Patent ID: 12261678

DETAILED DESCRIPTION

In describing the illustrative, non-limiting embodiments of the disclosure illustrated in the drawings, specific terminology will be resorted to for the sake of clarity. However, the disclosure is not intended to be limited to the specific terms so selected, and it is to be understood that each specific term includes all technical equivalents that operate in similar manner to accomplish a similar purpose. Several embodiments of the disclosure are described for illustrative purposes, it being understood that the disclosure may be embodied in other forms not specifically shown in the drawings.

Referring toFIG.2, in one example embodiment, only the electronics of a single small satellite302is shown. In this embodiment the structure10is an assembly or common small satellite that is connected to other antenna assemblies in a large antenna array5(FIG.3), such as in the antenna assembly300and array100of the '266 Patent (FIG.1(a)). The overall system forms an Altitude and Orbit Control System (AOCS) that can include the common satellite10, control satellite200, and/or ground station. The control satellite200can be fixedly connected to the small satellites302, such as at the center of the array as in5(FIG.3(a)).

The structure10is flat and rectangular or square, with the communication components (e.g., antenna elements19) at one side surface7(FIG.3(a)) facing the Earth (nadir) to communicate with user devices (e.g., cell phones) and an opposite side surface9(FIG.3(a)) facing in the opposition direction (zenith) with solar cells17that generate solar power for use by the electronic components, e.g., a processing device12, antennas19, battery16, and antenna front end modules18.

Each structure10also has one or more connectors14, such as a hinge, joint, spring or tape-spring connector, that connect the structure10to one or more neighboring similar structures10. As shown, the structure10can be rectangular or square and encompass multiple antenna elements300, and one or more connectors can be positioned at or along one or more of the edges or sides of the structure10. It will be recognized that the system can utilize any suitable connection system, for example such as the one shown and described in U.S. Patent Pub. Nos. 2020/0361635, 2020/0366237, and 2020/0365966, the entire contents of which are hereby incorporated by reference.

The connectors can be subject to bending or flexing in the operating configuration. For example, the maximum flex at the connector14might be several degrees. Any flex results in a deviation of the antenna elements from the planar configuration in which the communication side (and/or the solar side) of the plurality of antenna elements are planar. That deviation is undesirable since it can affect beam formation.

In this example embodiment, it is therefore desirable for the array5to be substantially flat on both the solar side and the communication side, i.e., that the individual structures10are flat on both sides and that they are planar or co-planar with one another on both sides so that the overall array5is planar on both sides. However, the structure10and/or array5is subject to forces in space that can cause the structure10or array5to flex or bend.

FIGS.3(a), (b)show a nearly circular large planar phased array5formed in space by integrating many small satellite structures10. Each small satellite could host a processing device12(e.g., processor) and several antenna elements19. Hundreds of such small satellites together could form a large phased array with thousands of overall antenna elements. In one embodiment, each small satellite (referred to here as a micron) has an antenna assembly with antenna elements19arranged in four rows and four columns in a square shape. The overall phased array formed by the interconnection of several small satellites could take a square, or a rectangular or a circularized shape as desired by the application. Any suitable small satellite can be utilized, such as shown and describe in U.S. Pub. Nos. 2020-0361635, 2020-0366237, and 2020-0365966, the entire contents of which are hereby incorporated by reference.

The antenna elements19are positioned at the communication side surface7of the array and the solar cells are positioned at the solar side9of the array5. The arrow11shows the boresight, which for a planar phased array refers to a normal to the array's plane. Any beam off-boresight is called an edge beam (e.g.,FIG.6)

The small satellites10communicate with end user devices (such as cell phones) on Earth, and with the central processor200, which in turn communicates with a gateway at the ground station. The signals communicated to/by the small satellites are aggregated together, such that the small satellites collectively transmit and receive signals to the end user devices. However, any bending or flexing of the array can cause the signals from the individual small satellites to deviate or be out of phase from the desired phase.

