Satellite and method of operation

Abstract

Small and light weight satellites are described which may be minimally controlled, or uncontrolled, satellites tumbling or spinning in orbit. To ensure sufficient power and communications is maintained an emergency attitude control subsystem (EACS) is included and configured to generate an angular momentum perturbation upon a detection of an undesirable orientation or an undesirable spin state of the satellite. This may be detected using an attitude determination sensor, or by monitoring the solar panels or state of charge of the batteries. Multiple antennas are also used, and each antenna is connected to a different side of the satellite to ensure that communications can be maintained whilst tumbling or spinning.

Description

SATELLITE AND METHOD OF OPERATIONINCORPORATION BY REFERENCE

[0001] The present application claims priority from Australian Provisional Patent Application No. 2023904225 titled “SATELLITE AND METHOD OF OPERATION” and fded on 22 December 2023, the content of which is hereby incorporated by reference in its entirety.

[0002] The following publications and applications are referred to in the present application and their contents are hereby incorporated by reference in their entirety:PCT / AU2013 / 000895 titled CHANNEL ALLOCATION IN A COMMUNICATION SYSTEM and filed on 14 / 08 / 2013 claiming priority from Australian Provisional Patent Application No. 2012903489 filed on 14 / 08 / 2012;PCT / AU2013 / 001078 titled COMMUNICATION SYSTEM AND METHOD and filed on 20 / 09 / 2013 claiming priority from Australian Provisional Patent Application No. 2012904130 filed on 21 / 09 / 2012;PCT / AU2013 / 001079 titled MULTI-ACCESS COMMUNICATION SYSTEM and filed on 20 / 09 / 2013 claiming priority from Australian Provisional Patent Application No. 2012904145 filed on 21 / 09 / 2012;PCT / AU2014 / 000826 titled A MULTIUSER COMMUNICATIONS SYSTEM and filed on 21 / 08 / 2014 claiming priority from Australian Provisional Patent Application No. 2013903163 filed on 21 / 08 / 2013;PCT / AU2015 / 000743 titled MULTICARRIER COMMUNICATIONS SYSTEM and filed on 9 / 12 / 2015 claiming priority from Australian Provisional Patent Application No. 2014904976 filed on 9 / 12 / 2014;PCT / AU2017 / 000058 titled TERMINAL SCHEDULING METHOD IN SATELLITE COMMUNICATION SYSTEM and filed on 24 / 02 / 2017 claiming priority from Australian Provisional Patent Application No. 2016900685 filed on 25 / 02 / 2016;PCT / AU2017 / 000108 titled POSITION ESTIMATION IN A LOW EARTH ORBITSATELLITE COMMUNICATIONS SYSTEM and filed on 16 / 05 / 2017 claiming priority from Australian Provisional Patent Application No. 2016901913 filed on 20 / 05 / 2016;PCT / AU2017 / 000286 titled SYSTEM AND METHOD FOR GENERATING EXTENDED SATELLITE EPHEMERIS DATA and filed on 21 / 12 / 2017 claiming priority from Australian Provisional Patent Application No. 2016905314 filed on 22 / 12 / 2016;PCT / AU2018 / 000151 titled SYSTEM AND METHOD FOR PREDICTION OF COMMUNICATIONS LINK QUALITY and filed on 28 / 08 / 2018 claiming priority from Australian Provisional Patent Application No. 2017903470 filed on 28 / 08 / 2017;PCT / AU2021 / 000027 titled SYSTEM AND METHOD FOR ADAPTIVE COMMUNICATIONS and filed on 29 March 2021 claiming priority from Australian Provisional Patent Application No. 2020901049 filed on 3 / 04 / 2020.

[0003] The content of each of these applications is hereby incorporated by reference in their entirety.TECHNICAL FIELD

[0004] The present disclosure relates to satellites. In a particular form the present disclosure relates to the design and operation of small satellites and minimally orbitally controlled satellites.BACKGROUND

[0005] Delivering satellites into orbit is expensive, with cost scaling directly to their size and weight, which in turn relates to satellite complexity. One class of small satellites that have increased in popularity in recent years are cube satellites, also known as CubeSats which are small satellites based on a form factor of a cube with a side of length 10cm. They are typically described in terms of the number of cube units (U) occupied by the satellite. For example, a 1U CubeSat is a 10cm x 10cm x 10cm cube. They are typically low mass, e.g. a 1 U CubeSat is typically no more than 2kg, and frequently use commercial off the shelf (COTS) components with minimal or no redundancy. They are typically designed for Low Earth Orbit (LEO) missions, although some CubeSats have been developed for Lunar and Mars missions.

[0006] The benefit of the standardised form factor and weight limit is that it simplifies the design and allows the use of standardised launchers which can easily be integrated into a ridesharing mission with minimal risk to the main mission. That is when launching a larger satellite (or at least a satellite much larger than a CubeSat), there may be spare space in the launch vehicle allowing one or more standard CubeSat launchers to be accommodated in the launch vehicle to allowing several CubeSats to be launched with the main mission. Alternatively, a small (and cheaper) rocket may be used to launch multiple CubeSats or small satellites. These may be loaded into launchers in a small rocket capable of reaching a Low Earth Orbit (LEO) which then deploys the CubeSats into LEO orbits.

[0007] Compared to more traditional satellite systems, these small and light weight systems often sacrifice system redundancy to achieve a compact low-cost design in order to reduce the barrier of entry. However, whilst these designs have reduced the costs compared to traditional larger satellites, these costs are still significant and a typical 3U CubeSat may still cost hundreds of thousands of dollars to design and launch.

[0008] Thus, it would be desirable to further reduce size and cost. Space is a harsh environment, and hence further reducing size or even maintaining a small size and weight is a challenging design task. Whilst CubeSats may use COTS components and a low redundancy design to minimise cost, size and weight, they must fundamentally be capable of performing mission critical functions such as collection of sufficient solar power to power the satellite (and mission components) as well as being capable of maintaining communications with a ground station in order to transmit mission data and receive control commands. In order to reduce size, complexity and cost, many small satellites do not attempt to actively control their orbit, and in some cases, are designed to automatically de-orbit after some time frame (e.g. 3-5 years) due to natural orbital decay mechanisms such as atmospheric drag effects in LEO orbits. As such they often do not have thrusters or only carry a small amount of propellant and restrict thruster use for emergency conditions such as collision avoidance, rather than for station keeping purposes (e.g. to maintain a particular orbit). We will refer to such satellites as minimally orbitally controlled satellites, which will be understood to include satellites without any orbital control capabilities (i.e. uncontrolled) and / or those designed to allow natural orbital decay. Without intervention such small satellites will typically tumble in orbit (i.e. rotate around each of the three axes). Thus, small satellites typically contain attitude determination and control systems, often referred to simply as attitude control systems, to prevent or control tumble and / or spin (i.e. to allow rotation about a particular axis and prevent rotation about other axes). This is to ensure that solar panels are pointed towards the sun, antennas are capable of communicating with a ground station (e.g. pointed in the direction of the ground station) and if relevant, the payload is pointed in a desired direction (e.g. a camera or directional sensor in an earth observation system. Attitude determination and control systems frequently comprise of always-on reaction wheels, magnetorquers, thrusters for adjusting attitude as well as attitude determination sensors such as star trackers, sun sensors, earth sensors, angular rate sensors, and GPS receivers and antennas.

[0009] It is also a frequent problem with satellite designs that the size and weight can quickly balloon due to positive feedback loops. For example, if the power requirements of a mission component are high, then larger solar panels and batteries are required to power it. This adds additional size and mass, which then makes the satellite more difficult to control, and thus the size of the attitude control systems must be increased by adding further reaction wheels and / or thrusters, which then adds further mass and cost, as well as consuming more power and space, which may then require larger solar panels and batteries. As such design of small satellites whilst reducing cost is a challenging design task.

[0010] There is thus a need to provide improved designs for satellites, including small satellites, or to at least provide a useful alternative to existing satellite designs.SUMMARY

[0011] According to a first aspect, there is provided a satellite comprising: an on-board computer (OBC) comprising at least one processor and at least one memory, a telemetry tracking and command subsystem (TT&C), an electrical and power subsystem, (EPS) comprising one or more batteries, a payload module, an emergency attitude control subsystem (EACS) configured to generate an angular momentum perturbation upon a detection of an undesirable orientation or an undesirable spin state of the satellite; one or more solar panels mounted to one or more sides of the satellite and electrically connected to the one or more batteries; and an RF front end comprising at least two antennas wherein each antenna is connected to a different side of the satellite and the OBC is configured to select a set of one or more antennas to transmit or receive on.

[0012] In a further form, the satellite further comprises an attitude determination system (ADS) comprising one or more attitude sensors to determine a pointing direction of the satellite, and the ADS is configured to detect an undesirable orientation or an undesirable spin state of the satellite and to instruct the EACS to generate an angular momentum perturbation upon the detection. In a further form, the one or more attitude sensors comprises an inertial measurement unit and / or one or more sun sensors on one or more solar panels.

[0013] In one form, the OBC is configured to utilise the at least two antennas based on either: selecting all of the at least two antennas and determining a sequence in which each antenna in the sequence is switched on for an active time period and then off wherein the sequence is either an ordered sequence which is continuously repeated or is a random sequence or is a pseudo random sequence; or grouping the at least two antennas into a set of complementary antenna groups of two or more and determining a group sequence in which each antenna in a group is switched on for an active time period, and then off and continuously cycling through the group sequence; or selecting at least two antennas and simultaneously receiving or transmitting on the at least two antennas; or measuring an orientation performance indicator for each of the at least two antennas and selecting at least one antenna based on the orientation performance indicators such that over an extended time period the selected at least one antenna is used for a longer amount of time than a non-selected antenna or an antenna with a lower orientation performance indicator.

[0014] In a further form, the orientation performance indicator is a performance indicator and the OBC determines a performance weight for each antenna and utilising the one or more antennas is performedsuch that over the extended time period an antenna with more favourable weight is switched on for a longer total time than an antenna with less favourable weight. In one form, the orientation indicator for each antenna is a pointing direction indicator which is a measure of the pointing direction of the antenna with respect to a reference direction determined using the one or more attitude sensors. The order of the sequence or paired sequenced and / or the active time period may be determined based on the weightings such that over an extended time period antennas with favourable weights switched on for longer times than antennas with less favourable weights, and in the case of selecting at least two antennas and simultaneously receiving or transmitting on the at least two antennas the weights are used to select the at least two antennas, and in the case of measuring an orientation performance indicator, the weights are calculated using the orientation performance indicators.

[0015] In one form, selecting at least two antennas and simultaneously receiving or transmitting on the at least two antennas is performed using the coarse pointing direction of each of the at least two antennas determined using the one or more attitude sensors, and the OBC is configured to use beam forming or signal processing of received or transmitted signals using the coarse pointing directions to compensate for effects of a varying pointing direction or tumbling of the satellite.

[0016] In one form, the EACS comprises a DC motor that drives movement of a mass wherein the DC motor is switched on for an activation time period upon detection of an undesirable orientation or an undesirable spin state of the satellite. In a further form, the DC motor does not have a speed controller, and / or one or more of the one or more DC motors drive a shaft, and one or more of the one or more masses are on the shaft, and an axis of the shaft is inclined with respect to at least one side of the satellite on which one or more solar panels are mounted.

