TWI901982B - Satellite for operation in orbit around earth, method for operating the satellite, and satellite sar phased array antenna for imaging the earth's surface - Google Patents

Abstract

A satellite radar antenna array is formed as a generally planar structure comprising a plurality of panels. Each panel comprises a plurality of antenna transmit/receive modules; each antenna transmit/receive module comprises a stack of planar elements comprising an RF board and an antenna board, and one or more heat sink components for conducting heat from the RF board; and each RF board comprises a plurality of RF front ends. The stack may additionally comprise a power board.

Description

Methods for operating satellites orbiting the Earth, and satellite synthetic aperture radar phase array antennas for imaging the Earth's surface.

This invention relates to the field of satellite design and operation.

Synthetic aperture radar (SAR) is one of the known methods for observing Earth from satellites in space. Traditionally, these satellites are large, expensive, and limited in number, with only a handful in space. Due to the complexity and cost involved, SAR satellites are typically built by governments or government-backed space agencies.

The emergence of companies like SpaceX demonstrates that commercial organizations are increasingly involved in the space sector, introducing new and innovative technologies to reduce the cost of space missions and making space more accessible to both governments and businesses. This shift is largely driven by private capital rather than government funding, and commercial players in this field are collectively referred to as part of the "new space" industry, rather than the "old space" industry. This shift is making space more accessible to commercial organizations, for example, by significantly reducing launch costs, thus stimulating innovation and new technologies in the field.

The challenge for satellite designers includes how to build a large constellation of small satellites for Earth observation that can deliver data at a fraction of the cost of traditional SAR Earth imagery (and revisit it more frequently than traditional SAR satellites).

Satellite design must be inexpensive and easy to build, and its imaging capabilities must meet the resolution and accuracy requirements of typical space missions. Furthermore, they need to be small enough to take advantage of "new space" commercial launch opportunities to minimize launch costs, allowing small satellites to be launched at a fraction of the cost of traditional satellite launches, for example, through SpaceX's "rideshare" launches, which can send multiple aircraft from different organizations (including governments and commercial entities) into space as part of a single mission.

Smaller satellite sizes require a SAR design that is much smaller than traditional satellites and addresses some of the problems associated with small SAR systems, such as achieving acceptable image quality with a smaller SAR aperture.

The satellite design and construction methods described below address some problems. However, the present invention is not limited to solutions to these problems, and some embodiments of the present invention can solve other problems.

This invention provides a simplified introduction to a selection of concepts, which are further described in detail below. This invention is not intended to identify key or fundamental features of the claimed subject matter, nor is it intended to determine the scope of the claimed subject matter.

In a first embodiment, a satellite for operation in orbit around the Earth is provided below, comprising: a propulsion system, an attitude determination and control system (ADCS), and a radar antenna array; wherein the radar antenna array is formed in a generally flat structure, including a plurality of panels, each panel including a plurality of antenna transmission/reception modules; each antenna transmission/reception module includes a flat component stack comprising an RF board and an antenna board, and one or more heat dissipation components for heat conduction from the RF board; and each RF board includes a plurality of RF front ends.

As will be explained below, this configuration of the RF front end (where power components are positioned close to the amplifiers they supply) increases power density, which in turn leads to increased image resolution achievable with a small satellite.

In the technical field, the term "board" refers to a plate that supports one or more components, such as antenna elements, amplifiers, or power supply components. In other words, the components are referred to as being included in the board, such that the term "board" refers not only to, for example, a substrate, but also to the components that support it.

The numerous advantages described herein are due to the specific architecture of the antenna array, which can be supplied separately for use with any suitable propulsion and ADCS, and therefore, an antenna array for use with one of the satellites described herein is also provided herein.

Therefore, in another embodiment, a satellite radar antenna array is provided, which is formed as a generally flat structure including a plurality of panels, wherein: each panel includes a plurality of antenna transmission/reception modules; each antenna transmission/reception module includes a flat component stack comprising an RF board and an antenna board, and one or more heat dissipation components for heat conduction from the RF board; and each RF board supports a plurality of RF front ends.

The following additional features may be provided in one of the satellite radar antenna arrays or a satellite as described herein: Each RF board includes between 8 and 32 RF front-ends, for example, 16 front-ends. Each front-end may include a power amplifier, a transmit/receive switch, and a low-noise amplifier.

The satellite or antenna array can be configured to operate as a phase array. Each RF front end may include a number of phase shifters in both the transmission and reception directions.

This antenna board can support an array of antenna components, which may include surface-mount antenna components.

Each RF front-end can be configured to drive multiple antenna components.

The area of the radar antenna array can be between 1 ^m² and 5 ^m² , depending on the situation, between 1 ^m² and 3 ^m² .

The radar antenna array can have an aspect ratio between 2 and 24.

The radar antenna array may include a row of panels, with adjacent panels pivoting relative to each other, thereby allowing the array to be folded for transport and deployment in space.

A solar panel array can be supplied to the radar antenna array according to a fixed relationship.

The solar panels can be configured in a generally flat arrangement extending in one of two opposite directions parallel to the extension direction of the radar array, wherein the plane of the solar panels is perpendicular to the plane of the radar array.

These solar panels can be configured such that adjacent panels can pivot relative to each other, allowing the solar panel array to be folded for transport and deployment in space.

A large-capacity memory module can be provided for storing data acquired by the radar antenna array, and a transmitter can be directly connected to the large-capacity memory module to download the data.

The flat component stack may further include a power distribution board. The one or more heat dissipation elements may include a heat dissipation plate in contact with the RF board. The heat dissipation plate may be disposed between the power distribution board and the RF board such that the heat dissipation pad, the power distribution board, and the adjacent main surfaces of the RF board are in contact for heat conduction. The one or more heat dissipation elements may include one or more structures configured to conduct heat from the RF board to the antenna board.

The radar antenna array can operate at wavelengths between 4 GHz and 12 GHz, or depending on the situation, between 8 GHz and 12 GHz, or between 4 GHz and 8 GHz.

The satellite and radar antenna arrays described below are intended to operate as a synthetic aperture radar, but may also have other applications.

The ADCS is configured to rotate the satellite around an axis parallel to its direction of travel along an orbit. Therefore, the entire satellite can be manipulated, for example, as part of an image acquisition operation. This can be achieved through additional electronic steering by operating the array as a phase array.

The following further provides a satellite for operation in orbit around the Earth, comprising: a propulsion system, an attitude determination and control system, a power supply for powering one or more solar panels, and an active phase array synthetic aperture radar "SAR"antenna; wherein: the SAR antenna has an area of less than 5 ^m² , and has a peak transmission power of greater than 1,000 W per square meter or an average transmission power of greater than 200 W per square meter during imaging.

This phased array SAR antenna can be configured to fold for transport and deployed in orbit in a single deployment.

The satellite could have a mass of less than 1000 kg, for example, less than 843 kg, less than 500 kg, or less than 200 kg.

For any satellite employing the aforementioned technical solutions, the peak transmission power during imaging may exceed 1000W per square meter or the average transmission power may exceed 200W per square meter.

The peak transmission power can be in the range of 1000W to 5000W per square meter, or the average transmission power can be in the range of 200W to 1000W per square meter.

The following also provides a method for operating a satellite to acquire image data according to any of the foregoing technical solutions, the method including limiting the time each orbit of a radar antenna array is turned on to avoid overheating of the radar antenna array.

For example, the acquisition of imagery data could be limited to 2% to 4% of the satellite's orbital duration, such as 3%.

As will be apparent to one skilled in the art, the features of the various forms and embodiments of the present invention can be suitably combined, and can be combined with any form of the present invention.

