WO2017195202A1 - Gestion d'énergie de satellite - Google Patents

Gestion d'énergie de satellite Download PDF

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Publication number
WO2017195202A1
WO2017195202A1 PCT/IL2017/050518 IL2017050518W WO2017195202A1 WO 2017195202 A1 WO2017195202 A1 WO 2017195202A1 IL 2017050518 W IL2017050518 W IL 2017050518W WO 2017195202 A1 WO2017195202 A1 WO 2017195202A1
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Prior art keywords
satellite
energy
functional elements
attitude
future
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English (en)
Inventor
Nisim YECHEZKEL
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Israel Aerospace Industries Ltd
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Israel Aerospace Industries Ltd
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Classifications

    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles
    • B64G1/10—Artificial satellites; Systems of such satellites; Interplanetary vehicles
    • B64G1/1021—Earth observation satellites
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles
    • B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
    • B64G1/40—Arrangements or adaptations of propulsion systems
    • B64G1/411—Electric propulsion
    • B64G1/413—Ion or plasma engines
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles
    • B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
    • B64G1/42—Arrangements or adaptations of power supply systems
    • B64G1/428—Power distribution and management
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles
    • B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
    • B64G1/42—Arrangements or adaptations of power supply systems
    • B64G1/44—Arrangements or adaptations of power supply systems using radiation, e.g. deployable solar arrays
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles
    • B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
    • B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
    • B64G1/244—Spacecraft control systems
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles
    • B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
    • B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
    • B64G1/36—Guiding or controlling apparatus, e.g. for attitude control using sensors, e.g. sun-sensors, horizon sensors
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
    • B64G1/00—Cosmonautic vehicles
    • B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
    • B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
    • B64G1/36—Guiding or controlling apparatus, e.g. for attitude control using sensors, e.g. sun-sensors, horizon sensors
    • B64G1/361—Guiding or controlling apparatus, e.g. for attitude control using sensors, e.g. sun-sensors, horizon sensors using star sensors