[0017] In one form, the EACS comprises a magnetorquer that is activated for an activation time period upon detection of the undesirable orientation or the undesirable spin state of the satellite.

[0018] In one form, the ADS is an attitude determination and control system (ADCS) which is configured to use the one or more the magnetorquers to perform coarse control of attitude to control capture of solar light by the one or more solar panels.

[0019] In one form, the EPS further comprises one or more sensors for determining a battery charge level of the one or more batteries and the EACS is further configured to monitor a battery charge level of the one or more batteries, and an undesirable orientation or an undesirable spin state of the satellite is detected based on analysing a time history of the battery charge level of the one or more batteries and identifying a trend of decreasing battery charge level overtime.

[0020] In one form, the satellite (or component of it) has a rectangular cuboid shape, and each antenna of the at least two antennas is a patch antenna, and each of the at least two antennas is located on a different side of the satellite, and may be collocated with one or more solar panels.

[0021] In one form, the satellite further comprises a shielded enclosure for housing at least the OBC, TT&C, and EACS wherein the shielded enclosure is manufactured of a material configured to reduce the effects of radiation on the components housed with the enclosure, and comprises one or more bulkheads to allow passage of electrical connections through the shielded enclosure. In a further form, the shielded enclosure is a substantially rectangular cuboid housing with a front side, a rear side, a left side, a right side, an upper side and a lower side, and the front side, rear side, left side and right side form four sides of the satellite, and an upper panel is mounted to the upper side of the shielded enclosure and comprises a printed circuit board (PCB) supporting one or more solar panels on an exterior side of the upper panel, and an interior side of the upper panel is divided into two portions, wherein a first portion is mounted to the upper side of the shielded enclosure and is electrically connected to the bulkhead in the upper side of the shielded enclosure, and a second portion extends in a direction normal to the front side and supports one or more solar panels, and a PCB is connected to the lower side of the shielded enclosure and is electrically connected to the bulkhead in the lower side of the shielded enclosure, and supports one or more solar panels, and the at least two antennas comprises a first dipole antenna mounted to the left side and a second dipole antenna mounted to the right side such that when deployed the first and second antennas have an orthogonal configuration.

[0022] In a further form, the left side and the right side each have a stepped profile, and prior to deployment the first dipole antenna and the second dipole antenna are held in a stowed configuration such that they are housed under the panel, and upon deployment, each antenna is released and extends laterally beyond an edge of the panel until they reach a respective stop formed by a step feature in each of the left side and right side.

[0023] In one form, the payload module is integrated into the OBC wherein the payload module comprises a set of instructions stored in the at least one memory of the OBC which are executed by the at least one processor of the OBC to provide payload functionality.

[0024] According to a second aspect, there is provided a deployment system comprising a plurality of satellites of the first aspect wherein the plurality of satellites is dimensioned to form complementary groups of satellites that can be stacked inside a deployment container, and a release mechanism is provided between each satellite in the complementary group of satellites. In a further form the deployment container is a 3U or 6U container.

[0025] In one form, each satellite in the stack is deployed at different points in time to spread their distribution across an orbit.

[0026] In one form, the release mechanism is a spring and bum wire integrated in a PCB forming a solar panel of each satellite.

[0027] According to a third aspect, there is provided a method for operating the satellite of the first aspect, comprising: monitoring for an undesirable orientation or an undesirable spin state of the satellite and upon detection of the undesirable orientation or the undesirable spin state of the satellite, generating, by the EACS, an angular momentum perturbation; and utilising a set of one or more antennas to transmit or receive on.

[0028] In one form, utilising a set of one or more antennas to transmit or receive on comprises: selecting all of the at least two antennas and determining a sequence in which each antenna in the sequence is switched on for an active time period and then off and continuously cycling through the sequence; or grouping the at least two antennas into a set of complementary antenna groups of two or more and determining a group sequence in which each antenna in a group is switched on for an active time period, and then off and continuously cycling through the group sequence; or selecting at least two antennas and simultaneously receiving or transmitting on the at least two antennas; or measuring an orientation performance indicator for each of the at least two antennas and selecting at least one antenna based on the orientation performance indicators such that over an extended time period the selected at least one antenna is used for a longer amount of time than a non-selected antenna or an antenna with a lower orientation performance indicator.

[0029] In one form, the orientation performance indicator is a performance indicator, and the method further comprises: determining a performance weight for each antenna and utilising the one or more antennas is performed such that over an extended time period an antenna with a more favourable performance weight is switched on for a longer total time than an antenna with a less favourable performance weight.

[0030] In one form, the orientation performance indicator for each antenna is a pointing direction indicator which is a measure of the pointing direction of the antenna with respect to a reference direction.

[0031] In one form, selecting at least two antennas and simultaneously receiving or transmitting on the at least two antennas is performed using the coarse pointing direction of each of the at least two antennas,and the method further comprises performing beam forming or signal processing of a received signal or signal to be transmitted using the coarse pointing directions to compensate for effects of a varying pointing direction or tumbling of the satellite.

[0032] In one form, the method further comprises determining a coarse pointing direction of the satellite, and the orientation indicator for each antenna is a time average measure of the pointing direction to the sun based coarse on the pointing direction.

[0033] In one form, detecting an undesirable orientation or an undesirable spin state of the satellite comprises monitoring a battery charge level of the one or more batteries and determining a trend of decreasing battery charge level over time.

[0034] According to a fourth second aspect, there is provided a method for deploying the deployment system of the second aspect, comprising: triggering each release mechanism at different points in time.BRIEF DESCRIPTION OF DRAWINGS

[0035] Embodiments of the present disclosure will be discussed with reference to the accompanying drawings wherein:

[0036] Figure 1 is a functional block diagram of a satellite according to an embodiment;

[0037] Figure 2A is a schematic diagram of a low earth orbit satellite communications system using a satellite according to an embodiment;

[0038] Figure 2B is a schematic diagram of a satellite with an elongated box shape tumbling in orbit and associated variation in the capacity of the batteries according to an embodiment;

[0039] Figure 2C is a schematic diagram of a satellite with an elongated box shape tumbling in orbit in a favourable orientation and spin state and associated variation in the battery charge level of the batteries according to an embodiment;

[0040] Figure 2D is a schematic diagram of a satellite with an elongated box shape tumbling in orbit in an unfavourable orientation and spin state and associated variation in the battery charge level of the batteries according to an embodiment;

[0041] Figure 3 A is a schematic diagram of an emergency attitude control subsystem comprising a mass on a shaft driven by a DC motor according to an embodiment;

[0042] Figure 3B is a schematic diagram of an emergency attitude control subsystem comprising a mass on an inclined shaft driven by a DC motor according to another embodiment;

[0043] Figure 3C is a schematic diagram of an emergency attitude control subsystem comprising two perturbation units each comprising a mass on a shaft driven by a DC motor wherein the two shafts are orthogonal according to an embodiment;

[0044] Figure 3D is a schematic diagram of a satellite tumbling in orbit in an unfavourable orientation and spin state before and after an angular momentum perturbation initiated by the emergency attitude controller and the associated variation in the battery charge level of the batteries according to an embodiment;

[0045] Figure 4 is a schematic diagram of a satellite with a cube shape tumbling in orbit and the associated antenna pairs used for transmitting and receiving according to an embodiment;

[0046] Figure 5A is a perspective view of a satellite with an elongated box shape with two orthogonal antennas in a pre -deployment configuration and which is configured to be stackable according to an embodiment;

[0047] Figure 5B is a second perspective view of a satellite with an elongated box shape with two antennas in a deployed configuration according to an embodiment;

[0048] Figure 5C is a lower view of a satellite with an elongated box shape with two antennas in a deployed configuration and with the rear solar panels and lower panel removed to show the interior of the shielded housing according to an embodiment;

[0049] Figure 5D is a side view of a stack of four of the satellites with an elongated box shape shown in Figure 5A suitable for loading in a 3U canister for deployment according to an embodiment;

[0050] Figure 5E is an end view of a stack of four of the satellites with an elongated box shape shown in Figure 5A suitable for loading in a 3U canister for deployment according to an embodiment;

[0051] Figure 5F is a top view of a stack of four of the satellites with an elongated box shape shown in Figure 5A suitable for loading in a 3U canister for deployment according to an embodiment;

[0052] Figure 6 is a plot of the diversity mode antenna gain (dBi) vs. angle (0) during a tumble in a plane which includes the worst individual antenna gains according to an embodiment;

[0053] Figure 7A is a side view of the structural layout of a stack of 4 of the satellites shown in Figures 5A to 5F loaded into a 3U canister according to an embodiment;

[0054] Figure 7B is a side view of the structural layout of the shielded section shown in Figure 6 according to an embodiment;

[0055] Figure 8A is a plot of the displacement over a time for each of the four of the satellites shown in Figure 5A starting prior to deployment and for a short time frame after deployment according to an embodiment;

[0056] Figure 8B is a plot of the displacement over a time for each of the four of the satellites shown in Figure 5A starting prior to deployment and for a long time frame (longer than the short time frame shown in Figure 8A) after deployment according to an embodiment; and

[0057] Figure 9 is a flow chart of a method for operating a satellite according to an embodiment.

[0058] In the following description, like reference characters designate like or corresponding parts throughout the figures.DESCRIPTION OF EMBODIMENTS

[0059] Referring now to Figure 1 there is shown a functional block diagram of a satellite 10 according to an embodiment. The satellite comprises a shielded enclosure 110 that houses components 140, at least two antennas 120 and one or more solar panels 130. In this embodiment the components 140 include an on-board computer (OBC; 144), an electrical and power subsystem, (EPS; 145), which is connected to one or more solar panels 130, an emergency attitude control subsystem (EACS; 146), a telemetry tracking and command subsystem (TT&C; 147), an RF front end 148, and a payload module 149 (also referred to as simply the payload). Bulkhead connectors 141 are provided through the shielded enclosure 110 to connect components with exterior components located outside of the shielded housing. For example, the RF front end 148 may be connected to the antennas via the bulkhead connectors 141. The components 140 may be mounted on printed circuit boards (PCBs; 142) located within the shielded enclosure 110.The satellite 10 may further comprise an attitude determination system (ADS) or an attitude determination and control system (ADCS) and the EACS 146 may be a subsystem of the ADS / ADCS 146’ . The shielded enclosure 110 may be manufactured of a material configured to reduce the effects of radiation on the components 140 housed with the enclosure. For example, the shielded enclosure 110 may be constructed of 3mm thick Aluminium. However, it will be understood that the material and / or thickness can be varied (increased or decreased) based on the mission parameters such as mission life, orbit, expected radiation environment, and radiation sensitivity of components. In some embodiments theshielding 110 could be omited, for example if radiation hardened components are selected or failure can be tolerated. The payload module 149 may be contained in whole or in part within the shielded enclosure, and may be separate from or integrated with the OBC (e.g. share the processor, memories or other components including the PCB). That is the payload module may be a unitary module, or a distributed module, including a distributed module in which payload functionality is provided by other hardware components and sensors. If a shielded enclosure 110 is included, one or more bulkhead connectors 141 may also be provided to allow passage of electrical connections through the shielded enclosure, or other alternative wiring arrangements can be used to allow signals to pass through the shielded enclosure. The satellite 10 may be a small satellite however it is to be understood that the various features and subsystems described herein may be used on any satellites, and use is not limited to use only on small satellites.