FIGS.4(a),4(b),5(a),5(b)show the gradual bending effect in large arrays from the center towards the edge of the arrays. The individual antenna assemblies are rigid, but are mechanically coupled to one another by the connectors14. Those connectors hold the small satellites together, but are subject to bending or flexing, and tend to oscillate inwards and outwards at low frequency, with maximum displacements at the extremities, depending on the external forces. As the mass per unit aperture is reduced, the stiffness of the array is reduced and the array encounters greater flexure. The arrows inFIGS.4(a),5(a)point to the boresight of the array.

FIGS.4(a),4(b)illustrate the inward (towards the arrow11or boresight) flexing of the array, andFIGS.5(a),5(b)illustrate the outward (away from the arrow11or boresight) flexing of the array. The array is imagined to be in a nominal plane that is normal or perpendicular to the boresight. The flexing causes deviation from the nominal plane.

The example ofFIGS.4,5, shows a single bend20in the array5. However, other bends are possible, for example bends that only partially extend along the array, bends that are offset from the center diameter, bends that extend at other positions and locations that do not pass through the center of the array, and multiple bends. And, while the bend20is shown having a sharp angle between the left and right halves22,24, the bend can be more curved. And, the left and right halves22,24need not be planar, but can be curved due to slight deviations or bends at connectors between antenna modules10.

Referring toFIG.6, the top row of measurements shows the expected radiation patterns from a 10.3 m planar array ofFIG.3while forming the beams towards the boresight of the array and edge of the footprint. The bottom row of measurements inFIG.6shows the expected radiation patterns for an array ofFIGS.4,5with 8.7□ bending from nominal plane without any compensation for change in antenna elements' position while beamforming. As shown, there is a distorted radiation pattern with dual main-lobe and the reduced array gain due to bending of the array.

WhileFIGS.4,5show uniform bending effect, there could be random perturbations in the position of each small satellite structure10while deploying and attaching to neighboring small satellites.FIG.7(a)shows a nominal orientation for a structure10, andFIGS.7(b)-7(c),8(a)-8(b)show such perturbations in the form of in-plane and out-plane tilts and displacements for a plurality of antenna elements.FIG.7(a)and the top drawings ofFIGS.7(b),7(c), each show a single structure10(which may have multiple antenna elements).FIGS.8(a),8(b), and the bottom drawings inFIGS.7(b),7(c)show multiple structures10coupled to one another by connectors, and bending at each connector. General displacements such as those inFIG.8are not controllable as described in the current embodiment, but such displacements are observable using the determination methods described herein if they are sufficiently large. Thus,FIGS.4,5,7,8depict the main types of perturbations that are likely to occur, any of which may increase the distortion of the radiation pattern. The embodiment described herein has significant utility for near-planar arrays but may be utilized in systems with larger non-planarity if a majority of such non-planarity is the result of modal perturbations.

FIG.9is one embodiment of the disclosure in which the structure10inFIG.7has a single Global Positioning System (GPS) unit that reports carrier phase data. A collection of such antenna elements collectively contain a set of GPS units50whose data is used to calculated a Carrier Phase Differential GPS (CD-GPS) array51of differential GPS solution212. Each structure10also includes an Inertial Measurement Unit49, and the collective IMU data is compiled in an array52of filtered IMU measurements515. Such IMUs may be inertial sensing units which integrate acceleration only and/or directly detect angular motion, or other such inertial measurement devices.

The embodiment inFIG.9also includes a structure estimator54a computed torque model62based on that structure estimator, a torque command66issued to a torque mechanism, and a feedback element64from that torque mechanism. In one embodiment, the GPS sensors50are positioned as close as possible to the IMU sensors49. The GPS sensors50and/or IMU sensors49are coupled to the structure elements10; each structure10can have one or more GPS sensor50and one or more IMU sensor49, or one GPS sensor50and/or one IMU sensor49can be associated with multiple structures10. In one embodiment, the IMU sensors49are accelerometers that detect an acceleration.