Definitions

  • the presently disclosed subject matter relates to satellite energy consumption regime
  • Satellites are equipped with limited energy resources. Energy consumption regime onboard the satellite is regulated in order to extend the satellite life time.
  • Earth satellite imagery includes images which are collected by a camera located onboard a satellite orbiting the Earth. Satellites configured for obtaining imagery of the Earth (herein “imagery acquiring satellites") belong to Low-Earth orbits (LEO) category of satellites which occupy a region of space from about 180 kilometers to 2000 kilometers above the Earth surface. While it would have been advantageous for imagery acquiring satellites to orbit the Earth at low altitudes in order to improve the resolution of the acquired images, altitudes below 500-600 kilometers are generally avoided because of the substantial atmospheric-drag at such altitudes.
  • LEO Low-Earth orbits
  • the atmospheric-drag generates orbital decay, where the satellite altitude gradually decreases, and thus necessitates frequent orbit re-boost maneuvers in order to direct the satellite back to a higher altitude orbit.
  • Re-boost maneuvers are made with the help of thrusters, which consume the limited fuel stored onboard the satellite, and thus contribute to shorting the satellites life time.
  • a computer-implemented method of reducing satellite drag comprising an imaging subsystem configured to capture images of the Earth, wherein the satellite is orientated in a substantially vertical attitude while capturing images, which increases drag compared to the satellite being oriented in a substantially horizontal attitude
  • the method comprising using a processor for: determining information indicative of current position and attitude of the satellite; obtaining information indicative of areas on Earth sought to be captured by the imaging subsystem; determining whether current position of satellite coincides with an area on Earth sought to be captured by the imaging system; if the satellite position does coincide, generating instructions to a control subsystem to adapt the satellite attitude to a substantially vertical attitude where the camera is directed towards the Earth; and generating instructions for controlling the imaging subsystem to activate the camera and capture images of Earth; and if the satellite position does not coincides with an area on Earth sought to be captured and the satellite is in a substantially vertical attitude, generating instructions to a control subsystem to adapt the satellite attitude to
  • the method according to the above aspect can optionally comprise one or more of features (i) to (iv) below, in any technically possible combination or permutation: I. wherein the satellite comprises ionic propulsion, the method comprising: responsive to a need to activate the ionic propulsion, determining whether the imaging subsystem is operating; and generating instruction to activate the ionic propulsion only if the imaging subsystem is not operating and the attitude of the satellite is in a substantially horizontal position. II.
  • the satellite comprises solar panels configured to absorb solar energy from the sun
  • the method further comprising: generating instructions, during a predefined period of time from entry to satellite day time, for orienting the satellite in a substantially horizontal position to minimize drag on the expense of exposure of the solar panels to sun-light.
  • the method further comprising: selectively operating one or more functional elements onboard a satellite, comprising operating at least one processor for: determining available solar energy during the operation of the one or more functional elements; determining total available energy including available stored energy and the available solar energy; determining required energy; wherein determination of the required energy takes into consideration estimated energy consumed by the one or more functional elements and the drag exerted on the satellite in a desired satellite attitude and satellite altitude; determining a difference between the total available energy and the required energy; and in case the difference complies with one or more predefined conditions, generating instructions for activating the one or more functional elements.
  • determining available solar energy during the operation of the one or more functional elements is based on at least beta angle and a satellite attitude.
  • a satellite control system configured to reduce satellite drag
  • the satellite control system is operatively connectable to a plurality of satellite subsystems including a first subsystem configured to control the attitude and position of the satellite and an imaging subsystem configured to capture images; wherein the satellite is orientated in a substantially vertical attitude while capturing images of Earth by the imaging subsystem that increases drag compared to the satellite being oriented in a substantially horizontal attitude;
  • the satellite control system comprises a control unit configured to: determine information indicative of current position and attitude of the satellite; obtain information indicative of areas on Earth sought to be captured by the imaging subsystem; determine whether current position of satellite coincides with an area on Earth sought to be captured by the imaging system; if the satellite position does coincide, generate instructions to operate the first control system for adapting the satellite attitude to a substantially vertical attitude where the camera is directed towards the Earth; and generate instructions to the imaging subsystem to activate the camera and capture images of Earth; and if the satellite position does not coincides with an area on Earth sought to be captured and the satellite is
  • the satellite control unit can optionally comprise one or more of features (i) to (iv) above, in any technically possible combination or permutation.
  • a machine-readable non-transitory memory device tangibly embodying a program of instructions executable by the machine for executing a method of reducing satellite drag, the satellite comprising an imaging subsystem configured to capture images of the Earth, wherein the satellite is orientated in a substantially vertical attitude while capturing images, which increases drag compared to the satellite being oriented in a substantially horizontal attitude, the method comprising: determining information indicative of current position and attitude of the satellite; obtaining information indicative of areas on Earth sought to be captured by the imaging subsystem; determining whether current position of satellite coincides with an area on Earth sought to be captured by the imaging system; if the satellite position does coincide, generating instructions to a control subsystem to adapt the satellite attitude to a substantially vertical attitude where the camera is directed towards the Earth; and generating instructions for controlling the imaging
  • the memory device according to the above aspect can optionally comprise one or more of features (i) to (iv) above, in any technically possible combination or permutation.
  • a computerized method of selectively controlling the operation of one or more functional elements onboard a satellite the satellite having limited energy resources and comprises a plurality of functional elements; the method comprising using at least one processor for: determining available solar energy during the operation of the one or more functional elements; determining total available energy including available stored energy and the available solar energy; determining required energy; wherein determination of the required energy takes into consideration estimated energy consumed by the one or more functional elements and the drag exerted on the satellite in a desired satellite attitude and satellite altitude; determining a difference between the total available energy and the required energy; and in case the difference complies with one or more predefined conditions, generating instructions for activating the one or more functional elements.