100: Satellite

110: Satellite Main Unit/Busboard/Aircraft Busbar/Satellite Main Unit/Satellite Busbar

115: Pillar

150: Solar Panels

155: Solar Panels

160: Radar Antenna Array

161: Panel

162: Panel

163: Panel

164: Panel

165: Panel

165a to 165d: Transmit/Receive (T/R) Modules

190: Attitude and Control System (ADCS)

195: Propeller

196: Radiator Panel/Panel

210: Antenna

400: Effective Load

401: Processing Board

402: Field Programmable Gate Array (FPGA)

403: Buffer Random Access Memory (RAM)

404: Flash High-Capacity Storage

421: Transceiver

450: Amplifier

461: Digital-to-Analog Converter (DAC)

462: Analog-to-Digital Converter (ADC)

510: Antenna Board

512: RF board

513: Heatsink/Radiator Plate

514:Distribution board

516: Support Structure

518: Support Structure

520: Power and Control Signal Connections

522: Power and Control Signal Connections

531 to 535: Radio Frequency (RF) Connectors

600: Frontend

601: Amplifier

603: Digital Phase Shifter

605: T/R switch

607: Amplifier/Low Noise Amplifier (LNA)

609: Digital Phase Shifter

620: Antenna Array

710: Antenna Patch

730: Antenna Subarray

800:RF board

801:Distribution board

802: RF Front End

803: Components

804: Components

805: Components

806: Components

807: Coaxial Connector

808: Components

809: Components

810: Components

1001: Power Supply

1002: Power Distribution System

1003: Computing System

1004:Communication system

1006: Antenna Block / Antenna

1007: Amplifier/Amplifier Block

1008: Power Distribution System

1009: Propulsion Controller

1020: Satellite Flow Chart

1030: Radar Antenna Array

1031: Attitude and Control System (ADCS)

1033: Synthetic Aperture Radar (SAR) Processor

1035: Buffer

1090: Propulsion System

1101: Electrical Power System

1102: Attitude and Orbit Determination Control System (ADCS/AOCS)

1103: Telemetry and Remote Control Systems

1104: Onboard Data Processing

1105: Effective Load Chain

1106: SAR Effective Load

1108: Main onboard computer

1109: Auxiliary Onboard Computer

1111: Surveillance Board

1115: Battery

1117: Power Regulation Module

1120: ADCS

1121: Propulsion Controller

1131: Main S-band transceiver

1132: Auxiliary S-band transceiver

1151: Main X-band Transmitter

1152: Auxiliary X-band Transmitter

1161: Signal Amplifier

Embodiments of the invention will be described by way of example only and with reference to the following figures, in which: Figure 1 is a perspective view of a satellite; Figure 2 is an alternative perspective view of the satellite shown in Figure 1; Figure 3 is a perspective view of the satellites of Figures 1 and 2 configured for transport; Figure 4 is a schematic diagram of a portion of an exemplary payload that can be carried by the satellites of Figures 2 and 3; Figure 5 is a schematic diagram of a transmission/reception module stack; and Figure 6 is an exemplary front end. Figure 7 is a schematic diagram of one of the related antenna patches; Figure 8 is an enlarged view of a portion of the radar antenna array shown in Figure 2; Figure 9 is a perspective view of an assembly including an RF board and a power distribution board that may be included in a transmission/reception module; Figure 10 is a magnified front view of a satellite in direction z from Figure 2; Figure 11 is a schematic diagram illustrating a feasible configuration of one of the satellite electronic components; Figure 12 shows a feasible configuration of one of the components shown in Figure 10. Figure 11 is a schematic diagram illustrating a feasible configuration of one of the components shown in Figure 10.

Common element symbols are used in all diagrams to indicate similar features.

The following description uses examples to illustrate implementations of the satellite design and operation. These examples represent the best way to put the invention into practice, and are currently known to the applicant, although they are not the only way to achieve this.

This invention relates to the design and operation of small satellites.

Currently, satellites in orbit around the Earth are typically defined by weight ranges, although the boundaries between these categories are somewhat unstable and arbitrary. These include cubic satellites, microsatellites, and others. One constraint distinguishing one type of satellite from another is its launch capability. Therefore, it is desirable for a satellite to weigh less than a certain maximum value so that it can be launched from a launch vehicle such as SpaceX's Falcon 9 rocket. For example, SpaceX offers rideshare launches, where multiple satellites can be launched as part of a single mission into different orbits. Compared to traditional custom missions, these rideshare programs offer a lower-cost alternative to sending satellites into orbit. Part of the rideshare program involves standardized payload configurations. For example, SpaceX offers an "exo-port" of 0.381m (15 inches) for a maximum mass of 454kg (1,000lb) or 0.610m (24 inches) for a maximum mass of 830kg in its rideshare program. Smaller satellites (e.g., with a maximum mass of 227kg or less) can also be accommodated. Sometimes, multiple satellites can be launched from a single exo-port, provided the total weight and volume of the rocket and fairing are within acceptable limits. The term "small satellite" as used here refers to the type of satellite currently available for commercial launches, and is generally, but not necessarily, limited to satellites weighing less than 1000kg.

Therefore, this invention relates to the design and operation of satellites having a weight of less than 830 kg, less than 454 kg, or less than 227 kg (to allow for the launch of two satellites from a single port).

Other launch providers are also feasible, such as the Vega rocket developed by Airanespace, the Italian Space Agency, and the European Space Agency. The Vega rideshare service utilizes small aircraft mission services and is designed to launch small payloads for scientific and Earth observation missions. The requirements and interface parameters of the various launch providers are similar, although they may not be identical.

Figures 1 and 2 are different perspective views of a satellite 100. The basic components of a satellite similar to those in Figures 1 and 2 are described in our earlier patent application WO2022079168A1. According to one example, a small satellite includes a propulsion system and an attitude determination and control system (ADCS), as known in the art. It should be noted that in the satellite design discussed herein, the ADCS can be controlled to manipulate the entire satellite to achieve desired positioning and orientation (e.g., viewing angle) during imaging operations.

Satellite 100 also features a propulsion system for controlling the satellite to generate thrust, described in more detail in WO2022079168A1.

The satellite includes a body or bus 110 that houses or supports components including the ADCS 190 and propulsion system, shown in Figure 1 as including four thrusters 195. The body 110 is also referred to in the art as a "bus" because it houses or supports bus components. The bus 110 may additionally house one or more batteries. The bus 110 may be partially enclosed, for example, to house and protect components. A housing provides a surface on which components can be mounted. In the satellites of Figures 1 and 2, a solar panel array is provided, comprising a solar panel 150 mounted on a rectangular surface of a bus 110 and additional solar panels 155 attached to the solar panel 150 by means of struts 115. The illustrated satellite has a total of five solar panels 150, 155 to form a solar array described further below. In the illustrated example, the solar array has five rectangular panels. The solar panels are arranged in a row, wherein adjacent panels can pivot relative to each other, thereby allowing the solar panel array to be folded for transport and then deployed in space. In the examples of Figures 1 and 2, panel 155 is deployable and panel 150 is mounted on bus 110. In one example, each panel contains 112 solar cells, which are positioned on the front side of the satellite (-Y plane, see Figure 2), connected to the satellite's batteries, and provide up to approximately 90W of power when directly facing the sun. The nominal average orbital power during operation can be approximately 350Wh. A radiator panel 196 (described further with reference to Figure 9) serves as both a radiator for thermal management within the satellite and, in this example, a radiator for attaching satellite 100 to a rocket during launch.

Other components included in a satellite can be provided to enable the satellite to perform a mission, and these are referred to as "payload".

In one example, the payload is an Earth observation system used to image the Earth's surface, called a synthetic aperture radar (SAR) system. This payload operates by transmitting radar signals from a SAR antenna to Earth and listening for the returning echoes. By recording the echoes, an image of the Earth's surface can be constructed from data including the time required for the echo to return (indicating position), the amplitude of the radar echo, and the phase information contained in the radar signal. Further position information is obtained from the frequency of the returned signal, which is attributable to the Doppler effect shift due to the satellite's motion relative to the Earth.