[0060] The OBC 144 comprises at least one processor and at least one memory and acts as main controller for the satellite 10 controlling the EACS, payload and TT&C radio. In some embodiments the OBC may be a low power and / or radiation tolerant / hardened microprocessor or microcontroller, or subsystem, including those manufactured by Microchip (ATMegaS128, SAM family, TSC695, AT697F, etc) Texas Instruments (e.g. MSP430, TMS570, etc), GomSpace (e.g. NanoMind), Space Micro (Proton family), AAC Clyde Space (Kryten-M3, Q7S, Sirius class, etc), Endurosat etc. The OBC may provide various services and perform some of the functionality of the EPS, EACS, TT&C and may also provide services to the payload module depending upon the mission requirements (i.e. may be a distributed part of the payload module).

[0061] The TT&C radio 147 provides a connection between the satellite and a ground or gateway station 30. The TT&C 147 may perform encoding / decoding and modulation and demodulation of data, and may comprise components such as transceivers, Low Noise Amplifier (LNA) and Power Amplifier (PA). Suitable TT&C radios include those manufactured by Planet Labs, AAC Clyde Space, GomSpace, EnduroSat, etc.

[0062] The RF front end 148 comprises at least two antennas 120. Each antenna is connected to a different side of the satellite and the OBC is configured to select a set of one or more antennas to transmit or receive on. The RF front end acts a front end of the TT&C radio, and may also be used for the payload module (if required; i.e., as a distributed part of the payload module). Suitable antennas include patch antennas, omnidirectional antennas and directional antennas including those manufactured by GOMSPACE (e.g. NanoCom ANT2000, NanoCom AM2150-O), Clyde Space (ANT-100), ISISPACE, Endurosat, etc.

[0063] The electrical and power subsystem, (EPS), 145 comprises one or more bateries and one or more sensors for determining a batery charge level of the one or more bateries. In one embodiment the sensorsare voltage sensors for measuring a voltage level of the batteries which is used to estimate the battery charge level, or an equivalent measure such as the available capacity of the battery or the state of charge defined as the available capacity / total capacity. Other sensing arrangements may also be used to estimate the battery charge level including sensors to measure the battery current and / or measuring battery current and voltage. The EPS provides power to the entire satellite via the solar panels and batteries enabling sufficient power to run the satellite when in sunlight and in eclipse. The EPS may include current and voltage sensors to monitor voltage and current of the batteries and solar panels, and may be used to estimate the battery charge level or equivalently the remaining available capacity of the battery or state of charge. The satellite may be operated in different operational modes based on the remaining battery capacity.

[0064] The emergency attitude control subsystem (EACS) 146 is configured to generate an angular momentum perturbation upon detection of an undesirable orientation or an undesirable spin state of the satellite. In some embodiments the satellite may comprise an attitude determination system (ADS) comprising one or more attitude determination sensors used to determine a pointing direction of the satellite. In some embodiments the attitude determination system (ADS) is an attitude determination and control system (ADCS) further comprising components (e.g. magnetorquers) to perform attitude control, including coarse attitude control during normal flight operations. The EACS may be configured to perform the detection of the undesirable orientation or an undesirable spin state, or the detection may be performed by other system components, either as a service for the EACS or the EACS may be under the control of the OBC, ADS or ADCS which performs the detection and then instructs, or triggers, the EACS to generate the angular momentum perturbation. Detection may be performed using one or more attitude determination sensors, e.g. of the ADS / ADCS and / or by monitoring one or more systems or components (e.g. battery charge level of EPS). In one embodiment the attitude determination sensors of the ADS / ADCS comprises one or more inertial measurement units (IMU) and / or one or more coarse sun sensors on solar panels for attitude determination which are used to perform the detection of the undesirable orientation or spin state, and may also be used for general attitude determination in normal flight. For example the sun sensor may comprise a single sensor such as a photodiode or photodetector whose output signal is measured, or multiple photodiodes or photodetectors may be arranged as a linear array, orthogonal linear arrays or grid arrangements. In other embodiments entire individual solar panels may be used as a coarse sun sensors by monitoring the output of each solar panel, or the comparative output of different solar channels. As will be discussed below generating an angular momentum perturbation by the EACS may be performed by one or more perturbation apparatus such as a basic DC motor (e.g. optionally without requiring a speed controller) driving movement of a mass (rotationally or linearly) or by magnetorquers if present. In one embodiment the OBC provides a simple low-rate determination and control service to perform the detection, and the satellite does not rely on or require on accurate pointing or fine attitude control (e.g., is allowed to tumble). The EACS can also be used as anend-of-life / disposal mechanism by generating angular momentum perturbations as part of end-of-life operations to disturb the drag state and increase drag to induce de-orbiting. The EACS may be used on small satellites and may be used to in place of more complex, heavy or costly attitude control systems. However, it is to be understood that use of the EACS as described herein may also be used on any satellite, for example to provide redundancy in case more complex systems fail.

[0065] The payload module 149 may be any suitable payload depending upon the mission requirements. For example, in the case of a communication system the payload module may comprise a payload computer combined with one or more transceivers to form a Software Defined Radio (SDR) to provide access node functionality in the communication system. The communication system may be an embodiment of the communication systems described in PCT / AU2013 / 001078 titled COMMUNICATION SYSTEM AND METHOD and the satellite may be configured as an access node. In this embodiment, when receiving the RF transceivers pass IQ samples to the SDR for demodulation and decoding. For transmitting the SDR generates and sends IQ samples to one of the RF transceivers for transmission via the RF front end 148 and antennas 120. The antennas 120 may be configured to transmit on VHF, UHF, L and S bands and may be omnidirectional and / or patch antennas. In other embodiments, the payload module 149 may be a payload computer combined with a sensor, such as a camera, spectrometer, optical or RF or magnetic sensing apparatus. The OBC may be configured to boot, wake from sleep, shutdown and reboot the payload computer. In some embodiments the payload module comprises a payload computer in addition to the OBC, i.e. separate processor(s) and memory, and may be a physically separate board or a separate physical module to the OBC. In some embodiments the payload module may be integrated into the OBC, and share the same processor, memory or other components including the PCB as the OBC. For example, the payload module may comprise one or more processors and one or more memories located on the same PCB as the processor and memories of the OBC. In some embodiments the payload module may comprise a set of instructions stored in the OBC memory (or memories) which are then executed by the OBC processor(s) to perform payload functionality (e.g., communications processing). The OBC may provide a virtual server environment to run the payload module, and may be isolated or sandboxed to prevent the payload module interfering with mission critical functions.

[0066] Figure 2A is a schematic diagram of a low earth orbit satellite communications system using a satellite 10 as an access node according to an embodiment. The system 1 may be used to provide Internet of Things (loT) connectivity with remotely located sensors and assets 22 for a plurality of users 42. The satellite communications system 1 is comprised of one or more satellites 10 in a low earth orbit (LEO) 12 which acts as an access node for a plurality of terminals 20. The satellite 10 communicates with a plurality of geographically distributed gateways 30, such as ground stations, which are in communication with core network infrastructure 40. In some embodiments the satellites perform on-board processing ofreceived transmissions including decoding and processing of received packets, and may perform on-board processing to reduce the amount of data sent over a downlink. In some embodiments satellites may be configured to digitally capture the transmission and forward the digital capture over the downlink for decoding and processing by network infrastructure. The satellite 10 has a field of view 14 comprising multiple ground-based terminals 20 with communicate with the satellite 10 using radio frequency signals over a radio link 16, which may be a unidirectional downlink or a bidirectional link, and are connected to, mounted on, or in communication with sensors or assets 22. The satellite 10 communicate with a gateway 30 using radio frequency signals over a radio link 18, which may act as a downlink or uplink, e.g. to transmit commands and data to the satellite 10. The core network infrastructure 40 may include cloudbased servers which may manage the system and network routing, and provide an application interface to forward data to and from the terminals to a plurality of users 42 or otherwise provide an interface to allow users to access data provided by terminals 20. Multiple users may be supported, each communicating with different terminals 20. The gateway may send data and commands to the satellite 10 over radio link 18. In other embodiments a satellite may be used in an earth observation system which omits user terminals 20. In these embodiments the satellite 10 comprises a payload module 149 that collects observational data in the field of view 14, such as radio-based measurements and optical images, which is sent to a gateway 30 over radio link 18. In some embodiments the payload module 149 performs space-based sensing such as measurements of magnetic fields or radiation data, such as for space weather applications.

[0067] In this embodiment the satellite 10 is a minimally orbitally controlled satellite and is shown tumbling in orbit. Thus, in this embodiment, to ensure sufficient power for the satellite each side is covered in solar panels. This ensures that at any point in time the satellite is receiving light and charging the batteries to power the satellite. However, in other embodiments fewer sides may be fitted with solar panels, including fitting only one side with solar panels. As the number of solar panels is reduced, the duty cycle may also be reduced to compensate for the reduced available power or recharging capabilities. In this embodiment, in order to maintain continuous communications with the gateway and terminals in the field of view, each side of the satellite comprises a patch antenna collocated with a solar panel, which are mounted on a PCB board. In other embodiments, rather than placing patch antennas on all sides, fewer but still multiple patch antennas may be used, where each patch antenna is located on a different side of the satellite (but in this embodiment not on all sides). However, in other embodiments alternative antenna types, e.g. wire dipole or tape measure antennas, may be located on multiple or each edge or comer, and transmission and reception may be performed jointly on all antennas, or combinations of antennas as discussed below. In some embodiments a pair or orthogonally orientated dipole antennas are used.

[0068] Figure 2B shows another embodiment of a satellite with an elongated box shape tumbling in orbit and the associated variation in the battery charge level (C) of the batteries, which may also be referred toas the power level, state of charge, or remaining capacity (of the total capacity) of the battery, according to an embodiment. In this embodiment the satellite has a panel like shape in which the upper and lower sides have an area much larger than the area of the side walls. The largest dimension is the length followed by the width, both of which are substantially larger than the depth. As discussed below this would allow packing (stacking) of multiple satellites into a launcher such as a standard 3U or 6U launcher. The solar panels are mounted on upper and lower sides. In this embodiment the front, rear, first (left) and second (right) sides of the satellite form a shielded enclosure and house the electronics components (OBC, TT&C, EPS, batteries, etc) within the shielded enclosure. The satellite comprises two antennas - one each mounted to the left (first) side and right (second) side (i.e., on opposite sides). During deployment the antennas extend outward to an angle of 135° with respect to the opposing sides so the two antennas become orthogonally directed when deployed.

[0069] In this embodiment the satellite is tumbling along orbit 12 and as the solar panels are located on two opposing sides, the amount of light collected, and power generated will vary over time as shown in Figure 2B. For example, in Figure 2B the satellite 10 is shown in a first orientation 13, a second orientation 13 ’ and a third orientation 13 ”. In the first orientation 13 both solar panels are orientated almost orthogonally to the sun and thus are generating very little power. In the second orientation 13’ the upper solar panels are directed towards the sun and thus generates substantial power. Similarly in the third orientation 13’ the lower panels are directed toward the sun and thus generates substantial power. In some embodiments solar panels may be located on at least two sides so that as the satellite 10 tumbles along orbit 12, at least one panel will typically be viewing the sun at any point in time ensuring sufficient power can be collected to charge the batteries and power the mission.