  • the method according to the above aspect can optionally comprise one or more of features (i) to (vii) below, in any technically possible combination or permutation.
  • the one or more functional elements is an imaging subsystem configured to capture images of the Earth.
  • the one or more functional elements include a plurality of functional elements, wherein in the event that the total available energy is sufficient for the operation of only part of the plurality of functional elements, the method further comprising: selecting one or more functional elements from the plurality of functional elements to be activated based on the respective energy consumption and priority rank of each one of the plurality of functional elements.
  • determining available solar energy during the operation of the one or more functional elements is based on at least beta angle and a satellite attitude.
  • the method further comprising using the processor for determining operational parameters of one or more functional elements at a future time: determining estimated future available solar energy at the future time based on at least a future beta angle, a future satellite attitude and future satellite altitude which are expected at the future time; determining estimated future total available energy including estimated future available stored energy and the estimated future available solar energy; determining estimated future required energy at the future time; wherein determination of the required energy takes into consideration estimated energy consumed by the one or more functional elements and the drag exerted on the satellite at the future time; wherein the drag is estimated based on a desired satellite attitude and a desired altitude of the satellite at the future time; in case the difference between the total estimated future available energy and the estimated future required energy complies with one or more predefined conditions, generating instruction for adapting the altitude at the future time according to the future altitude and generating instructions to activate the one or more functional elements at the future time.
  • the method further comprising: in case the difference between the total estimated future available energy and the estimated future required energy is insufficient, adapting the one or more operational parameters in order to reduce the future required energy and increase the difference.
  • the one or more operational parameters include a desired altitude of the satellite at the future time
  • the method comprising: increasing the value of the future altitude in order to reduce the required energy; calculating an updated estimated future required energy; calculating the difference between the estimated future available energy; and if the difference between the total estimated future available energy and the updated estimated future required energy complies with one or more predefined conditions, generating instruction to activate the one or more functional elements at the future time.
  • the method further comprising repeatedly calculating the difference, each time with a different altitude value until the difference complies with the one or more conditions.
  • a satellite control system configured for selectively operating one or more functional elements onboard a satellite to: determine available solar energy during the operation of the one or more functional elements; determine total available energy including available stored energy and the available solar energy; determining required energy; wherein determination of the required energy takes into consideration estimated energy consumed by the one or more functional elements and the drag exerted on the satellite in the desired satellite attitude and satellite altitude; determine a difference between the total available energy and the required energy; and in case the difference complies with one or more predefined conditions, generate instructions for activating the one or more functional elements.
  • the satellite control unit according to the above aspect can optionally comprise one or more of features (i) to (vii) above, in any technically possible combination or permutation.
  • a machine-readable non-transitory memory device tangibly embodying a program of instructions executable by the machine for executing a method of selectively operating one or more functional elements onboard a satellite, the satellite having limited energy resources and comprising a plurality of functional elements; the method comprising: determining available solar energy during the operation of the one or more functional elements; determining total available energy including available stored energy and the available solar energy; determining required energy; wherein determination of the required energy takes into consideration estimated energy consumed by the one or more functional elements and the drag exerted on the satellite in the desired satellite attitude and satellite altitude; determining a difference between the total available energy and the required energy; and in case the difference complies with one or more predefined conditions, generating instructions for activating the one or more functional elements.
  • the memory according to the above aspect can optionally comprise one or more of features (i) to (vii) above, in any technically possible combination or permutation.
  • a computerized method of controlling a satellite comprising solar panels configured to absorb solar energy from the sun there is provided a computerized method, a satellite control unit, comprising: generating instructions, during a predefined period of time from entry to satellite day time, for orienting the satellite in a substantially horizontal position to minimize drag on the expense of sun-light exposure to the solar panels, while striving to increase exposure of the solar panel to the Sun as much as possible under these constraints.
  • a satellite control system a non-transitory storage device and a satellite are contemplated to be capable of generating instructions, during a predefined period of time from entry to satellite day time, for orienting the satellite in a substantially horizontal position to minimize drag on the expense of exposure of Sun light to the solar panels.
  • Fig.l is a schematic illustration of a satellite, in accordance with examples of the presently disclosed subject matter
  • Fig. 2a is a schematic illustration of satellite oriented in a substantially vertical attitude with respect to the Earth, in accordance with examples of the presently disclosed subject matter
  • Fig. 2b is a schematic illustration of satellite orientated in a substantially horizontal attitude with respect to the Earth, in accordance with examples of the presently disclosed subject matter;
  • Fig. 3 is a functional block diagram of a satellite control system, in accordance with examples of the presently disclosed subject matter;
  • Fig. 4 is a flowchart illustrating a sequence of operation performed in accordance with examples of the presently disclosed subject matter;
  • Fig. 5a is a schematic illustration demonstrating ⁇ angle, in accordance with examples of the presently disclosed subject matter
  • Fig. 5b is a schematic illustration demonstrating ⁇ angle, in accordance with examples of the presently disclosed subject matter.
  • Fig. 6 is a flowchart illustrating a sequence of operations performed in accordance with examples of the presently disclosed subject matter.
  • Satellite control system 300 comprises at least one computerized device (e.g. control unit 320) comprising a processing circuitry with data processing capabilities.
  • the processing circuitry comprises more processing devices such as a processor (e.g. digital signal processor (DSP), microcontroller, field programmable circuit (ASIC), etc.) and can further comprise or be otherwise operatively connected to a computer memory.
