Spacecraft propulsion system and method of operating the same

Abstract

A method of operating a spacecraft propulsion system includes injecting electrons into the plasma surrounding the spacecraft before generating an ion stream and continuing to inject a sufficient amount of electrons after starting to generate the ion stream to keep the spacecraft at a positive potential. This method can be implemented in a single thruster. In a spacecraft with multiple thrusters, the same method can be implemented in each thruster. In the case where the propulsion system includes multiple thrusters, the method may include: operating at least one of the thrusters as a drive thruster and operating at least one of the thrusters as an auxiliary or "standby" thruster. The electron source of at least one auxiliary thruster can be operated before the generation of the ion stream to inject electrons into the plasma surrounding the spacecraft.

Description

Technical Field

[0001] The present invention relates to spacecraft propulsion. The present invention can be used for communication satellites and any other type of spacecraft. Background Art

[0002] Communication satellites orbiting the Earth are being increasingly used for various purposes, such as tracking the position of objects, surveying sites, and observing changing weather patterns. Satellites and other spacecraft are typically equipped with an electric propulsion system, including one or more ion thrusters. In an ion thruster, electrons are extracted therefrom to generate a positive ion vapor, ionizing a neutral gas. The ions are accelerated, for example, by a potential difference between consecutive grids to generate thrust. Subsequently, the extracted electrons are injected into the ion stream to re-neutralize the gas and disperse in space. Alternatively, ions can be extracted from a metal, in which case the generated electrons can be ejected to avoid the metal becoming negatively charged. In the art, the source of the extracted electrons is referred to as a neutralizer.

[0003] As used herein, the term "thruster" refers to a propulsion mechanism that includes an ion source for generating an ion stream to propel a spacecraft, and at least one electron source for emitting electrons, such as to neutralize the ion stream or balance the spacecraft charge.

[0004] Ideally, an orbiting satellite has state-of-the-art electronics and other equipment that needs to operate optimally and be extended for a period of time without replacement. Therefore, one of the problems in spacecraft development is to minimize damage to components during its operation.

[0005] Embodiments of the present invention are not limited to solving such problems and may include solutions to other problems. Summary of the Invention

[0006] In some aspects, the present invention provides a method of operating a spacecraft propulsion system including a thruster, wherein at least one thruster includes an ion source for generating an ion stream to propel the spacecraft and an electron source for emitting electrons. The method includes operating the thruster by controlling the electron source to limit the generation of electrons, gradually rising from a closed state of the electron source to an open state and / or gradually descending from an open state of the electron source to a closed state.

[0007] This control can be achieved by a control signal, which can be a voltage that gradually increases or decreases, such as a ramp. By controlling the electron source in this way, the rate of increase and decrease of electron emission is actively constrained, such as by the control signal, rather than being passively constrained by the capabilities of the electron source hardware components. The advantages of this active control are further explained with reference to specific embodiments of the present invention.

[0008] The methods further described herein may include injecting electrons into the plasma surrounding the spacecraft before generating an ion flow and continuing to inject electrons after the ion flow begins. The continuous injection is to keep the spacecraft at a positive potential relative to the surrounding plasma, although this may not be measurable in practice.

[0009] The method of the present invention can be implemented in a single thruster. In a spacecraft with multiple thrusters, the same method can be implemented in multiple thrusters.

[0010] If the propulsion system includes multiple thrusters, the method according to certain embodiments of the present invention may include operating at least one thruster as an auxiliary or "backup" thruster. The electron source of at least one auxiliary thruster can be operated before the ion flow is generated to inject electrons into the plasma surrounding the spacecraft. Subsequently, after the electron flow generated by the auxiliary thruster reaches a pre-set level, the ion source and electron source for driving the thruster can be operated to drive the spacecraft while continuing to operate the auxiliary thruster to inject electrons to keep the spacecraft at a positive potential. In this method, the ion source of the auxiliary thruster does not operate.