The field of synthetic aperture radar (SAR) is well-known in technology, and the first civilian SAR satellite was launched in 1978. However, SAR satellites have traditionally been quite large (>1000 kg), complex, and very expensive to design, build, and launch. They have traditionally required dedicated launches to reach space, making them prohibitively expensive for commercial, non-governmental applications. Implementing SAR in smaller satellites presents many challenges, and these satellites can then be launched by new commercial launch providers at significantly reduced costs.

The exemplary satellite 100 shown in the figure includes a radar antenna array 160 extending from bus 110 in two opposite directions to provide one of two "wings" with a generally flat structure. The satellite's "flight" direction is typically parallel to these opposite directions, for example, parallel to the longitudinal extent of the antenna array and, in this example, the longitudinal extent of the solar panel array. The radar antenna array 160 includes antenna elements and associated electronic circuitry systems, as described further below. The radar antenna array 160 is shown mounted on or adjacent to a rectangular surface of bus 110. The radar antenna array 160 can be configured as segments, referred to herein as panels, folded for transport on Earth and into space, and unfolded during deployment, similar to a solar panel. Solar panel supports 115 can be constructed such that solar panels 155 can be transported close to and positioned away from the main body 110, as illustrated in satellite deployment diagrams. The structure of the main body 110 and the radar antenna array 160 can be collectively referred to as the aircraft frame.

Figure 3 is a perspective view of one of the satellites in Figures 1 and 2, where the radar antenna array 160 and the solar panel array are folded against bus 110 during transport and launch to minimize the volume of satellite 100. Once the satellite is launched into space and separated from the rocket, it is instructed (e.g., from a ground station, also referred to in the art as a ground segment, not shown) to deploy the solar panels and radar antenna array. This is done by unfolding the solar panel array and radar antenna array 160 until they reach the configuration shown in Figure 2.

The designs described herein are not limited to a configuration including a central body or bus, and other configurations of the ADCS and thrusters are possible.

In the deployment configuration shown in Figures 1 and 2, bus 110 is cubic. Referring to Figure 1, solar panels 150 and 155 are configured in a generally flat arrangement extending in one of two opposite directions parallel to the extension direction of radar antenna array 160, wherein the plane of the solar panel is perpendicular to the plane of radar antenna array 160. The two planes may be parallel to their respective sides of the cube. In this configuration, the plane of radar antenna array may face the Earth and the plane of solar panel may face the Sun. It should be noted that the solar panels, once deployed, are in a fixed relationship with radar antenna array 160, contrary to known designs in which the solar panels may be relative to the satellite body and thus the radar antenna array may rotate or pivot. The lack of control over solar panels helps reduce the overall weight of the satellite.

Radar antenna array 160 includes a plurality of antenna elements mounted on its main surface opposite to bus 110. Other components may be included in the radar, as known in the art, including power distribution components and amplifiers, examples of which are described in our earlier patent application WO2020094872A1 and discussed further below. The antenna elements collectively constitute what is commonly referred to as a satellite antenna. This constitutes a SAR antenna, i.e., an antenna through which SAR signals are transmitted and received.

The illustrated satellite 100 includes a separate antenna 210, referred to as the downlink antenna, for communication with a ground station and/or other space or spacecraft. Therefore, in this exemplary satellite, the antenna array is primarily used for the transmission and reception of SAR signals. Data exchanged via antenna 210 may include raw or processed SAR data downloaded to Earth, instructions for satellite operation, and other data familiar to those skilled in the art. The downlink antenna 210 may have an aperture smaller than that of the antenna array 160. In addition to the downlink antenna, the illustrated satellite 100 may also include one or more additional antennas for telemetry to the ground. These telemetry antennas are typically smaller than the SAR antenna or the downlink antenna 210.

Referring to Figure 2, the antenna element can be a patch antenna arranged in a common plane. The total area of the antenna element (i.e., the total area of the radar antenna array 160 in the satellites of Figures 1 and 2) defines the antenna aperture. In the example shown in Figure 1, the aperture has an area of 3.2m x 0.4m = ^1.28m² and an aspect ratio of 16 in the flight direction, meaning the length in the flight direction is 16 times the width. In another example, the width of the antenna array is increased to a height of 0.8m ( ^2.56m² ), while still using a similar design and modular construction. In this case, the aspect ratio is 4. Generally, the aperture can have any size between about 1 ^m² and about 3 ^m² or between about 0.5 ^m² and about 5 ^m² , and the aspect ratio can be between 2 and 24. The antenna is shown in Figure 1 as divided into 5 panels 161, 162, 163, 164, and 165.

Antenna elements can operate as a phase array to steer the radar beam in either or both of azimuth and elevation. Steering can be active, as described further below. The term "active" in this context refers to amplification and, where appropriate, also to power supplied individually or in groups to the antenna elements rather than collectively to the entire antenna array.

Radar antenna array 160 is included in what is referred to as the satellite "payload." The term "payload" generally refers to the satellite-borne instruments used to perform the satellite's intended functions (such as Earth observation). Therefore, any equipment other than the satellite propulsion system, attitude determination and control system (ADCS), and other essential requirements for satellite operation can be considered part of the payload. Therefore, some components of radar antenna array 160 that may be included in the payload and housed in bus 110 may be included in the payload.

In one example, the radar antenna array may include a single-channel, VV-polarized phase array synthetic aperture radar. Three main imaging modes are known in SAR imaging: strip, spot, and scan. In the satellite illustrated in Figure 1, this has been achieved using phase array techniques.

There is a bias in the technology against using active phased array SAR technology on small satellites, such as those of the type discussed above, which can be launched by commercial providers. In some circles, this is considered infeasible. Other designs, such as those using SAR reflectors and those using waveguides instead of active phased arrays, are actively under development. Through a series of developments described further here, the inventors have successfully deployed active phased array technology in small satellites and achieved outstanding image resolution. For example, the skew distance resolution in the cross-track direction has reached 0.5m, providing a resolution of approximately 1m (depending on the viewpoint). Even higher resolutions are possible. Resolution can also be achieved in the azimuth direction (e.g., as low as 25cm).

One problem with using a phased array SAR with a small satellite is thermal management. This is particularly important in the case of the satellite described here, which has a relatively small SAR antenna array. A traditional expectation of a small satellite is that a high-resolution image cannot be achieved using an active phased array SAR because the small SAR antenna cannot transmit sufficient power to produce a good image quality. Generally, image quality varies depending on the total transmitted power (which is related to the size of the antenna and the power density transmitted by the antenna). In the exemplary satellite described here, high-quality imagery is made possible by designing the satellite to operate at a power density higher than previously thought achievable to achieve good imaging performance.

A key feature of the satellite disclosed herein is the compact placement of amplifiers and antenna elements (e.g., in the satellite's wings rather than in the main bus. This reduces signal loss, aids in thermal management, and allows for higher transmission power density. According to one example, the antenna elements included in antenna array 160 have dedicated amplification positioned close to the antenna elements. This can be a modular configuration, such that amplification and, where appropriate, associated power components are distributed across radar antenna array 160. In this modular configuration, each panel 161 to 165 may include a module, or each panel may include multiple modules, with separate amplification provided for each module. This distributed amplification configuration also provides a high degree of resilience against component failures that would otherwise occur in a centralized amplification system. In one particular example, each module includes multiple RF front-ends, each connected to one or more antenna elements, such as surface-mount antenna elements. Each front-end may include a power amplifier and a low-noise amplifier. In the illustrated example, each front-end uses a 20-watt power amplifier chip from Qorvo in Greensboro, North Carolina, to power eight antenna elements. Depending on the specific antenna design, other RF power amplifiers with different wattages powering different groups of antenna elements are also feasible. For example, a lower-wattage (e.g., 5W) power amplifier can be used with an antenna four times larger to achieve the same transmission power. Similarly, a higher-wattage (e.g., 30W) power amplifier can be used with a smaller antenna to achieve the same transmission power, provided appropriate thermal management is employed. In one example, power amplifiers between 5W and 30W can be used.