[0070] Typically, the tumbling will be random so that on average at least one solar panel will be viewing the sun most of the time so that the battery charge level of the battery remains in a safe range where there is sufficient power to support all mission activities. As discussed above the battery charge level of the battery can be determined using a voltage sensor and appropriate voltage thresholds determined for safe / normal operations (e.g. Vbat> Vsafe). However, in some cases the satellite can preferentially tumble or spin about one or more axes and conservation of angular momentum will thus bias the satellite into an orientation about a particular spin axis or axes. In some cases, this can be beneficial such as shown in Figure 2C, where the satellite is spinning around the normal axis of the upper solar panel which is pointed approximately towards the sun 11 allowing the batteries to charge and potentially reach a maximum capacity of the battery (i.e. fully charged). In some embodiments the satellite may be deliberately placed into an advantageous spin state, for example by the use of magnetorquers, where the satellite spins about a preferred axis such as to ensure charging of solar panels, or to orient the antennas in a desired reference direction such as towards the earth.

[0071] However, in other cases, such preferential tumbling or spinning about a particular axis can be detrimental, as shown in Figure 2D, where the satellite is spinning around the normal axis of the upper solar panel which is pointed approximately normal from a direction towards the sun 11. In these orientations the solar panels generate very little (if any) power and thus the battery charge level of the batteries decreases over time as mission functions consumer power (i.e. remaining capacity / state of charge decreases over time).

[0072] Thus, to provide some robustness against cases where the satellite gets stuck in an undesirable orientation or an undesirable spin state the satellite is configured to detect an undesirable orientation, or an undesirable spin state of the satellite and an emergency attitude control subsystem (EACS) is configured to generate an angular momentum perturbation upon the detection of the undesirable orientation or undesirable spin state of the satellite. The angular momentum perturbation may be induced by the temporary or short duration use of a magnetorquer or a DC motor to drive movement of a mass. In some embodiments the mass could be mounted on shaft driven by the DC motor thus forming a rudimentary reaction wheel (circular motion of the mass). The axis of the shaft may be inclined with respect to at least one side of the satellite on which one or more solar panels are mounted. That is the shaft could be orientated at an angle with respect to the side walls or the plane containing the most solar panels, to simultaneously induce spin on two or more axes (of the satellites). Multiple shafts and masses may be used, with each shaft orientated in a different direction. Other arrangements could be used such as using the DC motor to drive a cable or gear arrangement which drives the mass to move backwards and forwards along a track (linear motion of the mass). Choice of mass and specific arrangement could be chosen based on the specific dimensions and available space within the enclosure.

[0073] Figure 3A is a schematic diagram of an emergency attitude control subsystem (EACS) showing a first embodiment of a perturbation apparatus 300 comprising a DC motor 302 without a speed controller that drives rotation of a mass 304 via a shaft 306. The DC motor is switched on for an activation time period upon detection of an undesirable orientation or an undesirable spin state of the satellite. However, in other embodiments a speed controller, including a coarse speed controller may be included. Figure 3B shows a second embodiment of a perturbation apparatus 310 in which the motor and shaft are mounted on an 45° angle via a mount 312 with respect to two sides of the satellites or surfaces within the satellite (e.g. a PCB board and a side wall) to induce spin on two axes of the satellite. The mounting could also mount the motor at other angles, such as an angle with respect to each of the three principal axes of the satellites, and thus induce spin on all three axes. For example, the motor could be mounted in a comer and at an 45° angle with respect to each of the three sides forming the comer (and thus the 3 principal axes). In another example the mount may be mounted on a single surface such as PCB board, or other component, and support the motor in an orientation such that is angled with respect to two or three of the principal axes of the satellite. It will be understood that the motor could also be mounted at any angle (e.g. 10°, 20°, 30°,40°) to induce uneven (i.e. unbalanced) perturbations on the different axes. The mount may be constructed using struts, tooled or machined parts, be 3D printed, and be designed to support the expected loads and torque to maintain the motor in a fixed position during operation. Figure 3C is a third embodiment comprising two perturbation apparatus 320 330, in which the two-perturbation apparatus are configured such that the shafts are orthogonal to allow generation of perturbations in two or more axes (depending upon the alignment with the satellites axes). The activation time period 340 will typically be short - from a few seconds to tens of seconds. This is illustrated in Figure 3D which is a schematic diagram of a satellite tumbling in orbit in an unfavourable orientation and spin state, before and after an angular momentum perturbation initiated by the emergency attitude controller, and the associated variation in battery charge level of batteries according to an embodiment. In orientation 13 the satellite is spinning in an unfavourable orientation with the solar panels pointed orthogonally to the sun and spinning about the normal to the upper side. This unfavourable state continues as the satellite moves along the orbit, until, for example orientation 13’, at which point the battery charge level (i.e. remaining capacity) of the batteries drops below a threshold level triggering activation of the DC motor 302 for activation time period 340. This generates an angular momentum perturbation (or kick) such that the satellite orientation changes so that the solar panels are now able to collect sunlight and thus recharge the batteries. This is illustrated in third orientation 13” in which the upper solar panel is now approximately pointing towards the sun.

[0074] Reaction wheels system used by typical satellite attitude controllers (including small satellite attitude controllers) are typically always on and feature tight speed control to enable fine control over the satellite attitude and pointing direction of antenna or payload instruments. These features are not required in the current case where no or only coarse attitude control is required, and inclusion of speed control is optional. As such, in some embodiment the speed controller can be omitted to save weight, volume and substantially reduce complexity. However, in some embodiments the speed controller may be included, for example based upon the specific mission requirements and constraints, or ease of integration. In some embodiments a relatively coarse (low cost / complexity) speed controller could be included allowing use of a cheaper or less complex module compared to typical reaction wheel systems. Further as it is only required for emergency situations and is only used for short durations (rather than being always on) the power consumption is also very low. The mass can also be relatively small as it just has to be sufficient to induce a perturbation or change in angular momentum, and thus can be smaller than typically attitude controllers which are required to perform fine attitude control. The mass and motor can be housed within the shielded enclosure (if present). Further when using the motor to exit an unfavourable orientation state there is no requirement to select the direction or amount of change, as the satellite is already in an unfavourable orientation, and thus any other orientation is likely to be better. Further the system can continue to monitor the battery charge level after the perturbation, and if there is insufficient improvement in battery charge level, then an additional angular momentum perturbation could beperformed. That is the motor could be turned on for several short bursts over the space of a few minutes until sufficient recovery is observed. In some embodiments the EACS could be used to drive the satellites into a favourable orientation or spin state, and once this is attained, the EACS may then be used as required to maintain the favourable orientation or spin state. This may be performed in an iterative process in which the EACS is activated for an activation period, and then coarse direction or charge state monitored over a monitoring period to determine the effect. This activation / monitoring can be repeated as required until a favourable orientation or spin state is achieved. The EACS can then be used again later if the satellites drift out of the favourable orientation or spin state into an unfavourable orientation or spin state. Embodiments of the EACS as described herein may thus be used in small satellites, and or to enable construction of a small satellite. However, it is to be understood that embodiments of the EACS could be used in any satellites to save weight or cost, or where no or coarse attitude control is required. The choice of the mass and the DC motor (and number) may be selected based on the specific mission requirements and constraints.

[0075] Additionally, or alternatively the EACS may comprise one or more magnetorquers. In some embodiments the magnetorquers may also be used for attitude control as part of an ADCS system (e.g. where the EACS is a subsystem of the ADCS system). These may be activated for an activation time period upon detection of an undesirable orientation or an undesirable spin state of the satellite to generate an angular momentum perturbation. The magnetorquer may be located within the shielded enclosure, satellite housing, or located on a solar panel. As the perturbation effect of a magnetorquer is reliant on the strength of the existing magnetic field, the use of magnetorquers is only feasible for orbits where the magnetic field is of a sufficient strength, for example in LEO orbits. Solar storms can also affect the strength of the magnetic field. Thus, in some embodiments, such as non-LEO orbits including moon orbits, a basic DC motor and mass arrangement may be used in place of a magnetorquer, or alongside a magnetorquer as supplementary system for use in low magnetic field environments, or simply as a backup system in case of failure of the magnetorquer).

[0076] Detection of an undesirable orientation or an undesirable spin state of the satellite may be performed by monitoring and / or analysing sensor data from one or more sensors. In one embodiment detection is performed by analysing a time history of the battery charge level, or equivalently the state of charge or available capacity, of the one or more batteries and identifying a trend of decreasing battery charge level or state of charge over time. The battery charge level may be converted into another measure such as state of charge or available capacity. The state of charge is a ratio of the available capacity to the total capacity of the battery at the time of measurement. State of charge may also be measured using an alternative. Thus, if the satellite is using more power than it is collecting then the available capacity and state of charge will decline overtime, for example by using a time averaged value. The monitoring maybe performed by the OBC, accessing the EPS sensors, and provided as a service to the EACS. In oneembodiment the battery charge level is periodically sampled (e.g. by taking a measurement of the battery voltage and a moving average or a similar central measure of the distribution of samples (e.g. median, trimmed mean, etc) calculated over a moving time scale (e.g. 10 samples would be lOseconds if sampled at 1Hz). Other approaches such as fitting a linear or low order polynomial curve (e.g., 2ndor third order) to the data could be performed and a parameter determined indicating a trend of decreasing battery charge level, or equivalently the available capacity or state of charge, such as negative slope. Quality checks could be performed such as calculating a measure of variance (e.g. standard deviation or interquartile range), or calculating the coefficient of variation, or otherwise analysing the residuals to protect against false positive detections. However it is noted that in most cases a false positive will only use a small amount of power and lead to a reorientation which is likely to be similarly favourable from a power collection perspective and unlikely to drive the satellite into an unfavourable orientation (although even that would be detected by ongoing monitoring) Alternatively, the undesirable orientation or an undesirable spin state may be detected using a voltage threshold such as Vsafe and / or Vcritical (where Vsafe > Vcritical). In another embodiment the monitoring may comprise defining, determining or setting a threshold and then detecting if the battery voltage is less than the threshold. For example, the battery voltage could be sampled at some predetermined rate (1Hz) and a watchdog timer could be reset if the battery level remains above a threshold level (e.g. Vsafe and / or Vcritical). Thus, when the battery level drops below the threshold the watchdog timer will start counting down. If the battery voltage stays below this threshold for the watchdog time, then the watchdog timer will count down to zero triggering activation of the EACS (e.g. perturbation apparatus and / or magnetorquer) for the activation time. In some embodiments determining a trend or time monitoring of decreasing power may only be performed once the power level has dropped below a first threshold such as Vsafe, and which point the satellite switches to a safe operational mode switching off or restricting use of non-essential services (e.g., payload functionality). A second threshold Vcritical (Vcritical < Vsafe) is then used to trigger the angular momentum change set (i.e. the detection criteria is Vbat < Vcritical). The satellite may also be switched into a critical operational mode when the with only essential services provided (e.g. OBS, TT&C) on passing this second threshold. In some embodiments the second threshold voltage could be a voltage below Vcritical. That is, Vcritical is just used to switch operational modes, or the trigger criteria for a detection could be at least a minimum time duration during which the voltage remains below the Vcritical threshold. This duration could be any value in the range from minutes to many hours, thus reserving use of an angular momentum change (or kick) for emergency situations, e.g., when even shutting down to essential services is insufficient to enable the battery charge level to recover. The exact duration could be predetermined based on the specific mission (e.g. how long can the satellite be out of service for), the remaining battery capacity, the maximal power generation rate of solar panels, and / or the power consumption of essential services. The predetermined duration could be amended based on power usage once in flight. More complex triggering criteria for a detection could also be used, including multiple trigger criteria such as use of absolute threshold values and rates of power usage, historical analysis ofbatery charge data including time series analysis, and machine learning based approaches. For example, a classifier could be provided with the time history of batery charge data and a model trained to classify the likelihood the satellite is in an undesirable orientation, and once trained, generate a likelihood or trigger decision based on use of the classifier on the most recent time series of batery capacity measurements.