  • the computerized device can be for example a dedicated computer device or computer system or a general purpose computer device or computer system configured for operating as described herein below.
  • Fig. 3 illustrates a schematic of control system architecture in accordance with embodiments of the invention.
  • Functional elements in Fig. 3 can be made up of any combination of software and hardware, and/or firmware that performs the functions as defined and explained herein.
  • Functional elements in Fig. 3 illustrates may be centralized in one location or dispersed over more than one location. According to some examples of the presently disclosed subject matter, the system may comprise fewer, more and or different modules than those shown in Fig. 3 illustrates.
  • Satellite 100 comprises a number of subsystems including the structural subsystem, the telemetry subsystem, the power subsystem, thermal control subsystem and attitude and orbit control subsystem. Notably, not all subsystems are shown in Fig. 1. Satellite 100 further comprises an imaging subsystem (including at least one camera) and a communication payload configured to enable data transmission to and from the satellite.
  • the structural subsystem provides the mechanical infrastructure for holding together all the satellite functional subsystems and providing the required durability to withstand the mechanical stress and extreme temperatures during launch and while in orbit.
  • the telemetry subsystem is configured to transmit equipment operation data (e.g. to ground control system).
  • the power subsystem comprises solar panels 103 configured to absorb and convert solar energy into electrical power and charge the batteries that are configured to store electric power and supply the power to the satellite subsystems.
  • the thermal control subsystem is configured to regulate the internal temperature of the satellite subsystems and protect the onboard equipment from the extreme temperatures to which the satellite is exposed.
  • the attitude and orbit control subsystem comprises: sensors configured to determine the position and attitude of the satellite, and a flight control system configured to obtain and maintain the desired attitude of the satellite, keep the satellite in the desired orbital position and maintain the antennas directed to the desired position.
  • the attitude and orbit control subsystem further comprises various actuators, such as reaction wheels and thrusters which are controlled by the flight control system and enable to adjust and control the attitude and orbital position of the satellite.
  • the satellite's structural subsystem can have for example a cylindrical-like shape where the camera's aperture 105 is located at one end of the cylinder.
  • the camera is directed toward the Earth to enable capturing Earth satellite imagery thus positioning the satellite in a substantially vertical attitude (where the longitudinal axis of the satellite is in a substantially vertical position in respect of the Earth).
  • Fig. 2a is a schematic illustration showing a satellite in a substantially vertical attitude in respect of the Earth and Fig. 2b is a schematic illustration showing a satellite in a substantially horizontal attitude in respect of the Earth, according to examples of the presently disclosed subject matter.
  • Arrow 21 in Fig. 2a indicates that the camera is pointing downwards towards the Earth and arrow 23 in Fig. 2b indicates that the camera is pointing substantially horizontal to the Earth.
  • Due to the shape of the satellite when in vertical attitude, the drag exerted on the satellite is considerably greater than the drag exerted on the satellite in horizontal attitudes. As a result, when in vertical attitude considerable more propellant is spent to maintain the satellite in the desired orbit and specifically in a desired altitude.
  • substantially vertical attitude is used herein to include a range of angle values defined between camera/telescope bore-sight and nadir (e.g. if this angle is smaller than 45° the satellite is considered in substantially vertical attitude).
  • substantially horizontal attitude is used herein to include a range of angle values defined between camera/telescope bore-sight and the satellite instantaneous velocity vector (e.g. if this angle is smaller than 20°, or larger than 160°, the satellite is considered in substantially horizontal attitude).
  • the presently disclosed subject matter includes a computerized method and computerized system (being mountable on a satellite) which enable imagery acquiring satellites to reduce the altitude of the satellite orbit for operating the imaging subsystem at a closer distance from the Earth, while reducing the overall increase in drag exerted on the satellite and thus reducing the satellite energy consumption.
  • the satellite while the satellite is orbiting the Earth it is not required to continuously capture images of every part of the Earth. For example, it may be desired to capture images of the ground and not of the seas, or it may be desired to capture images of only specific areas of the Earth such as specific geographical areas or populated areas only.
  • the imaging subsystem is required to operate only during the time the satellite is located above those areas which are sought to be captured by the imaging subsystem.
  • a control system 300 which is configured to control the attitude of the satellite.
  • the control system switches between one attitude where the satellite is positioned in a substantially horizontal attitude and a second attitude where the satellite is positioned in a substantially vertical attitude. Switching from one attitude to the other depends on the current position of the satellite along its orbit as illustrated below with reference to Figs. 3 and 4.
  • Fig. 3 is a functional block diagram of satellite control system, in accordance with an example of the presently disclosed subject matter.
  • satellite control system 300 comprises sensor units, operatively connected to attitude and orbit subsystem 301, which is configured to provide information indicative of the attitude and position of the satellite.
  • Sensor units can include for example a GPS unit and star tracker unit.
  • control unit 320 is further operatively connected to one or more actuator control unit 303, imaging subsystem data unit 305, and imaging subsystem control unit 307.
  • Fig. 3 also shows ground control station 350 located remotely from satellite and configured in general to monitor and control the operation of the satellite.
  • Ground control station 350 is configured to communicate with the satellite e.g. for receiving telemetry data from the satellite and transmitting instructions and other data to the satellite.
  • Ground control station can be located at some location on earth (not limited to being located on land).
  • Imaging subsystem data unit 305 can include (or be otherwise connected to) a data- repository (including non-transitory computer memory) configured for storing information indicating when the imaging subsystem should operate and capture images of the Earth.
  • a data- repository including non-transitory computer memory
  • information indicating when the imaging subsystem should operate and capture images of the Earth can be provided (possibly in real-time) to the satellite by an external source, for example, from ground control station 350.
  • One or more actuator control units 303 are configured to control the operation of a respective actuator (including for example reaction wheels and thrusters) which enables in turn to control the attitude of the satellite.
  • Actuator control units 303 can be configured as part of the attitude and orbit control subsystem.
  • Imaging subsystem control unit 307 is configured to control the imaging subsystem including generating instruction to operate a camera or any other imaging device and capture images.