[0011] When operating multiple thrusters, any one or more of the following operations can be performed to supplement the gradual turn-on or turn-off of the electron source:

[0012] - Operating at least one Electron source of each driving thruster to inject electrons into the plasma surrounding the spacecraft before operating the ion source to generate an ion flow to propel the spacecraft;

[0013] - Continuing to operate at least one electron <!--2-->source of each driving thruster after the ion source of the same driving thruster has stopped operating After that, electrons are injected into the plasma;

[0014] - Starting to operate an Any electron source of a driving thruster before starting to operate the Each ion source (or operating successively) of the driving thruster to generate an ion flow to propel the spacecraft;

[0015] - After all ion sources have stopped operating <!--7-->(or stopped successively), continuing to operate All <!--6-->Drive the thruster to At least one electron source;

[0016] - Starting the ion sources to operate successively;

[0017] - Stopping the ion sources to operate successively.

[0018] A spacecraft propulsion system can be designed to implement the methods described herein. Accordingly, in another aspect, the present invention provides a spacecraft propulsion system including at least one thruster, the thruster including an ion source for creating an ion stream to propel the spacecraft and at least one electron source, and a controller configured to operate the propulsion system according to any of the methods described herein.

[0019] The method according to an embodiment of the present invention can be implemented in an existing spacecraft propulsion system. Accordingly, in another aspect, the present invention provides a computer-readable medium containing instructions that, when implemented in a processor of a spacecraft propulsion system controller, cause the system to operate according to any of the methods described herein.

[0020] It is understood that this abstract is provided to briefly introduce some concepts that will be further described in the "Detailed Description" section below.

[0021] This abstract is not used to determine the key features or essential features of the claimed subject matter or to determine the scope of the claimed subject matter. BRIEF DESCRIPTION OF THE DRAWINGS

[0022] Some embodiments of the present invention will be described by way of example, with reference to the following figures, in which:

[0023] Figure 1 is a schematic representation of a satellite component according to some embodiments of the present invention.

[0024] According to certain embodiments of the present invention, Figure 2 、 3 and 4 show different perspective views of the satellite.

[0025] Figure 5 is a schematic diagram of a thruster according to certain embodiments of the present invention.

[0026] Figure 6 are satellite components, and the voltages and capacitances that may be generated between them and the surrounding plasma due to the operation of the propulsion system according to certain embodiments of the present invention.

[0027] Figure 7 is a circuit model of the electrical properties of a satellite component and the surrounding plasma according to certain embodiments of the present invention.

[0028] Figure 8 is a graph showing the variation of the potential difference between a patch and a spacecraft frame over time according to an embodiment of the present invention.

[0029] Figure 9 is a timing diagram of the operation sequence of an electron source and an ion source according to certain embodiments of the present invention.

[0030] It is understood that the drawings are not necessarily drawn to scale. Detailed implementation manners

[0031] The present invention will be understood from the following detailed description of embodiments, which is intended to describe rather than limit. For the sake of brevity, some common features, methods, system programs, components, and circuits are not described in detail herein.

[0032] Embodiments of the present invention provide systems and methods for operating a spacecraft propulsion system. Before discussing these in detail, spacecraft components in which embodiments of the present invention can be implemented are described.

[0033] Figure 1 is a schematic representation of a satellite component according to some embodiments of the present invention. One-way solid arrows between components represent power connections, two-way solid arrows represent RF signal connections, and dashed lines represent data connections.

[0034] Some components are located on the satellite body, represented by rectangle 120, and some are located on the wing, represented by rectangle 130. Figure 1 The satellite shown includes a power supply 101 and a power distribution system 102. The power supply 101 and the power distribution system 102 supply power to a propulsion system 190, a propulsion controller 109, a computing system 103, and a communication system 104. The propulsion controller 109 is shown here as a separate item, but in fact it can form part of the computing system 103. Using control software implemented in one or more processors incorporated in the propulsion controller 109, or in response to received instructions, such as instructions from the computing system, the propulsion controller can be configured according to certain embodiments of the present invention to implement methods. When transmitting instructions from the computing system 103, the computing system can be considered to include the propulsion controller. One of the functions of the propulsion controller 109 can be to output control signals to the ion source and electron source of the thrusters in the propulsion system 190, which will be further described below.