The ability to supply power to individual antenna elements or groups of antenna elements, rather than to the entire antenna array as a whole, enables the phase array to be actively controlled.

The characteristics of an example satellite and its payload, as shown in Figures 1 and 2, are listed in Table 1.

In the examples above, the frequency band is the X-band (8 GHz to 12 GHz). Some of the designs discussed here can operate in the C-band, with frequencies ranging from 4 GHz to 8 GHz.

The overall satellite design, including the payload, differs from existing satellites in many ways and has been demonstrated to achieve outstanding performance. However, the cost-effectiveness of building such a satellite can be very high. For example, the payload can be designed using commercially available off-the-shelf components and standard PCB manufacturing processes. The small size and relative simplicity of the satellite are also important factors in reducing manufacturing and launch costs.

Figure 4 shows a schematic block diagram of a portion of an example payload 400. The illustrated payload has a core unit called a processing board 401, which contains an application processor that can be one of the field-programmable gate arrays (FPGAs) 402 (shown in Figure 4), buffered random access memory (RAM) 403, and flash memory 404. The FPGA 402 can store up to 100 individual waveforms for imaging. Each waveform can be generated in a ground segment for a specific imaging pattern and geometry (e.g., tilt) and uploaded to the satellite.

The waveform can be fed to a differential digital-to-analog converter (DAC) 461, which provides separate I and Q signals based on the baseband frequency. The signal is transmitted from the satellite to a target (e.g., a target to be imaged) via transceiver 421 and amplifier 450. The echo signal is then received due to reflection of the transmitted signal from the target. Upon reception on the satellite, the received signal is received by transceiver 421 and converted to digital using an analog-to-digital converter (ADC) 462. The sample from the ADC is passed through a decimation filter and compressed (not shown). Due to the high data rate, the data is stored in a RAM buffer and then written to flash memory for later downlink processing.

The baseband signal is modulated to an intermediate frequency (IF) at transceiver 421 and then up-converted to RF. After a further amplification stage, the signal is split and fed to five outputs on radar panels 161 to 165. On the receiver side, the same 5-channel splitter design is used in a combiner mode, and the signal is fed through a low-noise amplifier (LNA) and down-converted, filtered and amplified by an IF, and then demodulated to become the I and Q baseband components. The receiver gain can be digitally adjusted by the IF, although in practice this feature is of little use because the ADC can tolerate a wide dynamic range.

The components shown in Figure 4 can be housed in bus 110. In the exemplary configurations of Figures 1 and 2, the radar antenna array 160 is divided into five panels 161 to 165. Each of these panels may contain four transmit/receive (T/R) modules, for a total of twenty modules. Each module may include multiple RF front-ends, as further described with reference to Figure 6. The center panel 163 can be attached to the aircraft bus 110 and panels 161, 162, 164, and 165 (two on each side) can be deployed against the bus 110 for transmission, as shown in Figure 3. Figure 2 indicates a set of T/R modules 165a to 165d.

A beamforming network used to transmit signals from bus 110 to a panel or module may include a phase-matched coaxial cable harness and a microstrip splitter for separating signals between modules. The remaining phase-matching of the harness can be calibrated using a custom-designed linear scanner in an anechoic chamber.

Each T/R module 165a to 165d, similar to all other modules, can be composed of a stack, for example, a group of flat elements arranged side-by-side, one above the other, as shown in Figure 5. Here, the flat elements include a power distribution board 514, a heat dissipation board 513, an RF board 512, and an antenna board 510. The heat dissipation board 513 is an example of one or more heat dissipation elements primarily used to absorb and conduct heat from the power distribution board and the RF board. Support structures 516 and 518 are also provided relative to the edges of their respective boards, which can also act as heat dissipation elements to conduct heat from the RF board and possibly the power distribution board to the antenna board. As mentioned above, an important aspect of the satellite design described herein is the placement of the amplifier close to the antenna elements. This achieves significant advantages compared to previous designs. A further advantage can be achieved by placing the power components closer to the amplifier receiving the power, rather than (for example) at bus 110.

Heat is generated during the operation of the radar antenna array 160 to collect imaging data. In the structure shown in Figure 5, heat can be conducted from the amplifier board 512 to the antenna board 510 via the heat sink 513 and/or support structures 516, 518, from which the heat can be radiated. Heat can also be conducted from the RF board to the antenna board via the RF connectors. The opposing surface of the power distribution board 514 can point towards the sun, while the antenna array is directed towards the Earth, and therefore heat can be expected to be radiated from the satellite via the antenna board rather than the power distribution board. However, sometimes the back of the power distribution board can point towards space; in this case, they can also be used for radiating heat. The opposing surfaces of the four power distribution boards 514 can be seen on the back of the antenna module 165 shown in Figure 3 in its mounted position.

The power distribution board 514 can support power components such as regulators and switches for RF components on the bias RF board, and, where appropriate, support one or more functions such as (but not limited to) telemetry collection and control of electronic beam steering.

RF board 512 can support multiple (e.g., sixteen) identical front-ends as shown in Figure 6. Furthermore, the RF board can support a driver amplifier and separate splitter/combiner networks for both the transmit and receive sides. An example front-end 600 and associated antenna patch are schematically illustrated in Figure 6.

The heatsink or plate 513 may comprise a sheet of a suitable material with good thermal conductivity, such as a metal (e.g., aluminum). The heatsink 513 may be sandwiched between the power distribution board 514 and the RF board 512 such that adjacent main surfaces of the heatsink 513, power distribution board 514, and RF board 512 are in contact to facilitate heat conduction from the RF board. Holes are provided in the heatsink to allow power and control signal connections between the power distribution board and the RF board. In one example, a suitable thickness of the heatsink is 0.5 mm to 3 mm.

The components mounted on RF board 512 are shown connected to antenna elements supported on the antenna board via RF connectors 531 to 535 (which may be coaxial snap-fit connectors). This schematic shows only five connectors. In practice, the number of connectors may correspond to the number of RF front-end components.

Figure 6 shows a component suitable for a front-end 600 connected to one of the antenna patch arrays 620. Sixteen front-end 600 components can be mounted on a single RF board 512. Each front-end may include a 20W-class solid-state power amplifier (PA) 601 on the transmit side, a digital phase shifter 603 in front, and a T/R switch 605 behind, and a front-end LNA 607 and a digital phase shifter 609 on the receive side. Amplifiers 601 and 607 are switched on and off to follow the transmitted waveform. The antenna is shown as a patch array, with eight patches fed from a single front-end and connected to the RF board (e.g., via connectors 531 to 535). Antenna patch array 620 is a subarray of radar antenna array 160.

The number of antenna elements driven by a single front end depends on the specific geometry chosen for the radar antenna array, such as the number and size, or the panel and module. This number is a power of 2, such as 4, 8, 16, 32, etc. This is convenient but not mandatory.

Therefore, it should be understood that a single board can be used for multiple transmit/receive modules. In this example, the same RF board 512 contains 16 PAs, 16 LNAs, transmit/receive switches, and other components. Providing multiple front-ends on the same board is advantageous because it allows high power to be transmitted from the antenna in a way that is both energy-efficient and space-saving. Generally, the number of front-ends included on a single RF board is limited only by mechanical and thermal considerations.

Figure 7 is an enlarged view of a portion of the radar antenna array of Figure 2, showing antenna panels 165 and 164 with four modules 165a, 165b, 165c, and 165d. Each module includes sixteen rows of antenna patches, with eight antenna patches in each row. An antenna patch is indicated by symbol 710, and symbol 730 indicates a row or subarray of antennas. Sixteen front-ends (not visible in Figure 7) on the RF board behind the antenna panels drive each antenna module, with each front-end driving a row of eight antenna patches. Each antenna patch may be referred to as an element of the SAR antenna or radar antenna array 160.