[0077] In some embodiments the EACS may also include atitude sensors (e.g. IMU and / or sun sensors), or the satellite may include an ADS that comprises atitude sensors, or ADCS that comprises atitude sensors and magnetorquers and performs at least coarse atitude control. Coarse atitude control may be provided by magnetorquers to try and keep the solar panels mostly directed towards the sun with the EACS used in emergency situations. The ADCS may be configured to use the magnetorquers to perform coarse control of atitude to control capture of solar light by the one or more solar panels. For example, the ADCS may operate a control loop and use the magnetorquers to try and maintain solar capture (or charge level) above a threshold amount, or the ADCS may atempt to improve or optimise solar capture by using the magnetorquers to keep solar panels pointing towards the sun, including to maximise the number or area of solar panels pointing towards the sun. The attitude sensors may also be used to determine a coarse orientation of the satellite, and thus a coarse pointing direction for one or more of the solar panels, or sides of the satellite. Detection of undesirable orientation or the undesirable spin state of the satellite may be performed by analysing the time series of pointing directions of one or more sides of the satellite, for example with respect to a reference direction such as towards the sun or towards the earth. For example, if the solar panels are located on two sides such an upper side and a lower side, the average pointing direction of each side, and or the variation in pointing direction could be analysed. A detection of an undesirable orientation or an undesirable spin state could be determined by comparing the percentage of time over some measurement time period that each or either of the sides (or solar panels) are pointed towards a reference direction such as the sun (which can be determined based on knowledge of the orbital location). A threshold angular range, such as the pointing direction with 30 degrees of the direction to the sun, could be used to determine if the side is pointing towards the sun. The detection criteria could be defined as requiring that, over a monitoring duration, such as 24 hours or the last 20 orbital passes, the solar panels were pointed towards the sun (e.g. based on with 30degrees) for less than a minimum value. The minimum value could be a percentage such as 1% or 5% or it could be a duration of time, and could be determined based on the satellite specifics (e.g. typical power consumption and maximum possible charging rates). Other more complex criteria could be used based on further analysis of the pointing directions of the solar panels or satellite sides, including rates of change or variations in the pointing direction which may indicate tumbling, and whether tumbling is about a specific axis or axes which may indicate the solar panels receiving insufficient solar radiation.

[0078] Complex trigger criteria for determining or triggering detection of an undesirable orientation or an undesirable spin state could be determined based on analysis of pointing directions and battery charge levels as discussed above. The detection, including monitoring and analysing of sensor data or performance characteristics may be performed solely by the EACS subsystem, or the detection may be provided as a service to the EACS by another subsystem, such as the OBC, or the EACS subsystem may be a distributed system in which monitoring and analysis is performed by one or more of the OBC, TT&C and ACS subsystems. Multiple criteria could be defined to indicate that the satellite has entered an undesirable orientation or an undesirable spin state of the satellite, with each criteria defined in relation to a specific sensor, a specific performance characteristic, or a specific combination of sensors and / or performance characteristics. The monitoring tasks of each sensor may be distributed through the different satellite subsystems, wherein on triggering of a detection criteria, a detection message is sent to the EACS which then generates the angular momentum perturbation. The OBC may also be configured to monitor the different sensors and determine when a triggering condition is met (i.e. a detection) and then send a detection message to the EACS to generates the angular momentum perturbation. The detection message may specify the activation time for the angular momentum perturbation.

[0079] Another consideration for improving the design of satellites, including small satellites, is to use multiple antennas on different sides and to control, for example by the OBC, which antennas are selected for transmission and reception, or transmission and reception may be performed on multiple antennas, and beamforming may be performed based on orientation information, such as that provided by the ADS or ADCS. Selecting which of the one or more antennas may be performed in various ways, and several of the following ways may be combined.

[0080] In one embodiment, all of the at least two antennas are selected, and a sequence is determined in which each antenna in the sequence is switched on for an active time period and then off (until next switched on). In this embodiment there is thus a simple ping-pong from one antenna to another in order to improve the on-average performance. The sequence maybe an ordered sequence which is continuously repeated, or it may be a random sequence or a pseudo random sequence. The OBC may be used to select the type of sequence and to control cycling through an ordered sequence. The ordering in an order sequence may be updated and changed at a later time. For example, in the case of antennas on each side of the satellite the ordered sequence could be a linear sequence (1, 2, 3, 4, 5, 6, 1, 2, 3, 4, 5, 6, . . . ). In the case of a random sequence, e.g. (1, 5, 2, 3, 5, 4, 1, 6 . . .), the sequence could be pre-generated as a long list which is repeated, or it may be comprised of a series of random sub sequences of a fixed length with each new random subsequence generated in real time (e.g. as the previous subsequence is reaching the end) or the random sequence may be generated in real time. e.g. a random number generator polled at each change time point. The amount of time each antenna is switched on may be a fixed or predetermine time period, or it may vary. For example, an ordered sequence may be repeated, but the duration that eachantenna in the ordered sequence is switched on for, i.e. the active time, is randomly varied. This active time duration may be a time period randomly selected between a minimum time, which could be zero, and a maximum time.

[0081] In another variation we can group antennas into a set of complementary antenna groups and determine a group sequence in which each antenna in the group is switched on for the active time period, and then switched off. Like the previous case, the sequence of groups can be an ordered sequence that is repeated so that the system continuously cycles through the group sequence, or the sequence of groups may be a random or pseudo random sequence. The groups may be pairs of antennas on adjacent surfaces, so they are orthogonal and thus have complementary antenna patterns. This is illustrated in Figure 4 in which the satellite is a cube with antennas on each surface. In this embodiment we number each surface from 1 to 6 (see Figure 4) and form orthogonal pairs (1,2), (1,4), and (3,5). In Figure 4, the satellite is tumbling. At time ti, we use pair (1,2), and at time t2, we use pair (1,4), and at time t3, we use pair (3,5). The use of orthogonal pairs means that as the satellite tumbles, it is likely that at least one or both of the pairs are directed towards the reference direction (e.g. a point on the Earth’s surface) or at least have a reasonably favourable orientation for communications (e.g. toward the Earth). In some embodiments a satellite may include multiple antennas which operate at different frequencies, and multiple antennas may be located on the same side, and the groups may comprise three or more antennas, which may operate at the same frequency band or in different frequency bands.

[0082] When selecting two or more antennas (e.g. using when using antenna groups) we may simultaneously receive or transmit on the at selected antennas and may use beam forming and signal processing techniques. This allows the satellite to exploit diversity generated by the tumbling of the satellite. The signals may be combined or beamformed based on the pointing directions of each antenna, for example as determined by the ACS / ADCS, or by using sun sensors or output of solar panels adjacent or co-located with the antenna. In some embodiments all antennas may be selected and signal combining, including beamforming, performed on all received or transmitted signals. In some embodiments we may measure an orientation performance indicator for each of the antennas. In one embodiment the orientation performance indicator is a receive performance indicator including SNR, PER, RSSI, etc. In another embodiment the orientation performance indicator is a pointing direction indicator which is a measure of the pointing direction of the antenna with respect to a reference direction. The reference direction may be toward the sun, toward the Earth, or a point on the Earth’s surface. In some embodiments the reference direction may be determined indirectly, for example by determining the direction to a first direction, such as the sun, and then the offset to the actual desired direction (the reference direction) such as a point on the Earth’s surface, which can be determined using the known locations of the Earth, sun, and orbit of the satellite. The pointing direction may be a coarse pointing direction determined by monitoring the relative solar power captured by each side or each solar panel which may indicate the instantaneous or a timeaveraged pointing direction of the solar panel and an associated antenna which is adjacent the solar panel or on the same side of the satellite as the solar panel. In another form the pointing direction of the antenna may be determined through the use of the attitude determination system (ADS). This may be used to determine the pointing direction of each antenna, or the time averaged pointing direction to determine which antenna or antennas, are, on average, pointing in the desired direction, and thus can be switched on more frequently. We can then select at least one antenna based on the orientation performance indicators. This may include continuously switching to the antenna with the best receive performance indicator at some rate (e.g. 1 Hz, l / 60Hz, etc). The rate may be based on an estimated rate of tumbling or spinning. In the above cases where we receive on two or more antennas, the received signals could be compared based on the performance metrics and the antenna with the best performance metric selected. This could be performed prior to decoding of the signal or after decoding of the signal. In the latter case the performance metrics may include decoding metrics. In other embodiments, the signals from multiple antennas including antenna groups could be combined such as by using maximal ratio combining or other beam forming and signal processing methods. In some embodiments each of the antennas is independently processed to obtain the set of all decoded packets. When choosing transmit antennas, antenna pointing metrics, or recent receive performance metrics may be used to select the transmit antennas.

[0083] In another embodiment we may bias the selection of antennas or the duration of time a selected antenna is switched on for. This may be determined using the orientation performance indicators including receive performance indicators, antenna pointing directions with respect to a reference direction as determined from the ADS / ADCS, or other relevant factors, for example based on the output of a solar panel, or panels adjacent or co-located with the antenna which can provide coarse attitude determination. In these embodiments (or variations) an antenna with a favourable orientation performance indicator will, over an extended time period, be used for a longer amount of time than an antenna with a lower orientation performance indicator. As described above, orientation performance indicators such as receiver performance indicators or average pointing direction with respect to a reference direction may be calculated. In one embodiment the average amount of power generated by a set of panels on each side of the satellite may be calculated to determine which side or sides are more often pointing towards the sun. We can thus determine an average pointing direction for each side and estimate an orientation metric that measures how often each antenna is directed towards the reference direction e.g. towards the surface of the earth or a specific point on the earth’s surface. In other embodiments we may estimate the pointing direction of each antenna using the attitude estimated by the ADS / ADCS, and this can be averaged over some time period to determine an average pointing direction, which can then be used to obtain an orientation metric that measures how often each antenna is directed towards the reference direction. The orientation metrics (e.g. average pointing directions) may then be used to select how often each antenna isswitched on such that the antennas which are on average directed in the reference direction (or a favourable direction) are switched on more often than those directed in less favourable directions.

[0084] In some embodiments the OBC may determine a performance weight for each antenna and selecting the antennas or duration of time that one or more antennas is selected for may be performed using the weights such that over an extended time period an antenna with more favourable weight is switched on for a longer total time than an antenna with less favourable weight. Thus the order of the sequence or group sequence and / or the active time period may be determined based on the weightings such that over an extended time period antennas with favourable weights are switched on for longer times than antennas with less favourable weights, and in the case of selecting at least two antennas and simultaneously receiving or transmitting on the at least two antennas the weights are used to select the at least two antennas, and in the case of measuring a receive performance indicator, the weights maybe calculated using the receive performance indicators.