The effect of modal array deformation caused by coupled flexure on beamforming can be minimized by considering the instantaneous position of each antenna element while computing the corresponding phases used for beamforming. This is accomplished by placing position sensors50(FIGS.2,7(a)) and IMUs49in several of the small satellites10for accurate position and rate estimation of each antenna element or minimum required sensors, to predict the uniform perturbation characteristic of the small satellites across array, to determine approximate position of each small satellite. For example, each small satellite10can have one or more sensors that are placed on the body of the structure, such as on the communication side (the side with the antennas19) or the solar side (the side with the solar cells17). The sensor placement is determined by the type of sensor to ensure observability of the sensed effect. In the current embodiment, the GPS antenna attached to the receiver50must face the GPS constellation and therefore must be on the solar side of the antenna element structure10, while the IMU may be placed near the center of the structure10.

As shown inFIG.11, each pair of GPS receivers must observe overlapping signals from the GPS constellation. The sensors can, for example, be part of the electronic circuits that form the small satellite. Reference X, Y and Z planes for a nominal array are shown inFIG.7(a). The sensors can be a standard Global Positioning System (GPS) receivers or other sensor devices that automatically estimate position in a global co-ordinate system to the accuracy of carrier phase differential GPS (approx. 2 cm relative accuracy).

The sensor50in this embodiment is a standard Global Positioning System (GPS) sensor device that automatically estimates position in a global co-ordinate system and provides carrier phase and pseudorange output data for individual received GPS signals. As shown inFIG.11, the GPS receivers50reside at different locations298and299on the satellite202. Each receiver senses its position160relative to common GPS satellites (exemplified here as135,137), then use carrier phase information152as it changes154over time150to determine relative location170of the receivers. A group of such sensors50with mutually common received signals results in a set of solutions51providing relative offsets170between the receivers50. This process is known as CD-GPS and is well-documented in academic and technical papers, though any suitable technique can be utilized. In order to perform the system operations described herein, the number of CD-GPS solutions51need only exceed some minimum requirement. There is no theoretical maximum to the number that may be used, as additional solutions tend to improve the solution.

The embodiment here further utilizes inertial measurement unit (IMU) sensors49that provide measurements of the motion of a subset of satellite elements52. The raw IMU data is filtered84(FIG.9) with e.g. a low-pass filter to produce a dataset with content at a higher rate than the structure estimator operation and with low-quality or outlying datapoints removed. Such IMU data is not required to coincide identically in time with the available CD-GPS data, but the sensors must be placed to capture oscillatory motion within the system. As with the GPS solutions, the number of IMU datapoints need only exceed some minimum requirement for system observability, but no theoretical maximum number exists.

Timing misalignments between CD-GPS and IMU solutions are resolved via propagation of the existing solution. These CD-GPS and IMU solutions are the inputs to a structure filter83that estimates the characteristic parameters describing the displacements of the spacecraft structural elements, as well as the persistent error (bias estimate) in the IMU rate measurements85. These bias estimates are applied to the available rate data solution60even if new CD-GPS data is not available. During these periods, the structure motion is propagated87solely using IMU data to provide an estimate of the current location of each spacecraft element90. Structure constants may be updated at each timestep88to reflect the most recent observations, or if no control is performed they may be updated only when CD-GPS is available. Note that the solution also requires as an input the baseline angular motion of the spacecraft system89. Once structural parameters are estimated, the positions X, Y, and Z of the structure elements can be determined as often as required to provide data to beamforming. Instantaneous solution accuracy is sub-centimeter and can be extrapolated to similar accuracy over very short timespans.