Figure 8 is a perspective view of an assembly including an RF board 800 and a power distribution board 801, which may be included in a transmission/reception module. The RF board 800 and power distribution board 801, shown in Figure 8, can be separated by a support structure that conducts heat from the RF board 800 and also functions as a heatsink. A heatsink fin or plate (not shown) may or alternatively be sandwiched between the RF board 800 and the power distribution board 801.

The components on the top surface of RF board 800 are visible in Figure 8. The ellipse 802 on the right side of RF board 800 indicates an RF front-end as shown in Figure 6, and symbol 807 indicates a coaxial connector. In one example, each RF front-end uses the same components and each front-end drives eight antenna elements. For example, components 805 and 806 are power amplifiers for the two RF front-ends, corresponding to PA 601 in Figure 6. Similarly, components 803 and 804 are transmit and receive RF switches corresponding to SW 605. Component 809 is a low-noise amplifier corresponding to LNA 607. Components 808 and 810 are phase shifters corresponding to phase shifters 603 and 609 in Figure 6.

The configuration of components described above has been used to achieve a power density greater than previously considered feasible for a SAR satellite. The peak radiated power from the 3.2 x 0.4 m antenna is 3.2 kW, equivalent to a peak power density of 2,500 W per square meter. This power is transmitted during the transmission phase of imaging. The antenna is in transmission approximately 20% of the time and in reception approximately 80% of the time. Therefore, the average power density during imaging is 500 W/ ^m² . Generally, some of the advantages of the design described herein can be achieved with peak power densities greater than 1000 W/ ^m² and/or average power densities greater than 200 W/ ^m² .

The same design can be used to double the total peak power to 5,000 W by, for example, doubling the area of the SAR antenna, while maintaining the same peak and average power densities. This can be easily achieved using the same modular components used in the exemplary satellites described in this invention. Alternatively, with appropriate thermal management, the total transmitted power can also be doubled by maintaining the same antenna size and doubling the peak power density to 5,000 W/ ^m² and the average power density to 1,000 W/ ^m² . In one example, for a small satellite capable of achieving high image resolution, the range of peak and average transmitted power densities is: 1,000 W/ ^m² to 5,000 W/ ^m² for peak transmitted power and 200 W/ ^m² to 1,000 W/ ^m² for average transmitted power. Other ranges are feasible. Total transmission power is limited by the size of the SAR antenna available on a small satellite and the ability of the design to handle the heat generated by higher power density, as described herein.

The entire payload can be self-contained within one of the onboard computer configurations in bus 110 using a differential asynchronous data bus (which can also be used to collect telemetry data from each front end). Additionally, two global synchronization signals generated by the processing board, following the RF waveform, may be present: one for triggering switching between TX and RX modes on the T/R module for each pulse, and another for triggering beam switching (e.g., in scan operation mode).

In scan mode, the SAR radar beam is oriented in both range and azimuth. To achieve high-speed synchronous beam switching, the oriented angle matrix can be broadcast to the T/R module before imaging, and the phase shifter oriented vector can be pre-calculated for each beam. These can then be sequentially loaded from a lookup table of each beam oriented trigger pulse.

Depending on the exact wavelength λ, the usable electron shunting angle is limited by the fact that the spacing between antenna elements is approximately 0.8λ in the range and approximately 5.2λ in the azimuth. Therefore, large-angle electron shunting can lead to grating lobes. This can be mitigated to some extent by a combination of mechanical and electronic shunting, which can be achieved using ADCS to manipulate the entire satellite, especially in the case of a small satellite. In other words, the entire satellite can be manipulated to orient the antenna relative to Earth, rather than relative to the satellite itself, which is feasible in a very large satellite. Therefore, in strip and scan modes, the entire satellite can be manipulated on the roll axis, for example, rotating around an axis parallel to the X direction in Figure 2 to point the antenna to the appropriate viewing angle. In spot mode, manipulation can be performed on all three axes to maintain a point on Earth and also maintain local V-polarization of the radar beam in the target area.

Payload data download is achieved via a primary X-band transmitter connected directly to the high-capacity memory 404 on the processing board 401 in Figure 4. A speed of up to 500 Mbps is achieved using 8-phase shift keying (PSK) modulation. The transmission power is 4W via a medium-gain patch antenna, implemented using the same technology as radar antennas. A redundant auxiliary X-band transmitter is provided for backup purposes.

Constructing a small SAR payload introduces numerous limitations and forces several trade-offs. A major challenge to image quality is the small aperture. To achieve reasonable sensitivity, a relatively high pulse power can be used to compensate for the reduced antenna gain. The resulting high power density presents challenges for thermal management, as previously described. An additional consequence of a small aperture is increased range and azimuth ambiguity, leading to image quality degradation. To reduce this ambiguity, many innovative methods can be employed to suppress ambiguity during processing, such as those using Do... The methods described by An et al. in "Experimental Demonstration of a Novel End-to-End SAR Range Ambiguity Suppression Method" 2022 IEEE RadarConf22, 2022, pp. 1-6, doi: 10.1109/RadarConf2248738.2022.9764267 and by Radius et al. in "Phase Variant Analysis Algorithm for Azimuth Ambiguity Detection" 2022 IEEE RadarConf22, 2022, pp. 1-4, doi 10.1109/RadarConf2248738.2022.9764162.

High transmission power presents challenges for power system design and thermal management. The feasible solar panel area for a small satellite is a limiting factor in the design and operational effective load of the power system. As shown in Figures 2 and 3, one way to increase the solar panel area and thus the power available for a small satellite is to have deployable solar panels that are folded for transport and launch into space and unfold when the satellite reaches orbit. In one example, the panel contains 112 solar cells, placed on a surface away from bus 110, and arranged in a generally flat configuration extending in two opposite directions parallel to the extension direction of the radar antenna array, wherein the plane of the solar panel is perpendicular to the plane of the radar array. When directly pointed to the sun, each panel provides up to approximately 90W of power. During nominal operation, the average rail power is approximately 350Wh.

Since solar panels cannot always be pointed towards the sun (e.g., when the satellite is in Earth's shadow), any limitations on the area of the solar panels can be at least partially compensated for by using a battery to accumulate and store energy between image acquisitions. An ultra-large battery can be used, which also allows the battery to operate in a manner that maximizes its lifespan (e.g., by never discharging it below 60% and never fully charging it). For example, the battery could comprise an array of modules, such as 20 modules, connected in parallel to deliver approximately 33 volts and capable of storing approximately 850 Wh. These could each consist of eight lithium-ion batteries connected in series. This allows for a constant voltage supply for nominal operation and withstands higher loads during SAR transmission in imaging mode. The battery array can be directly charged from the solar panel and can also supply power to a power regulation module (not shown) for further distribution. The battery array can be housed in bus 110.

Thermal management at the radar antenna array panel can be passive, for example, without power supply components, where heat from the power amplifier on the RF board is absorbed into the antenna structure (e.g., via a heat sink) and then conducted to the antenna panel for radiation. In the illustrated satellite, the thermal management of the antenna array is separated from the thermal management of the components within the bus.

Satellite thermal management can be achieved by providing one or more radiating surfaces configured to radiate heat accumulated at the satellite into space. For the wings, the antenna array surfaces are also used to radiate heat generated by the distributed power amplification system during imaging. For components located in the bus, a radiating surface can be disposed on the satellite bus 110 to radiate heat generated in the bus. The radiating surface may include any suitable material known in the art, such as a polyvinyl fluoride film.

Figure 9 is a magnified front view of a satellite from the direction-y (i.e., from the back of the bus and solar panel) of an example of the radiator panel 196 shown in Figure 1. Figure 9 shows a radiating surface disposed on a radiator panel 196 positioned on a surface of the bus. In this example, the radiating surface is on a surface opposite the solar panel such that it faces in a direction opposite to the outward-facing surface of the solar panel, thus generally pointing as far away from the sun as possible. It should be understood that this design can be used in a configuration with only one solar panel.