[0085] In embodiments where the satellite comprises attitude determination sensors knowledge of the attitude could be used to select or weight the most favourable antenna (based on attitude / pointing direction) or pair of favourable antennas to transmit or receive on. Thus, the OBC or transceiver system may be configured to use beamforming or signal processing techniques on received or transmitted signals to compensate for effects of a varying (e.g. poor) pointing direction or tumbling of the satellite.

[0086] We thus have a method of operating a satellite that ensures sufficient power and communications is maintained. Figure 9 is a flow chart of a method 200 for operating a satellite according to an embodiment. The method comprises monitoring 210 for an undesirable orientation or an undesirable spin state of the satellite battery charge level as described above, and generating an angular momentum perturbation using the EACS upon detection of an undesirable orientation or an undesirable spin state of the satellite. Additionally, the method comprises utilising, 220, a set of one or more antennas to transmit or receive on as described above.

[0087] In some embodiments the satellite is designed to be stackable, so that multiple satellites may be grouped together, and then the groups stacked and loaded into, and later deployed from a launcher (also known as a deployer), such as a standard 3U or 6U launcher. Figure 5A is a perspective view of a satellite with an elongated box shape with two antennas 121 122, shown in a stowed configuration and which is configured to be stackable according to an embodiment. Figure 5B is a second perspective view of the satellite shown in Figure 5 A with the two antennas 121 122 shown in a deployed configuration such that they are orthogonally orientated with respect to each other. The satellite has an upper panel 111, a lower panel 112, a front side 113, a rear side 114, a first side 115 (which is the left side when viewed from the front) and a second side 116 (which is the right side when viewed from the front). The upper surface of the upper panel 111 is covered by a first set of solar panels 131 and forms the upper surface of thesatellite. The lower surface of the satellite is substantially formed into two sections (or portions) - the first section being formed by the lower panel 112 over the shielded housing 110, and the second section formed by the lower surface of the upper panel 111. A second set of solar panels 132 are located on the lower surface - four on the first section (i.e., on lower panel 112) and four more on the second section (lower surface of the upper panel 111). The upper panel 111 and lower panel 112 may be formed as PCB, for example an FR-4 PCB, and the solar panels may be mounted to the PCB panels, or formed as an integral panel. The dimensions of the upper panel 111 and lower panel 112 are selected to fit the launcher dimensions, for example the width and length of the upper panel 111 is selected to fit within a 3 U launcher.

[0088] Figure 5C is a lower view of the satellite shown in Figure 5A and 5B with the two antennas in a deployed configuration and with the lower panel 112 and lower surface of the shielded housing 110 removed to show the interior of the shielded housing 110, along with the solar panels 132 forming the second section of the lower surface 112. Figure 5D is a side view of a stack of four of the satellites with an elongated box shape show in in Figure 5A suitable for loading in a 3U canister for deployment according to an embodiment. Figure 5E is an end view of a stack of four of the satellites with an elongated box shape show in in Figure 5A, and Figure 5F is top view of a stack of four of the satellites with an elongated box shape show in in Figure 5A.

[0089] Prior to deployment the first dipole antenna 121 and the second dipole antenna 122 are held in a stowed configuration such that they are housed under the upper panel 111 (Figures 5 A, 5D and 5E). In this embodiment the antenna mount comprises a spring biasing the antenna to rotate away from the side / stowed position one a release mechanism such as a bum wire is activated once the satellite is deployed. Upon deployment, each antenna is released (e.g. release mechanism is activated) allowing the antennas to extend laterally beyond the edge of the panel (when viewed from above) until they reach a respective stop formed by a step feature 117 in each of the left side 115 and right side 116 (see Figures 5B and 5C). This acts as a physical stop and limits the maximum angle the antenna mount can rotate to. In this embodiment the maximum angle is 135° (with respect to the side) so that when deployed the two antennas will adopt a 90° / orthogonal orientation. However in other embodiments different angles could be used as well as other antenna types including tape-measure type extending omnidirectional antennas, helical antennas (deployable spring) and patch antennas. In addition to the two dipole antennas 121 and 121, a GNSS patch antenna 123 is mounted on the rear side 114 for connection to a GNSS chip.

[0090] In this embodiment the shielded enclosure 110 is a substantially rectangular cuboid housing in which the front, rear, left, and right sides substantially form the front 113, rear 114, left 115 and right 116 sides of the satellite. The shielded enclosure also includes an upper side and a lower side. Figure 5C shows an interior view of the shielded enclosure 110 that houses components 140 including the OBC 144, EPS 145 (including batteries), EACS 146 (which may be an ADCS including the EACS), TT&C 147, RFfront end 148 and the payload module (e.g., an SDR processor) 149. The components may be mounted on internal PCB 142 and other supporting components may be provided such as a harness 143 as well as heat sinks for the payload module and batteries. The heat sinks may be connected to the shielded enclosure 110 via a thermal gap pad so that the heat can be conducted to the shielded enclosure, or via any other method known to those skilled in the art. The heatsinks may also be formed as supporting structures to act as battery holders and to hold the payload in place during launch. Bulkhead connectors 141 are provided through the shielded enclosure 110, for example to connect the solar panels to the EPS, the antennas to the RF front end.

[0091] Figure 6 is a plot of the diversity mode antenna gain (dBi) vs. angle (0) during a tumble in a plane which includes the worst individual antenna gains (the worst tumble plane). The combined gain curve 124 was estimated for the satellite shown in Figure 5B in which the dipole antennas 121 and 122 are orthogonal (e.g. have complementary antenna patterns), and each individual antennal peak gain was 3.5dBi. From Figure 6 we see that the combined gain curve 124 varies from around -1.5dBi to 2.5dBi. Thus, even in the case of a tumbling satellite, it is expected that at least one of the antennas will receive, with a typical antenna gain of OdBi or more (and thus OdBi may be used for link budget estimation).

[0092] Figure 7A is a side view of the structural layout of a stack of four satellites shown in Figures 5A to 5F for loading into a 3U canister according to an embodiment. That is the total height H and width W are selected to fit within a 3U canister. The four satellites 10a 10b 10c lOd each comprise upper solar panels 131, upper panel 111, shielded enclosure 110, lower panel 112 and lower solar panels 132 with respective labels a, b, c and d to indicate each satellite. Figure 7B is a side sectional view of the structural layout of the shielded section shown in Figure 7A according to an embodiment showing bulkhead connectors 141 provided in the upper and lower surfaces of the shielded enclosure 110. The bulkhead connectors 141 allow the power from the solar panels to enter the shielded region without creating EMI issues for the antennas. In this embodiment the shielded enclosure 110 is constructed of 3mm Aluminium and internal PCBs 142 are respectively located adjacent the upper surface 111 and lower surfaces 112. A connecting harness 143 is provided between the two PCBs 142. In this embodiment the PCBs 142 are 1.6mm thick FR-4 type PCBs. The various components 140 (e.g. OBC, TT&C, ADCS, EACS, EPS) are mounted to the PCB boards 142. Heatsinks (not shown) are formed in the side walls and used to support the batteries and payload and are connected to the PCBs 142 via the harness 143. The solar panels are electrically connected (soldered) to the outer PCB panels 111 and 112 which are 2mm thick FR-4 type PCBs and bolted into place, for example using M3 bolts in each comer of the panel.

[0093] As shown in Figure 7A the deployment system comprises the four satellites formed into two pairs where 10a and 10b form the first pair, and 10c and lOd form the second pair. Each pair comprises a satellite in an inverted configuration so the paired satellites form a rectangular slab where the shieldedhousing 110 of one satellite fits under the extended portion of the upper surface of the other satellite in the pair (to fill the gap under that portion). A small gap 118 is created between each pair of adjacent satellites by washers and screws which attach the solar panels to the PCB boards and a release mechanism 119 is provided in this gap 118 between each pair of adjacent satellites. The release mechanism may be a bum wire, spring, or other component configured to control the release of the satellites. More generally, the satellites may be dimensioned to form complementary groups of satellites that can be stacked inside a deployment container, or attached to a deployer, and a release mechanism is provided between each satellite in the complementary group of satellites. The deployment container may be a standard 3U or 6U container, or some other form factor.

[0094] In some embodiments the satellites are deployed at the same time. For example, all release mechanisms 119 may be triggered at once, or passive release mechanisms 119 such as springs may be used which act as soon as the satellite is deployed from the container. In some embodiments, each satellite in the stack is deployed at a different point in time to spread each of the satellites across the orbit. In some embodiments the release mechanism is a spring and bum wire integrated in the PCB forming the panel 111 112 of each satellite. In an embodiment where a helical spring antenna is used, this could be used as a spring to separate the satellite from the other satellites by deploying the antenna. Deployment sensing may be performed through momentum measurements. The satellites could be tethered together and then each tether released.

[0095] The timing of the deployment may be automated based on a predetermined time or orbital position. In another embodiment the deployment may be controlled by a ground instruction. In some embodiments the satellites are configured to communicate with each other when tethered together to coordinate deployment and / or when in close proximity shortly after deployment. This may be via a wired connection, such as via the springs used to separate the satellites or a wireless connection may be used. In one embodiment one of the satellites in the group may deploy its antennas to facilitate communication with the ground, and then act as the conduit for control signals between ground and all satellites in the group. Satellites may then be deployed from the group via control instructions.

[0096] In one embodiment the relative difference in mass when deployment occurs may be exploited to distribute the satellites, as the Nth deployed satellite is l / (4-N) the mass of the remainder. This is illustrated in Figure 8A which is a plot of the displacement over a time for each of the four of the satellites shown in Figure 5D starting prior to deployment and for a short time frame after deployment. The displacement of the first, second, third and fourth satellites to deploy are shown respectively as dashed line 802, long dash line 804, dash dot line 806, and solid line 808. Figure 8B is a plot of the displacement over a time for each of the four of the satellites shown in Figure 5D starting prior to deployment and for a long time frame (longer than the short time frame shown in Figure 8A) after deployment according to an embodiment. The displacement of the first, second, third and fourth satellitesto deploy are shown respectively as dashed line 802', long dash line 804', dash dot line 806', and solid line 808'. These plots are based on one satellite sequentially maintaining the same direction. The separation is quite asymmetrical, particularly for the satellite which is last to deploy (solid thick line 808 and 808'). In some embodiments the timing of when to split off each satellite is determined using orbital modelling / simulations. The orbital modelling may be performed to optimise a particular orbital distribution of satellites, or to achieve specific orbits associated with the satellite mission. In one variation the satellites are split into pairs, for example by activating release mechanisms 119c and 119d, and then each pair is split again. For example, the first pair can be slit by subsequently activating release mechanisms 119a and 119b, and the second pair can be slit by subsequently activating release mechanisms 119e and 119f. After splitting the 4 satellites into two pairs, splitting of each of the pairs into single satellites can be performed at the same time, or each pair can be further split into single satellites at different times. Orbital simulations / modelling may be performed to determine the timing of when the to split the four satellites into two pairs, and when the satellites forming each pair is split into single satellites.