The placements of the sensors49and50in the example realization are chosen to coincide with the maximum displacements due to the primary oscillation modes of the aperture structure. Because the beamforming phase is used to determine the phase compensation based on the displacement of the structure elements10from a planar configuration, the system must compensate for any additional flexing by mechanically correcting the relative positioning of the spacecraft elements302where possible. Where each small satellite10has its own sensors49and50, the deviation can be determined based on the position of that sensor. However, where a single sensor is provided for multiple small satellites, the deviation and correction for each small satellite10can be interpolated based on the positions of surrounding small satellites with respect to the sensor. In the embodiment shown here, these displacements are determined at the central satellite200(FIG.1(a)) by calculating the apparent modal excitation88. Determining the displacements resulting from these modes offers information that can be used to counteract their effects through a variety of means, and directly controlling these modes can improve spacecraft performance.

The present system resolves that measurement problem with a real-time estimation algorithm and integrates the solution with space-capable actuators to perform closed-loop control. The displacement filter56outputs the displacement data to the correction and control module58, which uses an actuator23operating on an external field. In one embodiment, the actuator23can be an electromagnetic torque rod that applies a physical displacement to the structure10and correct the overall displacement of the array5, reducing deflections about the bending axis20. The correction and control module58uses current and extrapolated deformation constant estimates and determines apparent excitation energy. It then applies the correction to the actuator23by a damping torque model62, which sends a torque command66to the torque rod.

The torque rod then applies a torque to the small satellite10to move it back toward the desired position at the point of maximum deflection, and the resultant torque is fed back to the model64. In this way the peak displacement decays. The control law associated with the estimated CD-GPS/IMU mode dynamics can take any basic nonlinear form, or a linear form if the structure is sufficiently rigid. Because mode motion is oscillatory as a function of the estimated mode shape constants88, and because digital compensation for some limited motion is possible, the selected control law need only force the system to remain within an allowable equilibrium displacements.

Thus, the processing device at the central controller200computes the commanded actuator input to move one or more of the plurality of antenna element structures10to correct for structural displacement of the antenna element structures. This computation may include prior known actuator command64, which can affect system dynamics and required torque inputs going forward. The correction and control module58also outputs the displacement data (e.g., the sub-centimeter displacement) to the digital beamformer68. The digital beamformer68computes and applies a phase correction to the beam. Any suitable technique can be utilized to apply an electronic correction, such as described in co-pending application no., which claims priority to 62/976,107, the entire contents of which are hereby incorporated by reference.

As a more specific example of the embodiment described herein, for systems with a significant first bending mode (seeFIGS.3-4), bending of the longest beam of the system about a perpendicular eigenaxis can a locally linearized as a 1-dimensional Euler-Bernoulli beam. Bending is approximately quadratic. The controller is designed only to stabilize the principal modes, so a constraint is placed on the energy imparted to the system (|u|<|umax|) to prevent possible excitation of additional modes. As noted above, control is imparted symmetrically about the inertial eigenaxes. Deformation energy is driven to 0, with a hysteresis enacted to limit chatter as the allowed maximum deformation is achieved.

Thus, as described above, the present system determines the amount of local movement of the structural array30, and then corrects that. It is further noted that co-pending application no., which claims priority to 62/976,107, determines the amount of flexing or bending of the structural array5, and then corrects that by performing beam forming techniques that compensate for the bending. The entire content of that application is incorporated herewith.

In one example embodiment, structure10is an antenna assembly with a solar panel that receives solar energy from the Sun and generates solar power for use by the structure. The overall structure is flat and rectangular or square, such as a tile, with the communication components (e.g., antenna elements) at one side surface facing the Earth (nadir) to communicate with user devices (e.g., cell phones) and an opposite side surface facing in the opposition direction (zenith) with solar cells that generate solar power for use by the electronic components—e.g., a processing device, antennas, antenna front end modules. Here the control satellite200is fixedly connected to the small satellites302, shown at the center of the array100and visible in the array5.

It is important that the structure10and the array5remain as flat (i.e., planar) as possible to maximize solar power generation by the solar side and communication with the Earth on the communication side. Thus, it is desirable for the array5to be substantially flat, i.e., that the structures10are flat and that they are planar with one another. However, the structure10and/or array5is subject to forces in space that can cause the structure10or array5to flex or bend. To correct for any bending or flexing, the structure10has symmetrically-placed GPS units50, inertial measurement units49, and actuators23on each small spacecraft.