In the examples of Figures 1, 2, and 9, panel 196 includes a radiating surface located on the surface of a busbar identical to one or more thrusters. In one example, the radiator panel may have a generally circular shape, although it is not limited to a circular shape and other shapes are possible.

Panel 196 may include a structural panel forming part of the existing structure of satellite body 110, as shown in FIG. 9. Here, panel 196 has a substantially circular shape, and a plurality of mounting holes surround its outer circumference. Thus, the radiation panel may be an integral part of one side of satellite bus 110. In one example, the radiation panel may even be part of the structure (and thus the mounting holes) for attaching the satellite to a release mechanism in a rocket when launched into space. The radiation panel may be made of aluminum having a relatively high thermal conductivity. Thus, the temperature is (relatively) uniform on the panel, and heat can be conducted from one or more components to the radiation panel in a first location where cooling is required and from the radiation panel to one or more components in a second location where heating is required. This is particularly advantageous for small aircraft due to strict mass constraints. Heat is transmitted and absorbed on the reverse side of the radiating surface, and also conducted to and from the structure via structural components (not shown) that connect the radiating surface to other parts of the aircraft. The satellite can use an ADCS to manage heat transfer to and from the structure between the first and second positions.

The decision to use only passive thermal management can lead to a limitation on the amount of time the payload can be turned on, as overheating can occur if it is turned on for too long. This, in turn, limits the number of images or the amount of image data that can be acquired in a single orbit. The limitation on the amount of time the payload can be turned on, and therefore the limitation on the amount of image data that can be acquired in a single orbit, leads to a necessary trade-off for small SAR, which means a limitation on the number of images that can be acquired and their duration in a single orbit. However, this has been found to be acceptable, especially in applications where maximizing the imaging time in each orbit is less important than, for example, minimizing the time between duplicate images of a particular area on Earth. Therefore, a method for operating a satellite as described herein is also provided to limit the on-time of the radar antenna array to avoid overheating. This can be achieved by limiting the total on-time during a single image acquisition period, or by limiting the total on-time during a single orbital period, or both. On-time control can be performed from the ground segment or programmed into the satellite's onboard computer. A suitable maximum value can be determined experimentally. In one example, an alarm limit is set such that if the antenna array exceeds a maximum angle of 45 degrees Celsius, imaging is cut off, even if individual components in the antenna array are rated for 80 degrees Celsius or higher. The alarm limit is significantly lower than the maximum capability of the components to improve reliability.

Similarly, a lower limit can be set so that imaging does not occur when the antenna array temperature drops below -25 degrees Celsius, even if individual components of the antenna array can be rated as low as -40 degrees Celsius. For an Earth monitoring satellite, the time required to complete one orbit around the Earth in low Earth orbit can be approximately 90 minutes. In one example, 75 seconds of imaging per orbit can be achieved within thermal constraints. A suitable maximum imaging time (i.e., the image data acquisition time during a single orbit) is 180 seconds, and a suitable maximum design single image duration is 120 seconds. In other words, with appropriate limitations on the effective payload uptime, a small satellite can be passively thermally managed without dedicated cooling equipment. According to some examples, the limitation could be approximately 3% to 4% of the orbital duration (i.e., the time required for the satellite to orbit the Earth once). However, higher limitations may become possible with greater thermal management capabilities.

Heat dissipation elements such as heat sink 513 and support structure 516 can act as a buffer, storing heat to be radiated at a later time. The amount of imaging time in an orbit is limited by the rate at which heat can be radiated from the antenna, making larger heat sinks and potentially more radiating surfaces helpful in increasing imaging time.

Imaging can be performed in different modes, and the effective imaging time can vary depending on the imaging mode. For example, the longest time for continuous imaging will typically be shorter than the longest time for another imaging mode in which the radar antenna array is switched on and off at different times in orbit. Disconnecting the antenna between imaging sessions within an orbit allows the antenna to cool to some extent, enabling it to image over a longer total time.

In the design discussed herein, heat will not be conducted back to the bus and radiated from the emitter, although this possibility cannot be ruled out. In the design discussed herein, the bus 110 and the radar antenna array are thermally isolated from each other.

Although a typical orbital period of 90 minutes for a satellite in low Earth orbit around the Earth does not necessarily mean that the satellite can repeatedly image a specific location on Earth every 90 minutes. Typically, a satellite does not continue on the same circular orbit around the Earth, but rather follows a more complex orbital pattern, preventing it from reappearing at the same location approximately every 90 minutes. This may be undesirable for continuous monitoring of positions on Earth. However, the orbital pattern of a single satellite can be designed so that the satellite can image the same location on Earth approximately once a day. Furthermore, by operating multiple satellites in a constellation, a point on Earth can be imaged by more than one satellite, providing more frequent monitoring of a specific location. In one example, a suitable constellation of small satellites can achieve repetition times of 24 hours or less, 12 hours or less, 6 hours or less, or even 4 hours or less. The small satellites according to the invention are particularly suitable for applications where frequent revisits to a specific location on Earth are required and where the amount of imaging that can be completed over an entire orbit is not critical. Their lower cost also makes them particularly suitable for deployment in large satellite constellations that allow for shorter repetition cycles, such as those with five or more satellites, 10 or more satellites, 15 or more satellites, or 25 or more satellites.

Figure 10 is a schematic diagram illustrating a feasible configuration of one of the satellite electronic components used in any of the satellite designs described herein. One-way solid arrows between components indicate power connections, two-way solid arrows indicate RF signal connections, and dashed lines indicate data connections.

In the configuration shown in Figure 10, some components are located at the satellite bus, indicated by rectangle 1020, and some are located at the radar antenna array, indicated by rectangle 1030. The configuration shown in Figure 10 includes a power supply 1001 and a power distribution system 1002. The power supply 1001 may include the solar panels and batteries described above. The power distribution system 1002 supplies power to a propulsion system 1090, which includes, for example, a thruster 195, a propulsion controller 1009, an attitude determination and control system (ADCS) 1031 (which may include the ADCS 190 of Figure 1), a computing system 1003, a buffer 1035, and a communication system 1004 (which may include a downlink antenna 210). Although presented as a single item, buffer 1035 may be included in computing system 1003 and may include the aforementioned high-capacity (e.g., flash) memory 404.

The propulsion controller 1009 is shown here as a separate item, but in reality, it can form part of the computing system 1003. The propulsion controller can be configured to control the propulsion system 1090, either by using control software implemented in one or more processors included in the propulsion controller 1009 or in response to, for example, instructions received from the computing system. When instructions are transmitted from the computing system 1003, the computing system can be considered to include a propulsion controller. One function of the propulsion controller 1009 can be to output control signals to the ion and electron sources of the thrusters in the propulsion system 1090.

Referring to Figure 5, the radar antenna array 160 shown in Figure 6 includes a plurality of modules, each including an RF board 512 on which an amplifier is mounted, a power distribution board 514 on which power components are mounted, and an antenna board as shown in Figure 7 supporting the antenna elements. Each RF board includes multiple front ends. In Figure 10, for simplicity, these components are uniformly represented as a single block. Thus, antenna block 1006 represents a plurality of antenna elements of all modules, amplifier block 1007 represents, for example, a plurality of RF front ends included in radar antenna array 160, and power distribution system block 1008 represents a plurality of power distribution components included in the plurality of modules of radar antenna array 160. Both power distribution systems 1002 and 1008 may include control logic known in the art.

Antenna 1006, together with amplifier 1007 and power distribution system 1008, forms the satellite's image acquisition equipment, as those skilled in the art will know. They can perform functions other than image data acquisition.