[0035] The power supply 101, the power distribution system 102, the computing system 103, and the communication system 104 are collectively referred to as the satellite "bus" in this technology. The communication system 104 can include one or more antennas (e.g., located on the satellite body). Alternatively, the communication system 104 can send and receive signals through one or more antennas on the wing 130.

[0036] Figure 1 The power supply 101 and the power distribution system 102 shown can also supply power to one or more sensors, not shown, which can be located on the body 120. These sensors form part of what is known in the art as the satellite "payload". The number and types of sensors can vary according to the intended use of the satellite.

[0037] In the case of an earth observation satellite, the payload may include one or more radar antennas 106 or antenna arrays, which may be located on one or more wings 130. Each antenna 106 or antenna array may have an associated amplifier 107, and receive power from power source 101 (e.g., via power distribution system 102) via power distribution system 108. Both power distribution systems 102 and 108 may include control logic as is known in the art.

[0038] Components of a satellite may be damaged during use, and once the satellite is in orbit it is more difficult to determine the source of any damage. The inventors have hypothesized that one particular source of damage is when a component is subjected to excessive voltage or current or power dissipation, known as electrical overstress "EOS". This may result in immediate damage or failure, or may shorten the life of the component. Certain components may be more susceptible to damage than others, particularly certain antenna types.

[0039] Amplifier 107 has a two-way data communication link with computing system 103 (through power distribution system 108 in the illustrated example) and can be configured to send data, such as data related to received radar signals, to computing system 103. The data can be processed by communication system 103, for example, to provide earth contour data, which can then be output to communication system 104 for onward transmission. Alternatively, raw data can be output by computing system 103 to communication system 104 for remote computing system processing on Earth or on another satellite. Computing system 103 can send data, such as operating instructions, data requests, and other signals familiar to those skilled in the art, to amplifier 107 through power distribution system 108.

[0040] The communication system 104 may communicate with ground stations or other satellites using radio frequency communications, optical (eg, laser communications), or any other form of communication known in the art.

[0041] Figure 2 , Figure 3 and Figure 4 is a perspective view of a satellite 140 orbiting in space, which may include Figure 1 components, such as Figure 3 shown. Figure 2 The satellite includes a main body 110, Figure 1 Certain parts of the body 120 may be installed therein, or Figure 1 Certain components of the device may be mounted thereon. The body 110 is also referred to in the art as a "bus" because it may house or support bus components. The body 110 may also house one or more batteries. The body 110 may be partially enclosed, such as to house and protect components. A housing may be used to mount components. Figure 2In the example, the solar panel 150 is installed on a rectangular surface of the main body 110, and other solar panels 155 are connected to the panel by struts 115.

[0042] The satellite 140 includes a generally planar structure extending from the main body 110 in two opposite directions to provide two "wings" 160. The structure shown with the wings 160 is installed on or beside a rectangular surface of the main body 110. As Figure 4 shown, it is formed in sections so as to be foldable for transportation and unfoldable in use. The main body 110 and the wings 160 are collectively referred to herein as the spacecraft frame and have electrical characteristics further described below.

[0043] One or more antennas as described above may be installed on the satellite "wings". Figure 4 The antenna array 180 is shown removed from the satellite to illustrate and may include patch antennas known in the art. As is known in the art, other components may be installed on the wings, including power distribution components and amplifiers, examples of which have been described in the previous patent application GB - A - 2598793. It has been found that the power control switch of the satellite described in that earlier patent application is particularly prone to damage. Different components may be more prone to damage due to different component arrangements.

[0044] The satellite 140 has a propulsion system 190 for generating thrust to maneuver the satellite. The propulsion system 190 is Figure 3 most clearly visible in and in this embodiment is installed on the main body 110 on a surface opposite to the solar panel 150.