The processing device12in the common satellite10transmits required data to the control satellite200in real time, which derives the initial estimates of small displacements due to modes of the structure flex11via the CD-GPS solution51and IMU data52at time t=0. The control satellite200then determines the apparent effect of the bending of the structure11and estimates the modal contributions of the expected principal modes via structure constants. Approximately 1/10 second later, additional filtered IMU data52is received, and calculated bias and the spacecraft angular rate89are removed86from the sensed rate. The system is propagated forward87. Structure constants may be updated88if acceleration or rate data suggests a large error in the estimate, resulting in a final displacement estimate90and58.

This data is used as an input to the damping torque computation62which, along with any previous output torque64, results in a new desired torque command66. These torque commands are sent to the appropriate small satellite actuators23as torque rod activation signals. Finally, the estimate90and58is sent to the beamformer68, which uses the data to improve its digital beamforming solution. This process repeats for 0.25 seconds until new CD-GPS data is available, at which point the estimated structure constants are updated using the position data.

When the structure10is configured as an antenna array5, it (e.g., antenna19or antenna elements302) communicates with processing devices on Earth, such as for example a wireless device including a user device (e.g., cell phone, tablet, computer) and/or a ground station. The present disclosure also includes the method of utilizing the structure10to communicate with processing devices on Earth (i.e., transmit and/or receive signals to and/or from). The present disclosure also includes the method of processing devices on Earth communicating with the structure10(i.e., transmit and/or receive signals to and/or from). In addition, while the structure10is used in Low Earth Orbit (LEO) in the examples disclosed, it can be utilized in other orbits or for other applications.

Still further, while the system has been described as for an array of antenna assemblies, the system can be utilized for other applications, such as for example data centers, telescopes, reflectors, and other structures, both implemented in space or terrestrially. The system of the present disclosure can also be utilized in combination with a phase correction system, such as shown and described in U.S. Application No., filed herewith, entitled Compensating Oscillations in a Large-Aperture Phased Array Antenna, claiming priority to U.S. Application No. 62/976,107, filed Feb. 13, 2020, the entire contents of which are hereby incorporated by reference.

In addition, it is noted that operation is described as occurring at the control satellite200, which may or may not be fixedly embedded in the array. However, operation can also be at the common satellite10processing device12if GPS and IMU data from other structures10is distributed in such fashion. In another embodiment of the present disclosure, data (such as position and attitude) can be transmitted from the satellite10and/or200(e.g., by the common satellite processing device12and/or the control satellite processing device, if such are not coincident) to a ground station. The ground station processing device can then determine the necessary correction and/or other flight information and transmit a control signal to the satellite10and/or200(e.g., common satellite processing devices12and/or control satellite processing device) to control the correction via the torque rod23, in addition to performing other ground-based tasks.

It is further noted that the drawings may illustrate and the description and claims may use several geometric or relational terms and directional or positioning terms, such as planar, linear, curved, circular, flat, left, and right. Those terms are merely for convenience to facilitate the description based on the embodiments shown in the figures, and are not intended to limit the disclosure. Thus, it should be recognized that the system can be described in other ways without those geometric, relational, directional or positioning terms. In addition, the geometric or relational terms may not be exact. For instance, walls or surfaces may not be exactly flat, or planar to one another but still be considered to be substantially planar because of, for example, roughness of surfaces, tolerances allowed in manufacturing, etc. And, other suitable geometries and relationships can be provided without departing from the spirit and scope of the disclosure.

The foregoing description and drawings should be considered as illustrative only of the principles of the disclosure. The system may be configured in a variety of shapes and sizes and is not intended to be limited by the embodiment. Numerous applications of the system will readily occur to those skilled in the art. Therefore, it is not desired to limit the disclosure to the specific examples disclosed or the exact construction and operation shown and described. Rather, all suitable modifications and equivalents may be resorted to, falling within the scope of the disclosure.