communicably coupled to the plurality of antenna assemblies, wherein the control satellite comprises the one or more processors. Claim 13 delete Claim 14 delete Claim 15 An antenna system according to claim 3, further comprising a housing fixedly mounted to the plurality of antenna assemblies, wherein the gyroscope component is rotatably mounted to the housing. Claim 16 delete Claim 17 An antenna system according to claim 1, wherein the desired orientation is offset by a controlled angle from the stable orientation. Claim 18 An antenna system according to claim 1, wherein the desired orientation is unstable and one or more actuators are configured to maintain the phase array in the desired orientation. Claim 19 An antenna system according to claim 1, wherein the desired orientation positions the phase array at an angle offset from orthogonality to the surface of the Earth. Claim 20 delete Claim 21 delete Claim 22 delete Claim 23 delete Claim 24 A method performed by an antenna system of any one of paragraphs 1, 3-4, 6, 10, 12, 15, 17-19.

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