[0097] In one embodiment the payload module 149 may be an access node for the satellite communication system illustrated in Figure 2A. In this embodiment the payload module 149 comprises a baseband processor, comprising a general-purpose processor, graphics processing unit, field programmable gate array, or similar computational architecture combined with one or more transceivers to act as a Software Defined Radio (SDR). In this embodiment when receiving the RF transceivers pass IQ samples to the baseband processor for demodulation and decoding. For transmitting the baseband processor generates and sends IQ samples to one of the RF transceivers for transmission via the RF front end and antennas. In a preferred embodiment, a GNSS (e.g. GPS, Galileo, GLONASS, BEIDOU, etc), receiver is also integrated into the satellite with a GNNS antenna located on the rear side (see Figure 5C) which is connected to a GNSS receiver circuit in the shielded enclosure. The GNSS may be used to provide the location of the satellite with respect to Earth as well as a timestamp that is used to update the on-board clock, and a 1PPS signal. This information can be used by the access node payload service to determine when the satellite will receive and transmit. Furthermore, it can be used for end-of-life deorbiting of the satellite.

[0098] The communication system and components illustrated in Figure 2A may be configured as described in the following patent applications:PCT / AU2013 / 000895 titled CHANNEL ALLOCATION IN A COMMUNICATION SYSTEM and filed on 14 / 08 / 2013 claiming priority from Australian Provisional Patent Application No. 2012903489 filed on 14 / 08 / 2012;PCT / AU2013 / 001078 titled COMMUNICATION SYSTEM AND METHOD and fded on 20 / 09 / 2013 claiming priority from Australian Provisional Patent Application No. 2012904130 fded on 21 / 09 / 2012;PCT / AU2013 / 001079 titled MULTI-ACCESS COMMUNICATION SYSTEM and fded on 20 / 09 / 2013 claiming priority from Australian Provisional Patent Application No. 2012904145 fded on 21 / 09 / 2012;PCT / AU2014 / 000826 titled A MULTIUSER COMMUNICATIONS SYSTEM and fded on 21 / 08 / 2014 claiming priority from Australian Provisional Patent Application No. 2013903163 fded on 21 / 08 / 2013;PCT / AU2015 / 000743 titled MULTICARRIER COMMUNICATIONS SYSTEM and fded on 9 / 12 / 2015 claiming priority from Australian Provisional Patent Application No. 2014904976 fded on 9 / 12 / 2014;PCT / AU2017 / 000058 titled TERMINAL SCHEDULING METHOD IN SATELLITE COMMUNICATION SYSTEM and fded on 24 / 02 / 2017 claiming priority from Australian Provisional Patent Application No. 2016900685 fded on 25 / 02 / 2016;PCT / AU2017 / 000108 titled POSITION ESTIMATION IN A LOW EARTH ORBITSATELLITE COMMUNICATIONS SYSTEM and fded on 16 / 05 / 2017 claiming priority from Australian Provisional Patent Application No. 2016901913 fded on 20 / 05 / 2016;PCT / AU2017 / 000286 titled SYSTEM AND METHOD FOR GENERATING EXTENDEDSATELLITE EPHEMERIS DATA and fded on 21 / 12 / 2017 claiming priority from Australian Provisional Patent Application No. 2016905314 fded on 22 / 12 / 2016;PCT / AU2018 / 000151 titled SYSTEM AND METHOD FOR PREDICTION OFCOMMUNICATIONS LINK QUALITY and fded on 28 / 08 / 2018 claiming priority from Australian Provisional Patent Application No. 2017903470 fded on 28 / 08 / 2017;PCT / AU2021 / 000027 titled SYSTEM AND METHOD FOR ADAPTIVECOMMUNICATIONS and fded on 29 March 2021 claiming priority from Australian Provisional Patent Application No. 2020901049 fded on 3 / 04 / 2020.

[0099] Various embodiments are configured to reduce size, weight and battery consumption and extend battery or mission life such as by omitting conventional attitude determination and control systems which typically use heavy and power intensive attitude controllers such as always on reaction wheels, and using an emergency attitude control system which generates a short duration angular momentum perturbation (kick) upon detection of an undesirable orientation or an undesirable spin state of the satellite (note that a simple ADS or ADCS may also be used with the EACS). This may be a basic perturbation apparatus such as an uncontrolled DC motor to drive movement of a mass, but stripped back of unnecessary components like a speed controller. Other simple motorised mechanism or magnetorquers may be used to induce tumble or perturbations. In some embodiment the satellite does not rely on accurate pointing systems andmay use a coarse attitude determination system (ADS or ADCS). Solar panels may be located to maximise solar energy harvesting within the available form factor (e.g. on the side or sides with the largest surface area), and power generation or battery charge state is monitored and used to trigger use of an emergency attitude control subsystem upon a detection of an undesirable orientation or an undesirable spin state of the satellite. In some embodiments one or more solar panels may be mounted to each of at least two sides of the satellite so that power may be generated as the satellite tumbles. The system design may use a single RF band or several RF bands and permit the use of omnidirectional antennas such as VHF, UHF, L-band or S-band antennas. Multiple antennas may be used and configured to have complementary antenna patterns allow the use of use of transmit or receiver diversity and / or signal processing on the received signals to ensure communications are maintained whilst only using low accuracy pointing or no pointing / full tumble (that is there are no hard pointing requirements for maintaining communications).

[0100] Embodiments of the satellite may feature heavy integration and simple construction, for example using only two PCBs for bus and payload. The satellites may be constructed to have a small form factor with densely populated electronics to maximise use of volume and also features simple assembly / disassembly and easy access fortesting. The use of deployable components may be minimised to reduce deployment risk. The cost may be kept comparatively low via use of commercial components and small form factor thus allowing a large number of satellites to be deployed over time. Components may be selected based on previous flight heritage or designed to be tolerant to radiation effects. The shielded enclosure 110 may be formed of a metal, such as aluminium, and can be used to enclose and protect critical electrical hardware (except solar panels) from radiation. In some embodiments the design features separation and protection of critical components including the OBC, TT&C communications as well as the use of watchdogs to catch and correct undesirable system states. The bulkheads 141 provide EMI fdtering for electrical connections / hamesses passing through the shield. Similarly careful component selection, power system design, routing, and shielding maybe used to avoid the generation of electromagnetic interference. The use of 4 PCBs (including solar panels) for all the electronic components on each satellite allows the construction of a satellite with a small and compact form factor.

[0101] Embodiments of the satellites described herein can be used to lower the cost of launch. The compact design allows multiple satellites to be grouped and stacked efficiently into a small volume. This may also permit a launch canister to contain multiple satellites, such as four satellites being stacked into a standard 3U or 6U deployer (other combinations may be used). Deployment may be controlled to obtain a desired orbital separation. The design allows use in both earth orbits (LEO, MEO, GEO) as well as lunar or mars orbits and do not require hard pointing requirements (e.g. there is no requirement to point a sensor or antenna in a particular direction). The embodiments as described herein may be utilised to design small satellites, or to lower cost and complexity of larger satellites. That is, it is to be understoodthat embodiments as described herein are not limited to use only in small satellites, but may be used in any satellites, including as backup, or redundant systems for more complex and costly systems.

[0102] Those of skill in the art would understand that information and signals may be represented using any of a variety of technologies and techniques. For example, data, instructions, commands, information, signals, bits, symbols, and chips may be referenced throughout the above description may be represented by voltages, currents, electromagnetic waves, magnetic fields or particles, optical fields or particles, or any combination thereof.

[0103] Those of skill in the art would further appreciate that the various illustrative logical blocks, modules, circuits, and algorithm steps described in connection with the embodiments disclosed herein may be implemented as electronic hardware, computer software or instructions, or combinations of both. To clearly illustrate this interchangeability of hardware and software, various illustrative components, blocks, modules, circuits, and steps have been described above generally in terms of their functionality. Whether such functionality is implemented as hardware or software depends upon the particular application and design constraints imposed on the overall system. Skilled artisans may implement the described functionality in varying ways for each particular application, but such implementation decisions should not be interpreted as causing a departure from the scope of the present invention.

[0104] The steps of a method or algorithm described in connection with the embodiments disclosed herein may be embodied directly in hardware, in a software module executed by a processor, or in a combination of the two. For a hardware implementation, processing may be implemented within one or more application specific integrated circuits (ASICs), digital signal processors (DSPs), digital signal processing devices (DSPDs), programmable logic devices (PLDs), field programmable gate arrays (FPGAs), processors, controllers, micro-controllers, microprocessors, or other electronic units designed to perform the functions described herein, or a combination thereof.

[0105] In some embodiments the On-Board Computer (OBC) module comprises one or more processors which may be one or more Central Processing Units (CPUs). The OBC module is configured to control the satellite and to perform some of the steps of the methods. The payload module may also comprise a computing module including one or more CPUs. A CPU may comprise an Input / Output Interface, an Arithmetic and Uogic Unit (AUU) and a Control Unit and Program Counter element which is in communication with input and output devices through the Input / Output Interface. In one embodiment the OBS is a radiation hardened computing module such as a MSP430 microcontroller manufactured by Texas Instruments or similar module. The computing modules may comprise a single CPU (core) or multiple CPU’s (multiple core), or multiple processors and memory may be is operatively coupled to the processor(s) and may comprise RAM and ROM components, and may be provided within or external to the device or processor module. The memory may be used to store an operating system and additionalsoftware modules or instructions. The processor(s) may be configured to load and executed the software modules or instructions stored in the memory.

[0106] Software modules, also known as computer programs, computer codes, or instructions, may contain a number a number of source code or object code segments or instructions, and may reside in any computer readable medium such as a RAM memory, flash memory, ROM memory, EPROM memory, registers, hard disk, or any other form of computer readable medium. In some aspects the computer- readable media may comprise non-transitory computer-readable media (e.g., tangible media). In addition, for other aspects computer-readable media may comprise transitory computer- readable media (e.g., a signal). Combinations of the above should also be included within the scope of computer-readable media. In another aspect, the computer readable medium may be integral to the processor. The processor and the computer readable medium may reside in an ASIC or related device. The software codes may be stored in a memory unit and the processor may be configured to execute them. The memory unit may be implemented within the processor or external to the processor, in which case it can be communicatively coupled to the processor via various means as is known in the art.

[0107] Further, it should be appreciated that modules and / or other appropriate means for performing the methods and techniques described herein can be downloaded and / or otherwise obtained by computing device. For example, such a device can be coupled to a server to facilitate the transfer of means for performing the methods described herein. Alternatively, various methods described herein can be provided via storage means (e.g., RAM, ROM, a physical storage medium, etc.), such that a computing device can obtain the various methods upon coupling or providing the storage means to the device. Moreover, any other suitable technique for providing the methods and techniques described herein to a device can be utilized.

[0108] The methods disclosed herein comprise one or more steps or actions for achieving the described method. The method steps and / or actions may be interchanged with one another without departing from the scope of the claims. In other words, unless a specific order of steps or actions is specified, the order and / or use of specific steps and / or actions may be modified without departing from the scope of the claims.