[0045] As Figure 3 shown, the propulsion system 190 includes a plurality of thrusters 205, 210, 215, 220 which generate thrust for maneuvering the satellite 140 when required. Figure 3 The plurality of thrusters 205, 210, 215, 220 shown are located at the corners on one side of the main body 110 and may be equally spaced apart. However, in certain embodiments of the present invention, the propulsion system may have a different structure.

[0046] Figure 3 Each thruster shown may include an ion source for generating an ion stream to propel the spacecraft, and at least one electron source for injecting electrons into the ion stream to neutralize the ion stream. Embodiments of the present invention may include any thruster of this general type. Figure 5 An example is shown.

[0047] Figure 5The thruster shown includes an ion source 501, also known as an emitter, for creating an ion stream 503. In this example, two electron sources 505 are arranged in the form of filaments 507 on opposite sides of the ion source 501. As shown, one electron source is operating to emit electrons, namely the so-called neutralizer beam current 509, to neutralize the ion stream 503.

[0048] An extractor 511 and an extractor grid 513, known in the art and suitable for a suitable voltage, are used to attract ions out of the ion source 501. This will create a thrust and a reaction force 517 in the direction of arrow 515.

[0049] known to those skilled in the art Figure 5 The other components shown are not part of the present invention and include a heater 519, a reservoir 521, and a printed circuit board 525. The entire thruster assembly is housed in a housing 550.

[0050] It has been found that the operation of the neutralizer, i.e., the electron source 505, may cause the spacecraft to charge rapidly relative to the ambient plasma. By ejecting many negatively charged electrons, the spacecraft may have an overall positive charge as it now has more positive ions. In particular, this charging can occur very quickly, e.g., within 1 millisecond, if the electron source 505 is simply turned on and allowed to reach its on state at the speed allowed by the electron source hardware. This rapid charging can damage sensitive electronic components on the spacecraft, especially but not limited to antenna arrays and other components mounted in remote locations away from the satellite body, whether close to the thruster or not, such as the ends of the wings. The same effect occurs in reverse when the power is turned off.

[0051] Therefore, according to some embodiments, the rate of transition of the electron source from OFF to ON is controlled to constrain the emission of electrons, thereby controlling the rate of charge of the spacecraft frame. Similarly, the rate of transition of the electron source from ON to OFF can be limited to avoid an immediate stop in production.

[0052] Before further describing embodiments of the present invention in detail, first refer to Figure 6 discuss in more detail the possible causes of damage.

[0053] Figure 6 is a schematic diagram showing the components of a satellite 140 and the voltages and capacitances that may be generated between them and the surrounding plasma due to the operation of the propulsion system 190. Specifically Figure 6 shown is the propulsion system 190, the spacecraft frame denoted as 601, electronic components such as power switch elements denoted as 612, a metal structure 610 such as traces or areas on the satellite such as the antenna array 180, and the surrounding plasma 600.

[0054] According toFigure 6 , the spacecraft frame 601 is coupled to the metallic structure 610 through one or more electronic components 612 and the propulsion system 190. The potential difference that may occur between the metallic structure 610 and the spacecraft frame 601 is denoted as V1, and its associated capacitance is C1, while the potential difference that may occur between the metallic structure 610 and the plasma 600 is V2, and its associated capacitance is C2. The sum of V1 and V2 is denoted as V3. These and other electrical effects can be modeled according to individual discrete simulation elements, for example, using the well-known modeling program SPICE, as Figure 7 shown.

[0055] The current inventors have hypothesized that the cause of component damage and possible failure is due to the high voltage difference generated by the rapid change in the state of the propulsion system components from ON to OFF, particularly the electron source, but also including the ion source. In particular, the electron ejection of the spacecraft frame 601 causes a change in the potential of the frame 601 relative to the ambient plasma 600. At the same time, there is a natural (parasitic) capacitance C2 between the metallic structure 610, such as antenna patches and other metallic traces, and the ambient plasma 600. This capacitance C2 can hold the potential V2 of these metallic structures relative to the plasma for a short period of time. This may cause a change in the potential of the metallic structure 610, that is, an increase in V3, while the metallic structure maintains its potential, that is, V2 does not change relative to the plasma 600. This causes a voltage difference V1 between the metallic structure 610 and the patch. Such a voltage difference can damage components due to excessive electrical stress.