  • Fig. 4 is a flowchart illustrating sequence of operation performed in accordance with examples of the presently disclosed subject matter. Operations described below with reference to Fig. 4 can be executed for example by satellite control system 300 described above with reference to Fig. 3, however any description of operations in to Fig. 4 with reference to functional elements in Fig. 3 is done by way of example only and should not be construed as limiting.
  • the satellite As the satellite orbits the Earth it overpasses different areas of the Earth. Throughout the satellite motion it determines (e.g. continuously or repeatedly) its position and whether it is positioned over an area of the Earth which is desired to be captured by the imaging subsystem to obtain images thereof, or whether the satellite is located over an area of the Earth which is not desired to be captured by the imaging subsystem.
  • control unit 320 can be configured to obtain from various sensors (e.g. operatively connected to attitude and orbit subsystem 301) information indicative of the current position of the satellite and the area on the Earth which is currently being traversed (block 401) and obtain information indicative of the location of the areas which are desired to be captured (block 403). As mentioned above, this information can be obtained from imaging subsystem data unit 305 or from some other external source. Control unit 320 can be further configured to determine based on the obtained information whether the satellite is currently positioned over an area of the Earth which is desired to be captured (block 405).
  • various sensors e.g. operatively connected to attitude and orbit subsystem 301
  • information indicative of the location of the areas which are desired to be captured block 403
  • this information can be obtained from imaging subsystem data unit 305 or from some other external source.
  • Control unit 320 can be further configured to determine based on the obtained information whether the satellite is currently positioned over an area of the Earth which is desired to be captured (block 405).
  • the attitude of the satellite can be changed to a substantially vertical attitude where the aperture of the camera in the imaging subsystem is facing the Earth (in some cases substantially aligned with nadir) to enable the camera to capture images of the Earth (block 407).
  • control unit 320 which can be configured, responsive to determining that the satellite is currently positioned over an area of the Earth which is desired to be captured, to generate instructions to actuators control units to alter the attitude of the satellite to a substantially vertical attitude. The exact attitude would be determined according to the specific location of the area of interest to be captured by the camera. If for some reason the satellite is already in the vertical attitude the satellite can continue and operate in substantially vertical attitude while operating the imaging subsystem.
  • Control unit 320 can be further configured to generate instructions to activate the imaging subsystem which operates one or more cameras for capturing images of the Earth (block 409). If on the other hand it is determined that the satellite is currently positioned over an area of the Earth which is not desired to be captured, the attitude of the satellite can be changed to a substantially horizontal attitude where the optical axis of the imaging subsystem is substantially aligned with the satellite velocity and the imaging system is unable to capture images of the Earth (block 411). As before this can be accomplished with the help of control unit 320 which can be configured to provide instructions to actuators control units to alter the attitude of the satellite to a substantially horizontal altitude. If the satellite is already in the horizontal altitude no change to the attitude of the satellite is required.
  • Control unit 320 can be further configured to deactivate the imaging subsystem. Activating the imaging system only above areas which are sought after to be captured by the camera and maintaining the satellite oriented in a substantially horizontal attitude when flying above other areas enables to reduce the total satellite drag and thus to reduce the satellites energy consumption (including propellant and/or electrical energy stored in batteries), since less energy is allocated for altitude maintenance. Accordingly, a given satellite orbiting the Earth at a given altitude consuming a given amount of energy can be configured with control system 300 as disclosed herein to enable the same satellite to orbit the Earth at a lower altitude without increasing energy consumption.
  • the presently disclosed subject matter includes a method and system which enable to balance between real-time energy requirements needed for operating various functional elements in the satellite and the currently available energy. Based on a comparison between the required energy and available energy it is determined which of the functional elements onboard the satellite can be activated and which are maintained inactive or turned off.
  • the term "functional element" includes any device or circuitry onboard the satellite including any subsystem onboard the satellite or part thereof as well as any other device or circuitry which is operated as individual component separated from a specific subsystem.
  • Some functional elements are essential functional elements which are needed for maintaining the satellite properly operative (alive). Essential functional elements cannot be turned off during satellite operation. Other functional elements are non-essential and include systems which are not necessary for maintaining the satellite alive (e.g. imaging subsystem).
  • a non-essential functional element before a non-essential functional element is operated it is determined whether under the current conditions of energy availability the operation of the functional element is allowed or denied. Alternatively or additionally, it is determined whether under the current conditions of energy availability, an operating functional elements should be allowed to continue and operate or whether its operation should be terminated or at least reduced if possible.
  • the decision as to whether or not allow the operation of one or more non-essential functional elements (herein below also “desired functional elements") onboard the satellite is in general dependent on the available residual energy resource which can be defined as the difference between the overall available energy and the energy required for operating the essential functional elements. Only if the difference complies with one or more predefined conditions (e.g. the difference is greater than a certain predefined value) operation of the desired functional elements is allowed.
  • AE is the overall available energy during a certain time interval (e.g. per orbit);
  • SE is the available stored energy (e.g. in the onboard batteries);
  • SOE is the available solar energy (during a certain time interval), which is the energy which can be produced from absorption of the Sun-light on the solar panels;
  • RE includes the required energy for operating the satellite (the essential functional elements) as well as the desired functional elements.
  • Fig. 6 is a flowchart illustrating a sequence of operations performed in accordance with examples of the presently disclosed subject matter. While operations in to Fig. 6 are described below with reference to functional elements in Fig. 3 this is done by way of example only and should not be construed as limiting.
  • control unit 320 can be configured to access onboard energy resources (e.g. power subsystem 311) and obtain the current level of batteries depletion.
  • the currently available solar energy SOE is determined. According to one example, SOE is determined based on the angle between the satellite orbit around the Earth and the direction of the Sun-light (herein below “beta ( ⁇ ) angle").
  • Fig. 5a is a schematic illustration showing the Earth (52) orbiting the Sun (50) along orbit 54.