[0109] As used herein, the terms “estimating” or “determining” encompasses a wide variety of actions. For example, “estimating” or “determining” may include calculating, computing, processing, deriving, investigating, looking up (e.g., looking up in a table, a database or another data structure), ascertaining and the like. Also, “estimating” or “determining” may include receiving (e.g., receiving information), accessing (e.g., accessing data in a memory) and the like. Also, “determining” may include resolving, selecting, choosing, establishing and the like.

[0110] The reference to any prior art in this specification is not, and should not be taken as, an acknowledgement or any form of suggestion that such prior art forms part of the common general knowledge.

[0111] It will be understood that the terms “comprise” and “include” and any of their derivatives (e.g. comprises, comprising, includes, including) as used in this specification, and the claims that follow, is to be taken to be inclusive of features to which the term refers, and is not meant to exclude the presence of any additional features unless otherwise stated or implied.

[0112] In some cases, a single embodiment may, for succinctness and / or to assist in understanding the scope of the disclosure, combine multiple features. It is to be understood that in such a case, these multiple features may be provided separately (in separate embodiments), or in any other suitable combination. Alternatively, where separate features are described in separate embodiments, these separate features may be combined into a single embodiment unless otherwise stated or implied. This also applies to the claims which can be recombined in any combination. That is a claim may be amended to include a feature defined in any other claim. Further a phrase referring to “at least one of’ a list of items refers to any combination of those items, including single members. As an example, “at least one of: a, b, or c” is intended to cover: a, b, c, a-b, a-c, b-c, and a-b-c.

[0113] It will be appreciated by those skilled in the art that the disclosure is not restricted in its use to the particular application or applications described. Neither is the present disclosure restricted in its preferred embodiment with regard to the particular elements and / or features described or depicted herein. It will be appreciated that the disclosure is not limited to the embodiment or embodiments disclosed, but is capable of numerous rearrangements, modifications and substitutions without departing from the scope as set forth and defined by the following claims.

Claims

CLAIMS1. A satellite comprising: an on-board computer (OBC) comprising at least one processor and at least one memory, a telemetry tracking and command subsystem (TT&C), an electrical and power subsystem, (EPS) comprising one or more batteries; a payload module, an emergency attitude control subsystem (EACS) configured to generate an angular momentum perturbation upon detection of an undesirable orientation or an undesirable spin state of the satellite; one or more solar panels mounted to one or more sides of the satellite and electrically connected to the one or more batteries; and an RF front end comprising at least two antennas wherein each antenna is connected to a different side of the satellite and the OBC is configured to select a set of one or more antennas to transmit or receive on.

2. The satellite as claimed in claim 1, further comprising an attitude determination system (ADS) comprising one or more attitude sensors to determine a pointing direction of the satellite, and the ADS is configured to detect the undesirable orientation or the undesirable spin state of the satellite and to instruct the EACS to generate an angular momentum perturbation upon the detection.

3. The satellite as claimed in claimed 2 wherein the one or more attitude sensors comprises an inertial measurement unit (IMU).

4. The satellite as claimed in claim 2 or 3, wherein the one or more attitude sensors comprises one or more sun sensors on the one or more solar panels.

5. The satellite as claimed in any one of claims 1 to 4, wherein the OBC is configured to utilise the at least two antennas based on either: selecting all of the at least two antennas and determining a sequence in which each antenna in the sequence is switched on for an active time period and then off, wherein the sequence is either an ordered sequence which is continuously repeated, or is a random sequence, or is a pseudo random sequence; or grouping the at least two antennas into a set of complementary antenna groups of two or more, and determining a group sequence in which each antenna in the group is switched on for an active time period, and then off, and continuously cycling through the group sequence; or selecting at least two antennas, and simultaneously receiving or transmitting on the at least two antennas; ormeasuring an orientation performance indicator for each of the at least two antennas, and selecting at least one antenna based on the orientation performance indicators such that over an extended reference time period the selected at least one antenna is used for a longer amount of time than a nonselected antenna or an antenna with a lower orientation performance indicator.

6. The satellite as claimed in claim 5, wherein the orientation performance indicator is a performance indicator and the OBC determines a performance weight for each antenna and utilising the one or more antennas is performed such that over the extended time period an antenna with more favourable weight is switched on for a longer total time than an antenna with less favourable weight.

7. The satellite as claimed in claim 5 or 6 when dependent through claim 2, wherein the orientation performance indicator for each antenna is a pointing direction indicator which is a measure of the pointing direction of the antenna with respect to a reference direction determined using the one or more attitude sensors.

8. The satellite as claimed in in any one of claims 5 to 7 when dependent through claim 2, wherein selecting at least two antennas and simultaneously receiving or transmitting on the at least two antennas is performed using a coarse pointing direction of each of the at least two antennas determined using the one or more attitude sensors, and the OBC is configured to use beam forming or signal processing of received or transmitted signals using the coarse pointing directions to compensate for effects of a varying pointing direction or tumbling of the satellite.

9. The satellite as claimed in any one of claims 1 to 8, wherein the EACS comprises one or more DC motors that are configured to drive movement of one or more masses, and one or more of the one or more DC motors are switched on for an activation time period upon detection of an undesirable orientation or an undesirable spin state of the satellite.

10. The satellite as claimed in claim 9, wherein one or more of the one or more DC motors drive a shaft, and one or more of the one or more masses are mounted on the shaft, and an axis of the shaft is inclined with respect to at least one side of the satellite on which one or more solar panels are mounted.

11. The satellite as claimed in any one of claims 1 to 10 wherein the EACS comprises one or more magnetorquers that are activated for an activation time period upon detection of the undesirable orientation or the undesirable spin state of the satellite.

12. The satellite as claimed in claim 11, when dependent through claim 2, wherein the ADS is an attitude determination and control system (ADCS) which is configured to use the one or more themagnetorquers to perform coarse control of attitude to control capture of solar light by the one or more solar panels.

13. The satellite as claimed in any one of claims 1 to 12, wherein the EPS further comprises one or more sensors for determining a battery charge level of the one or more batteries, and the EACS is further configured to monitor a battery charge level of the one or more batteries, and detecting the undesirable orientation or the undesirable spin state of the satellite is detected by comparing the battery charge level of at least one of the one or more batteries with a threshold value, or by analysing a time history of the battery charge level of the one or more batteries and identifying a trend of decreasing battery charge level overtime.

14. The satellite as claimed in any one of claims 1 to 13 wherein the satellite has a rectangular cuboid shape, and each of the at least two antennas is a patch antenna, and each of the at least two antennas are located on a different side of the satellite.

15. The satellite as claimed in claim 14 wherein at least one patch antenna is collocated with a solar panel.

16. The satellite as claimed in any one of claims 1 to 15, further comprising: a shielded enclosure for housing at least the OBC, TT&C, and EACS, wherein the shielded enclosure is manufactured of a material configured to reduce radiation to a plurality of components housed with the enclosure, and comprises one or more bulkheads to allow passage of electrical connections through the shielded enclosure.

17. The satellite as claimed in claim 16 wherein the shielded enclosure is a substantially rectangular cuboid housing with a front side, a rear side, a left side, a right side, an upper side and a lower side, and the front side, rear side, left side and right side form four sides of the satellite, and an upper panel is mounted to the upper side of the shielded enclosure and comprises a printed circuit board (PCB) supporting one or more solar panels on an exterior side of the upper panel, and an interior side of the upper panel is divided into two portions, wherein a first portion is mounted to the upper side of the shielded enclosure and is electrically connected to a first bulkhead in the upper side of the shielded enclosure, and a second portion extends in a direction normal to the front side and supports one or more solar panels, and a PCB is connected to the lower side of the shielded enclosure and is electrically connected to a second bulkhead in the lower side of the shielded enclosure, and supports one or more solar panels, and the at least two antennas comprises a first dipole antenna mounted to the left side and a second dipole antenna mounted to the right side such that when deployed the first and second antennas have an orthogonal configuration.

18. The satellite as claimed in claim 17, wherein the left side and the right side each have a step feature, and prior to deployment the first dipole antenna and the second dipole antenna are held in a stowed configuration such that they are housed under the upper panel, and upon deployment, each antenna is released and extends laterally beyond an edge of the upper panel until they reach a respective stop formed by the respective step feature in each of the left side and right side.

19. The satellite as claimed in any one of claims 1 to 18, wherein the payload module is integrated into the OBC wherein the payload module comprises a set of instructions stored in the at least one memory of the OBC which are executed by the at least one processor of the OBC to provide payload functionality.

20. A deployment system comprising a plurality of satellites as claimed in claim 18 or 19 wherein the plurality of satellites is dimensioned to form complementary groups of satellites that can be stacked inside a deployment container, and a release mechanism is provided between each satellite in the complementary group of satellites.

21. The deployment system as claimed in claim 20, wherein the deployment container is a 3U or a 6U container.

22. The deployment system of claim 20 or 21 wherein each satellite in the stack is deployed at different points in time to spread their distribution across an orbit.

23. The deployment system of claim 20, 21, or 22, wherein the release mechanism is a spring and bum wire integrated in a PCB forming a solar panel of each satellite.

24. A method for operating the satellite of any one of claims 1 to 19 comprising: monitoring for an undesirable orientation or an undesirable spin state of the satellite and upon detection of the undesirable orientation or the undesirable spin state of the satellite, generating, by the EACS, an angular momentum perturbation; and utilising a set of one or more antennas to transmit or receive on.

25. The method as claimed in claim 24, wherein utilising a set of one or more antennas to transmit or receive on comprises: selecting all of the at least two antennas and determining a sequence in which each antenna in the sequence is switched on for an active time period and then off, and continuously cycling through the sequence; orgrouping the at least two antennas into a set of complementary antenna groups of two or more and determining a group sequence in which each antenna in a group is switched on for an active time period, and then off, and continuously cycling through the group sequence; or selecting at least two antennas and simultaneously receiving or transmitting on the at least two antennas; or measuring an orientation performance indicator for each of the at least two antennas and selecting at least one antenna based on the orientation performance indicators such that over an extended time period the selected at least one antenna is used for a longer amount of time than a non-selected antenna or an antenna with a lower orientation performance indicator.

26. The method as claimed in claim 25, wherein the orientation performance indicator is a performance indicator, and the method further comprises: determining a performance weight for each antenna and utilising the one or more antennas is performed such that over an extended time period an antenna with a more favourable performance weight is switched on for a longer total time than an antenna with a less favourable performance weight.

27. The method as claimed in claim 25 or 26, wherein the orientation performance indicator for each antenna is a pointing direction indicator which is a measure of the pointing direction of the antenna with respect to a reference direction.

28. The satellite as claimed in claim 25 or 26 wherein selecting at least two antennas and simultaneously receiving or transmitting on the at least two antennas is performed using a coarse pointing direction of each of the at least two antennas, and the method further comprises performing beam forming or signal processing of a received signal or signal to be transmitted using the coarse pointing directions to compensate for effects of a varying pointing direction or tumbling of the satellite.

29. The method as claimed in claim 26, further comprising determining a coarse pointing direction of the satellite, and the orientation indicator for each antenna is a time average measure of a pointing direction to a sun based on the coarse pointing direction.

30. A method as claimed in any one of claims 24 to 29 wherein detecting the undesirable orientation or the undesirable spin state of the satellite comprises monitoring a battery charge level of the one or more batteries and determining a trend of decreasing battery charge level overtime.

31. A method for deploying the deployment system of claim 21 comprising: triggering each release mechanism at different points in time.

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