[0056] The key here is that the parasitic capacitance C2 holds the potential V2 constant for a short period of time. If the rate of change of the frame potential is slow enough (slower than the capacitance C2 can "accommodate it"), the risk of damage to the electronic component 612 will be reduced.

[0057] Figure 7 A SPICE circuit model showing in more detail the electrical characteristics of the components of the satellite 140 and the surrounding plasma is shown. Specifically, Figure 7 the components modeled in the shown circuit include the propulsion system 190, the spacecraft frame 601, the electronic component 612, the metallic structure 610 (patch antenna used to represent the metallic area and / or metallic traces), the surrounding plasma 600, and an additional protective sheath 620, for example, a thin film covering a multi-layer insulation blanket as known in the art. In space, the plasma is equivalent to the ground on Earth. Any one of the spacecraft frame 601, the metallic structure 610, and the sheath 620 can form one "side" of a capacitor. The patch is of particular interest because it is directly connected to the electrical components, and thus any voltage difference on the patch will affect the electronic components, including, for example, semiconductor switches that have been found to be vulnerable to damage.

[0058] As Figure 7As shown, the propulsion system 190 is modeled as a current source I1 connected to a pair of diodes, D1 and D2, arranged in a parallel configuration in opposite directions, and the entire propulsion system is grounded. The connection of the propulsion system 190 to the electronic components 612, that is, the electrical connection due to the spacecraft frame 601 is modeled as a 1 meter ohm resistor R1.

[0059] according to Figure 7 , the spacecraft frame 601 connects the propulsion system 190 to the electronics 612. According to the model, the capacitance between the frame and the plasma is modeled as a 22nF capacitor, CFX. The electronics are modeled as an RC circuit, where R represents the resistance between the frame and the patch, RFP, rated at 3GOhm, and C represents the capacitance between the frame and the patch, CFP, rated at 36pF. The patch is connected to the sheath and the plasma and is modeled as a 108nF capacitor, CPS.

[0060] In summary, Figure 7 middle:

[0061] -CFX stands for Capacitive Frame of Plasma

[0062] -CFP means Capacitance Frame to Patch, equivalent to Figure 6 C1

[0063] -RFP stands for Resistance Framework for Patch

[0064] -CPS stands for Capacitor Patch to Sheath

[0065] -CSX stands for Capacitive Sheath to Plasma

[0066] - Figure 7 (CPS+CSX) is equivalent to Figure 6 C2 in.

[0067] It should be understood that referring to Figure 6 The sheath, not discussed, is an additional source of capacitance that may cause the voltage to increase in the same manner as the metal structure 610. Figure 6 It can be seen from the description that if CSX>>CFP.

[0068] Figure 8 shows the change in potential difference between the patch and the frame over time. Figure 8 In the example, V(patch) is equivalent to Figure 6 V1, V(box) in is equivalent to Figure 6 V3 in the figure. Therefore, the figure shows the change of V2. When the electron source is turned on at 1.0Ks, V2 rises sharply, may reach 250V, and then decays. This voltage spike may cause damage to components or parts of components such as PIN diodes. Such damage has been observed in satellite operations.

[0069] However, it can be seen that as the equivalent capacitance dissipates its charge, the voltage decays over time. Thus, if the production rate of electrons can be constrained, no peak will occur because the voltage will not increase.

[0070] Thus, according to some embodiments of the present invention, controlling the electron source to constrain the generation of electrons gradually ramps up from the off state of the electron source to the on state.