Amplifier block 1007 has a bidirectional data communication link with computing system 1003, which, in the illustrated example, is via power distribution system 1008 and can be configured to send data, such as data related to received radar signals, to computing system 1003. The data can be processed by computing system 1003, for example, to generate images, which can then be output to communication system 1004 for forward transmission. In the system illustrated in Figure 10, raw data is output from computing system 1003 to communication system 1004 for processing by a remote computing system. In Figure 10, a SAR processor 1033 may be located, for example, at a ground station or another processing location.

Raw SAR data can be stored in the satellite's memory (e.g., buffer 1035). In one example, 30 seconds of imagery can be stored at full resolution (bandwidth). More can be stored at lower resolution (e.g., 60 seconds at half resolution). In one example, a small satellite can have a 150 MB/s download link. At this data rate, downloading 30 seconds of full-resolution imagery data would take approximately 3 minutes.

During operation, such as in spotlight mode, approximately 5,000 pulses per second can be transmitted. This means that at any given time, there can be 27 pulses in the air.

The communication system 1004 can communicate with earth stations or other satellites using radio frequency communication, optical (e.g., laser communication), or any other form of communication as known in the art.

Figure 11 illustrates one of the feasible configurations of the components shown in Figure 10. Here, the bidirectional bus connections between the components are indicated by dual lines, with solid-lined arrows indicating power connections and dashed-lined arrows indicating data connections. The configuration typically includes a power system 1101, an attitude and orbit determination and control system (ADCS/AOCS) 1102, a telemetry and remote control system 1103, onboard data processing 1104, a payload downlink 1105, and a SAR payload 1106.

Onboard data processing can be performed by an onboard computer. Figure 11 shows the primary and auxiliary onboard computers 1108 and 1109 used for this purpose, which can share tasks and/or provide redundancy. The onboard computer can be used to limit the acquisition of image data for thermal management purposes, as described elsewhere herein.

A monitoring board 1111 may be provided, for example, including a circuit system for protecting onboard computers 1108, 1109 from power failures, as known in the art.

The power system 1101 includes the aforementioned solar panels 150 and 155, battery 1115, and power regulation module 1117.

The ADCS/AOCS 1102 includes an ADCS 1120 and a propulsion controller 1121 for controlling satellite orbit. These may include components for controlling, for example, the reaction wheel of an ADCS and the thrusters of a propulsion system 1090, as well as separate computational devices for controlling satellite attitude and orbit.

The telemetry and remote control system 1103 may include primary and auxiliary S-band (2 GHz to 4 GHz) transceivers for communication between the satellite and a ground segment.

The payload downlink 1105 includes primary and auxiliary X-band transmitters 1151 and 1152 for downloading SAR payload data to a ground segment.

The SAR payload 1106 includes a processing board 401, a transceiver 421, a signal enhancer 1161, and a radar antenna array 160 comprising 20 SAR modules. The signal enhancer 1161 provides power and control signals to the splitter board, as well as enhances the signal for the SAR array.

Any of the computers or computing systems described herein may be a known computing system in the art and may include one or more controllers, such as (for example) a central processing unit (CPU), a chip or any suitable processor or computing device, an operating system, memory, and storage. One or more processors in one or more controllers may be configured to operate the satellite as described above. For example, one or more processors within a controller may be connected to memory storing software or instructions that, when executed by one or more processors, cause one or more processors to operate the satellite.

The functions performed by the computers and computing systems described herein can be distributed in any suitable manner and do not necessarily require a dedicated computer. Therefore, for example, a computer or computing system may perform some functions of ADCS and AOCS in addition to onboard data processing.

Some operations of the methods described herein can be performed at a ground station by software in a machine-readable form (e.g., in the form of a computer program including computer program code), as in some examples. Therefore, some embodiments of the invention provide a computer-readable medium that, when implemented in a computing system, causes the system to perform some or all of the operations of any of the methods described herein. The computer-readable medium can be in a temporary or tangible (or non-temporary) form, such as storage media, including magnetic disks, flash drives, memory cards, etc. The software can be adapted to execute on a parallel processor or a serial processor such that the method steps can be performed in any suitable order or simultaneously.

It should be understood that the above benefits and advantages may relate to one embodiment or several embodiments. The embodiments are not limited to any or all of the embodiments that solve the described problem or that have any or all of the described benefits and advantages.

Unless otherwise stated, any reference to "a" article or "part" refers to one or more of such articles. The term "comprising" is used herein to mean including identified method steps or elements, but such steps or elements do not include an exclusive list and a method or apparatus may contain additional steps or elements.

Furthermore, when the term "comprising" is used within the scope of [implementation] or [scope of the patent application], this term is intended to encompass in a manner similar to the term "including," because when used in a claim, "including" is interpreted as a transitional word.

The diagrams illustrate illustrative methods. Although methods are shown and described as a series of actions performed in a specific sequence, it should be understood that methods are not limited to a specific sequence. For example, some actions may be performed in a different order than that described herein. Furthermore, one action may be performed concurrently with another. Moreover, in some examples, not all actions are required to implement one of the methods described herein.

The order of steps in the methods described herein is illustrative, but these steps may be performed in any suitable order or, where appropriate, simultaneously. Furthermore, steps may be added or substituted, or individual steps may be deleted from any part of the methods, without departing from the scope of the subject matter described herein. The manner of any of the foregoing examples may be combined with the manner of any of the other examples described to form further examples.

It should be understood that the above description of a preferred embodiment is given by way of example only, and various modifications can be made by those skilled in the art. The content described above includes examples of one or more embodiments. Of course, in order to describe the above embodiments, it is impossible to describe every conceivable modification and variation of the above apparatus or method, but those skilled in the art will recognize that many further modifications and substitutions of the various embodiments are feasible. Therefore, the described embodiments are intended to cover all such changes, modifications, and variations that fall within the scope of the appended patent application.

Examples and embodiments of this invention may be understood in accordance with the following terms:

1. A satellite for operation in orbit around the Earth, comprising: a propulsion system, an attitude determination and control system (ADCS), and a radar antenna array; wherein: the radar antenna array is formed as a generally flat structure, including a plurality of panels, each panel including a plurality of antenna transmission/reception modules; each antenna transmission/reception module includes a flat element stack comprising an RF board and an antenna board, and one or more heat dissipation components for heat conduction from the RF board; and each RF board supports a plurality of RF front ends.

2. As in Clause 1, each RF board comprises between 8 and 32 RF front-ends.

3. As with any of the foregoing terms, each RF front-end of the satellite includes a power amplifier, a transmit/receive switch, and a low-noise amplifier.

4. Any satellite mentioned in the foregoing clauses is configured for use in the operation of the radar antenna array as a phase array.

5. As in Clause 4, each RF front end includes a number of phase shifters in both the transmission and reception directions.

6. As with any of the foregoing terms, the satellite wherein the antenna plate supports an array of antenna elements.

7. Satellites as described in Clause 6, wherein such antenna elements include surface mount antenna elements.

8. Satellites as described in Clause 6 or 7, wherein each RF front-end is configured to drive a plurality of antenna elements.

9. As with any of the preceding clauses, the area of the radar antenna array is between 1 ^m² and 5 ^m² .

10. As with any of the preceding clauses, the area of the radar antenna array is between 1 ^m² and 3 ^m² .

11. As with any of the preceding clauses for a satellite, wherein the area of the radar antenna array has an aspect ratio between 2 and 24.

12. The satellite as described in any of the foregoing clauses, wherein the radar antenna array comprises a row of panels, wherein adjacent panels are pivotable relative to each other, thereby allowing the array to be folded for transport and deployment in space.

13. The satellite as described in any of the foregoing clauses includes a solar panel array that is in a fixed relationship with the radar antenna array.

14. As in any of the preceding clause 13, wherein the solar panels are configured in a generally flat arrangement extending in one of two opposite directions parallel to the direction of extension of the radar array, wherein the plane of the solar panels is perpendicular to the plane of the radar array.