  • the Sun illuminates the Earth in the direction indicated by arrow 58 (usually measured from the satellite- Sun vector).
  • Beta angle is the angle between the direction of the Sun illumination towards the Earth and the plane of the satellite's orbit around the Earth 56 (i.e. satellite orbital plane).
  • Fig.5b is another schematic illustration, demonstrating the same principles as in Fig. 5a.
  • ⁇ angle is close to 90° i.e. the satellite orbital plane is substantially perpendicular to the direction of the Sun-light illuminating the Earth.
  • the value of ⁇ angle can be assumed to be constant and can be re-calculated, for example once every few days.
  • the ⁇ angle can be calculated based on the orbit of the satellite determined in some inertial reference frame (determination of the ⁇ angle can be performed by onboard systems, for example control unit 320, based on information from GPS system) and the sun position is determined in the same reference frame (determined by onboard systems), ⁇ angle can be determined from a geometrical computation based on the satellite orbit and sun position as well known in the art.
  • ⁇ angle values closer to 90° (or -90 for negative ⁇ ) when the satellite is exposed to solar radiation it can maximize its SOE while maintaining either vertical or horizontal attitude (i.e. in ⁇ angle values closer to 90° (or -90 for negative ⁇ ) both low drag and high solar energy absorption can be obtained); the closer is the value of ⁇ angle to 0°, the less is the energy which can be generated by the solar panels, since the satellites spends more time in Earth's shadow.
  • the satellite can no longer maintain a horizontal attitude (with minimum drag force) and at the same time directs its solar panels towards the Sun.
  • the SOE value is calculated based on the value of ⁇ angle and the expected attitude of the satellite during the time the functional element is operating.
  • a respective battery recharging rate can be determined as well based, inter alia, on these two factors.
  • the camera is directed toward the Earth to enable capturing Earth satellite imagery thus normally positioning the satellite in a substantially vertical position.
  • the SOE value during the operation of the imaging subsystem can thus be calculated based on the value of angle ⁇ and the attitude of the satellite while operating the imaging subsystem.
  • the entire satellite energy balance can be simulated (at a certain ⁇ angle and a certain satellite attitude) and compute (e.g. based on a respective mathematical model) a respective battery discharge (depletion) level.
  • ground control unit 355 (which comprises a computerized device with a processing circuitry) in ground control station can be configured to execute the simulation and generate and transmit appropriate instructions to the satellite for controlling the operations of various onboard functional elements based on the result of this computation.
  • the simulation or part thereof can be executed at the satellite.
  • the mathematical model for calculating the battery depletion level includes the following components:
  • Model of the solar panels including data related to their size, and efficiency in which they convert solar energy to electric power.
  • Model of batteries including data related to their properties , state, and model of power distribution between satellite sub-systems and operational units.
  • the overall available energy AE is determined.
  • the overall available energy is based on SE and the calculated SOE.
  • RE can be calculated for example, by adding the estimated non-essential energy consumption (NEC) of the one or more desired elements to the essential energy consumption (EEC) which indicates the energy consumed by the essential functional elements onboard the satellite.
  • NEC estimated non-essential energy consumption
  • EEC essential energy consumption
  • Ion propulsion system is a type of propulsion system that belongs to the electric propulsion systems and consumes a significant amount of energy. Accordingly in satellites which operate ion thrusters to maintain the satellite in the correct orbit, more electric energy will be consumed during periods of time where the drag force is greater and ion thrusters are used more frequently.
  • NEC value (and RE value accordingly) is calculated while taking into consideration the estimated drag exerted on the satellite, and specifically the additional energy consumption resulting from the excessive drag.
  • the desired functional element is an imaging subsystem
  • the drag exerted on the satellite during the operation of the imaging device is determined and based on this drag the expected energy consumptions is determined as well.
  • Rho is atmospheric density, computed from a model, based on satellite orbit and solar activity
  • A- satellite cross section (silhouette on plane normal to velocity) computed from a model strongly affected by attitude v - satellite velocity with respect to atmosphere (computed from a model) the orbit is determined by solving the orbital equations of motion (which requires modelling of the relevant forces including the above drag force and solving Newton equations).
  • control unit 320 can be configured to estimate the drag force and the resulting energy consumption based on the expected attitude of the satellite during operation of the functional element in question.
  • the additional energy expected to be consumed during the operation of the functional element is added to the value of RE.
  • the duration and power of activation of the ion thrusters can be estimated for the purpose of maintaining altitude.
  • the power consumption of the ion thruster can thus be calculated and added to the power consumption of the functional element (e.g. imaging subsystem).
  • control unit 320 can be configured to generate instructions for activating the desired one or more functional elements based on the result of the processing.
  • control unit 320 can be configured to monitor the operation of various functional elements and determine whether or not to allow the functional elements to continue and remain operative or otherwise to shut down their operation.
  • Determining which of the functional elements can be operated from among the group of functional elements can be based on predefined decision logic stored in a designated data repository accessible to control unit 320. Various logical schemes can be used for determining activation priority of the different functional elements in the group.
  • Priority - operational priority of the functional elements indicating priority of the different functional elements in the group.
  • a functional element with higher priority is to be activated before a functional element with less priority.
  • Satellite position- the priority of the operation of certain functional elements may change, depending on the specific location of the satellite. For example, if the satellite is located over an area which is sought to be captured by the imaging subsystem, the priority of the imaging subsystem may increase while the satellite is located over these areas.
  • Planning of future operation includes for example, determining whether one or more desired functional elements can be activated at a certain time at the future. Planning of future operation can further include the calculation of various parameters (including for example, altitude and attitude of the satellite during the operation of the functional elements) which may be required in order make the operation of the functional element possible. After planning is completed appropriate instructions can be generated for operating the satellite and the relevant functional elements (e.g. subsystems) according to the planning.
  • control unit 320 can comprise prediction module configured for executing the related operations.
  • control unit 320 can be configured to access onboard energy resources (e.g. power subsystem 311) and obtain the current level of batteries depletion, calculate an estimated depletion rate on one hand and recharging rate (e.g. by the solar panel) and determine the estimated SE at the future time.