15. A satellite as described in Clause 13, wherein the solar panels are configured such that adjacent panels can rotate relative to each other so that the solar panel array can be folded for transport and deployment in space.

16. The satellite, as described in any of the foregoing terms, includes a mass memory for storing data acquired by the radar antenna array and a transmitter directly connected to the mass memory for downloading the data.

17. As with any of the foregoing clauses for satellites, wherein the flat element stack further includes a power distribution board.

18. As with any of the preceding clauses for a satellite, one or more of its heat dissipation elements include a heat dissipation plate that contacts the RF board.

19. As with the satellites of Clause 17 and 18, wherein the heat sink is disposed between the power distribution board and the RF board such that adjacent primary surfaces of the heat sink, the power distribution board, and the RF board are in contact for heat conduction.

20. The satellite as described in any of the foregoing terms, wherein the one or more heat dissipation elements comprise one or more structures configured to conduct heat from the RF board to the antenna board.

21. Any satellite as described in any of the foregoing clauses is configured to operate its radar antenna array at wavelengths between 4 GHz and 12 GHz, or as appropriate, between 8 GHz and 12 GHz, or between 4 GHz and 8 GHz.

22. Any satellite as described in any of the foregoing clauses, configured for use in the radar antenna array as a synthetic aperture radar.

23. A satellite as described in any of the foregoing clauses, wherein the ADCS is configured to cause the satellite to rotate about an axis parallel to the direction of travel along an orbit.

24. A satellite for operation in orbit around the Earth, comprising: a propulsion system, an attitude determination and control system, a power supply for powering one or more solar panels, and an active phase array synthetic aperture radar "SAR"antenna; wherein: the SAR antenna has an area of less than 5 ^m² , and has a peak transmission power of greater than 1,000 W per square meter or an average transmission power of greater than 200 W per square meter during imaging.

25. Satellites as described in Clause 24, wherein the phased array SAR antenna is configured to fold for transport and deployed once in orbit.

26. Any satellite as described in any of the foregoing clauses shall have a mass of less than 1000 kg.

27. Any satellite as described in any of the foregoing clauses shall have a mass of less than 843 kg.

28. Any satellite as described in any of the foregoing clauses shall have a mass of less than 500 kg.

29. Satellites as described in any of the foregoing clauses shall have a mass of less than 200 kg.

30. For any of the satellites mentioned in the foregoing clauses, wherein during imaging, the peak transmission power is greater than 1000 W per square meter or the average transmission power is greater than 200 W per square meter.

31. As with any of the preceding terms, wherein the peak transmission power is in the range of 1000W to 5000W per square meter, or the average transmission power is in the range of 200W to 1000W per square meter.

32. A method of operating a satellite as described in any of the foregoing clauses to acquire image data, comprising limiting the time for each orbit of a radar antenna array to be switched on to prevent the radar antenna array from overheating.

33. As in Clause 32, where the acquisition of image data is limited to 3% of the duration of the satellite orbit.

34. A satellite radar antenna array formed in a generally flat structure, comprising a plurality of panels, wherein: each panel includes a plurality of antenna transmission/reception modules; each antenna transmission/reception module includes a flat element stack comprising an RF board and an antenna board, and one or more heat dissipation components for heat conduction from the RF board; and each RF board includes a plurality of RF front ends.

100: Satellite

110: Main body/busbar

115: Pillar

150: Solar Panels

155: Solar Panels

160: Radar Antenna Array

190: Attitude and Control System (ADCS)

195: Propeller

196: Radiator Panel

Claims (29)

A satellite for operation in orbit around the Earth includes: a propulsion system, an attitude determination and control system (ADCS), and an Earth observation system for imaging the Earth's surface, the Earth observation system including a synthetic aperture radar (SAR) phased array antenna; wherein: the SAR phased array antenna is formed as a generally flat structure including a plurality of panels, each panel including a plurality of antenna transmission/reception modules; each antenna transmission/reception module includes a flat element stack, the flat element stack including: an RF board supporting a plurality of RF front ends; an antenna board; a power distribution board; and a heat dissipation plate for conducting heat from the RF board, the heat dissipation plate being sandwiched between the power distribution board and the RF board such that adjacent main surfaces of the heat dissipation plate, the power distribution board, and the RF board are in contact. For example, in request item 1, each RF board includes between 8 and 32 RF front-ends. For example, in the satellite of Request 1, each RF front end includes a power amplifier, a transmit/receive switch and a low noise amplifier. The satellite in Request 1 is configured to be used as a phase array antenna for the operation of the SAR phase array antenna. For example, in the satellite of Request 4, each RF front end includes a number of phase shifters in both the transmission and reception directions. For example, in the satellite of request item 1, the antenna board supports an array of antenna elements. For example, in the satellite of claim 6, the antenna elements include surface mount antenna elements. For example, in satellites such as Request 6 or Request 7, each RF front end is configured to drive a plurality of antenna elements. For example, the satellite in Request 1, wherein one area of the SAR phase array antenna is between 1 ^m² and 5 ^m² . For example, the satellite in Request 1, wherein one area of the SAR phase array antenna is between 1 ^m² and 3 ^m² . For example, the satellite in claim 1, wherein one area of the SAR phase array antenna has an aspect ratio between 2 and 24. As in claim 1, the satellite's SAR phase array antenna includes a row of panels, wherein adjacent panels can rotate relative to each other, thereby allowing the SAR phase array antenna to be folded for transport and deployment in space. The satellite, as claimed in claim 1, includes a solar panel array that is in a fixed relationship with the SAR phase array antenna. As in claim 13, the solar panels are configured in a generally flat arrangement extending in one of two opposite directions parallel to the direction of extension of the SAR phase array antenna, wherein the plane of the solar panels is perpendicular to the plane of the SAR phase array antenna. As in claim 13, the solar panels are configured such that adjacent panels can rotate relative to each other so that the solar panel array can be folded for transport and deployment in space. The satellite, as claimed in claim 1, includes a large-capacity memory for storing data acquired by the SAR phase array antenna and a transmitter directly connected to the large-capacity memory for downloading the data. As in claim 1, the satellite wherein the one or more heat dissipation elements include one or more structures configured to conduct heat from the RF board to the antenna board. The satellite in Request 1 is configured to operate its SAR phase array antenna at wavelengths between 4 GHz and 12 GHz, or as appropriate, between 8 GHz and 12 GHz, or between 4 GHz and 8 GHz. The satellite in Request 1 is configured for use as a synthetic aperture radar in the SAR phase array antenna. For example, the satellite in claim 1, wherein the ADCS is configured to rotate the satellite around an axis parallel to the direction of travel along an orbit. The satellite in Request 1 has a mass of less than 1000 kg. The satellite in Request 1 has a mass of less than 843 kg. The satellite in Request 1 has a mass of less than 500 kg. The satellite in Request 1 has a mass of less than 200 kg. For example, the satellite in Request 1, during imaging, has an average transmission power greater than 200 W per square meter. For example, the satellite in Request 1, wherein the peak power density is in the range of 1,000 W to 5,000 W per square meter, or an average transmission power is in the range of 200 W to 1,000 W per square meter. A method of operating a satellite such as any one of requests 1 to 26 to acquire image data, comprising limiting the time for each orbit of the SAR phase array antenna to be turned on to avoid overheating of the SAR phase array antenna. The method described in Request 27, wherein the acquisition of image data is limited to 3% of the duration of the satellite orbit. A satellite SAR phase array antenna for imaging the Earth's surface is formed as a generally flat structure comprising a plurality of panels, wherein: each panel includes a plurality of antenna transmission/reception modules; each antenna transmission/reception module includes a flat element stack comprising: an RF board supporting a plurality of RF front ends; an antenna board; a power distribution board; and a heat dissipation plate for conducting heat from the RF board, the heat dissipation plate being sandwiched between the power distribution board and the RF board such that the adjacent main surfaces of the heat dissipation plate, the power distribution board, and the RF board are in contact.