  • onboard energy resources e.g. power subsystem 3111
  • calculate an estimated depletion rate on one hand and recharging rate e.g. by the solar panel
  • Control unit 320 can be further configured to determine the estimated future SOE to be available at the future time (and deduce a respective battery recharging rate accordingly) based on the expected ⁇ angle and satellite attitude at the future time.
  • the estimated future available energy AE which is based on estimated future SE and the estimated future SOE, can then be determined.
  • Control unit 320 can be also configured to determine an estimated future required energy RE at the future time.
  • RE is calculated by adding the estimated energy consumed by all the functional elements which are intended to operate at the future time. As explained above, this depends on the expected altitude and attitude of the satellite at the future time, which affects the drag and thus affects the energy consumption.
  • Control unit 320 can be configured to generate instructions for activating the desired one or more functional elements based on the result of the processing.
  • some of the satellite's operational parameters can be adapted in order to increase the available energy. For example, assuming it is desired to activate the imaging subsystem at a future time, the altitude of the satellite at the time of activation can be increased in order to reduce the drag which is exerted on the satellite. A new RE value can be calculated and then a respective difference between AE and RE. This process can be reiterated until the difference between the values allows the activation of the desired one or more functional elements.
  • ionic propulsion ion thrusters
  • ionic propulsion is the efficient fuel consumption relative to other propulsion techniques such as chemical propulsion (efficiency can be defined as the impulse per unit mass flow of propellant).
  • efficiency can be defined as the impulse per unit mass flow of propellant.
  • the ionic propulsion cannot be activated. This is so, since the ionic propulsion also generates an undesired modification of the attitude, which must be very accurately controlled during imaging and thus using ionic thrusters may impede the imaging process.
  • the specific installation angle of the ionic thrusters is fixed (with respect to the satellite body).
  • the attitude of the thrusters is often chosen so that the thrusters can increase satellite velocity and by that increase satellite altitude only if the thrusters are activated while the satellite is in a specific attitude (usually horizontal attitude).
  • the thrusters would produce an undesired change to the satellite velocity if they are activated while the satellite is in a vertical attitude, or if the attitude of the satellite is changing (e.g. while the imaging subsystem is activated for capturing images) during thruster activation.
  • control unit 320 can be operatively connected to an ion-thrusters control unit 309 configured to control the operation of one or more ion thrusters installed in a respective satellite.
  • control unit 320 is configured to activate ionic propulsion at times when the imaging system is not operating and the attitude of the satellite is in a substantially horizontal position. Instructions to operate the ionic propulsion can be generated for example by control unit 320 and transmitted to the ionic propulsion unit.
  • control unit 320 controls the use of an ionic propulsion at times when the imaging system is not operating and the attitude of the satellite is in a substantially horizontal position.
  • Instructions to operate the ionic propulsion can be generated for example by control unit 320 and transmitted to the ionic propulsion unit.
  • a satellite is configured with solar panels configured to convert solar energy into electrical power.
  • the satellite control unit attempts to maintain satellite attitude that allows maximum exposure of the solar panels to the sun in order to maximize the solar energy which can be generated.
  • the need to maximize exposure of the solar panels to the sun goes hand in hand with the need to reduce the satellite's drag by maintaining the satellite in a substantially horizontal attitude.
  • the two requirements contradict and it becomes problematic to obtain both sufficient sun exposure on the panels and minimum drag. For example, this is so when the battery depth of discharge is significant (i.e. a significant part of the battery stored energy is depleted) and battery charging in the minimum drag attitude is insufficient to maintain satellite functionality.
  • the specific occasions during the orbit in which difficulty arises when attempting to obtain both sufficient sun exposure (i.e. sufficient for supplying the needed energy to maintain satellite functionality) on the panels and minimum drag includes the time of entry of the satellite into satellite day. This occurs when the sun is partially or completely blocked from the satellite by the Earth and the satellite progress along the orbit to be exposed to the sun once more.
  • the presently disclosed subject matter includes a method and system which enable to balance between the requirement of maximum sun exposure on the panels and requirement of minimum drag on the satellite.
  • minimizing the satellite drag is prioritized over the sun-light exposure to the solar panels.
  • the satellite control system is configured to prioritize minimizing the satellite drag and thus controlling the satellite attitude such that the drag is minimized on the expense that the absorption of sunlight at the solar panel is reduced.
  • the satellite control system is configured to maintain the attitude of the satellite to allow maximum sun-light exposure on the panels.
  • the satellite's attitude is adapted to obtain both minimum drag and maximum sun-light exposure.
  • the absorbed sunlight when in minimum drag attitude depends on ⁇ angle and on the particular location within the orbit; the attitude should depend on maximum allowed battery depth of discharge.
  • the satellite position in the orbit may allow it to go into minimum drag attitude and still obtain sufficient solar radiation for operating the satellite.
  • Control unit 320 can be configured to control the orientation of the satellite accordingly.
  • Control unit 320 When the Earth completely blocks the Sun from the satellite, Control unit 320 is configured to control the satellite's orientation to minimum drag (being in substantially in horizontal position).
  • control unit 320 During entry into satellite day, as the satellite is gradually exposed to the Sun, control unit 320 is configured to determine the orientation of the satellite based on the current battery depletion level and the ⁇ angle and generate instructions for adapting the satellite attitude accordingly. This is done while striving to reduce the drag as much as possible on one hand in order to avoid depletion of the battery below threshold and absorb as much solar energy as possible on the other hand. As explained above the closer is the ⁇ angle to 90°or -90 the more likely that it would be possible.
  • control unit 320 is configured to control the satellite attitude to maintain minimum drag and maximize exposure of the solar panels to the Sun. Notably, even at minimum drag attitude, there is one degree of freedom left (satellite rotation around velocity axis) and it can be used to increase the solar flux on the panels
  • the presently disclosed subject matter contemplates a computer program on a non-transitory computer memory device being readable by a computer for executing the method of the presently disclosed subject matter.
  • the presently disclosed subject matter further contemplates a machine-readable non-transitory computer memory tangibly embodying a program of instructions executable by the machine for executing the method of the presently disclosed subject matter.

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  • Engineering & Computer Science (AREA)
  • Remote Sensing (AREA)
  • Aviation & Aerospace Engineering (AREA)
  • Physics & Mathematics (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Power Engineering (AREA)
  • Astronomy & Astrophysics (AREA)
  • General Physics & Mathematics (AREA)
  • Plasma & Fusion (AREA)
  • Life Sciences & Earth Sciences (AREA)
  • Sustainable Development (AREA)
  • Control Of Position, Course, Altitude, Or Attitude Of Moving Bodies (AREA)

Abstract

La présente invention concerne un procédé informatisé et un système informatisé, qui permettent à des satellites d'acquisition d'images de réduire l'altitude de l'orbite du satellite pour actionner le sous-système d'imagerie à une distance plus proche de la Terre, tout en réduisant l'augmentation globale de la traînée exercée sur le satellite et en réduisant ainsi la consommation d'énergie du satellite.
PCT/IL2017/050518 2016-05-10 2017-05-10 Gestion d'énergie de satellite Ceased WO2017195202A1 (fr)

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IL245566 2016-05-10
IL245566A IL245566A0 (en) 2016-05-10 2016-05-10 Energy management in a satellite

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WO2017195202A1 true WO2017195202A1 (fr) 2017-11-16

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Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
RU2735655C1 (ru) * 2020-02-23 2020-11-05 Российская Федерация, от имени которой выступает Государственная корпорация по космической деятельности "РОСКОСМОС" Способ управления космическим аппаратом
CN116527105A (zh) * 2023-04-19 2023-08-01 北京理工大学 基于跨时隙调度的卫星数据下载时间最小化方法
WO2025226308A1 (fr) * 2024-04-21 2025-10-30 WildStar, LLC Simulation d'analyse de puissance d'un système satellitaire

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Publication number Priority date Publication date Assignee Title
GB1552086A (en) * 1976-11-23 1979-09-05 Matra Digital data-processing system
WO1994015834A1 (fr) * 1992-12-04 1994-07-21 Hughes Aircraft Company Regulation de la barre omnibus de distribution de courant dans un engin spatial par reglage de la position des panneaux solaires
EP2016755A2 (fr) * 2006-05-11 2009-01-21 Rafael-Armament Development Authority Ltd. Satellite en forme de missile sur orbite basse pour la surveillance terrestre électro-optique et d'autres missions
US20110261227A1 (en) * 2008-09-12 2011-10-27 Nikon Corporation Imaging apparatus

Patent Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB1552086A (en) * 1976-11-23 1979-09-05 Matra Digital data-processing system
WO1994015834A1 (fr) * 1992-12-04 1994-07-21 Hughes Aircraft Company Regulation de la barre omnibus de distribution de courant dans un engin spatial par reglage de la position des panneaux solaires
EP2016755A2 (fr) * 2006-05-11 2009-01-21 Rafael-Armament Development Authority Ltd. Satellite en forme de missile sur orbite basse pour la surveillance terrestre électro-optique et d'autres missions
US20110261227A1 (en) * 2008-09-12 2011-10-27 Nikon Corporation Imaging apparatus

Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
RU2735655C1 (ru) * 2020-02-23 2020-11-05 Российская Федерация, от имени которой выступает Государственная корпорация по космической деятельности "РОСКОСМОС" Способ управления космическим аппаратом
CN116527105A (zh) * 2023-04-19 2023-08-01 北京理工大学 基于跨时隙调度的卫星数据下载时间最小化方法
WO2025226308A1 (fr) * 2024-04-21 2025-10-30 WildStar, LLC Simulation d'analyse de puissance d'un système satellitaire

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