[0071] Typically, an electron source is controlled by an on / off signal, so the time spent between the off and on states is only limited by the hardware, similar to turning on a light bulb. According to some embodiments of the present invention, the duration of the ramp-up is longer than the time required to turn on the electron source, whereby the generation of electrons is constrained beyond the limitations of the electron source hardware. In other words, the rate of increase of electron emission is actively controlled by the control signal rather than passively constrained by the hardware components of the electron source. In this way, the increase in electron flow between on and off can occur more gradually.

[0072] Reference Figure 6 and Figure 7 A similar effect as described is believed to occur when the electron source is turned off, as indicated by the negative voltage spike at ok in Figure 8 Thus, according to some embodiments of the present invention, the electron source can be limited to gradually ramp down from the on state to the off state rather than being allowed to suddenly stop generating electrons.

[0073] This control can be achieved through a control signal, which can be a voltage that increases or decreases gradually, such as a ramp. The control signal can be a digital control signal. Thus, according to some embodiments, the method can be implemented without modifying the hardware of the propulsion system. Alternatively, the control signal can be implemented by using hardware components.

[0074] For example, the on and off states of one or both of the ion source and the electron source can be defined as 90% and 10% of the maximum operating capacity of the source, which is standard in the art. In other words, at 10%, the power supply is considered off, and at 90%, the power supply is considered on.

[0075] The thruster can include multiple electron sources, such as as Figure 5 shown, in which case each electron source can be turned on in the same manner, optionally one after another, to further avoid a rapid increase in electron flow.

[0076] To mitigate the damage caused by voltage spikes, some additional techniques for controlling the propulsion system are also proposed here, which can be used alone or in combination. They can be summarized as:

[0077] - By starting to generate “oversupplied” electrons before ion supply and / or continuing to generate them after ion supply stops,

[0078] - Controlling the ion source of the thruster in a manner similar to an electron source so that it gradually rises from the ON state to the OFF state and / or gradually falls from the ON state to the OFF state, and

[0079] - Operating an additional electron source, herein referred to as an auxiliary electron source, e.g., from a non-operating thruster - this may be used for “oversupply” and has the additional benefit of avoiding a step change in the neutralization current of the neutralizer in an operating thruster in the event of an emergency shutdown.

[0080] Thus, according to certain aspects of the present invention, electrons are injected into the plasma surrounding the spacecraft before creating an ion flow that accelerates positively charged particles. This can be achieved by increasing the neutralization current, thereby creating an electron plume before starting to generate a positively charged accelerating ion flow. Additionally, electrons can continue to be injected into the plasma while generating a positively charged ion flow. It should be understood that, according to some embodiments, the number of electrons may exceed the number required to neutralize the ions from the ion source. Thus, by controlling the number of electrons injected into the plasma, e.g., within a driving thruster, or by turning on the auxiliary electron source before using the driving electron source, the spacecraft can be maintained at a positive potential relative to the surrounding plasma and potentially reduce damage to sensitive components of the spacecraft.

[0081] Regardless of whether electrons are injected before the ion source starts operating, the rise or fall of the electron source itself helps to avoid a sudden rate of change in potential that could damage spacecraft components.

[0082] Taking Figures 2 - 4 the propulsion system shown as an example, according to some embodiments, the rise or fall can be performed in any neutralizer of any of the thrusters 205, 210, 215, 220. When there are multiple thrusters, the total number of thrusters can be considered when determining the number of electrons to be provided.

[0083] If there is no additional electron source for “oversupplying” electrons, this can be achieved by operating one or more thrusters at an ion emission rate lower than the electron emission rate, e.g., 75% ion emission and 100% electron emission. The ratio in practice may be determined by experimentation and / or modeling and will depend on specific operating conditions such as, but not limited to, plasma density, temperature (i.e., kinetic energy, i.e., the speed at which electrons move), and spacecraft area. The amount of “oversupply” may be a few milliamperes and will be suitable for certain sizes of spacecraft and orbits (larger spacecraft collect more electrons, and lower orbits have denser plasma).

[0084] When an additional electron source is used for an "excess supply" of electrons, this can come from a non-operating thruster. Thus, in the operation of multiple thrusters, one or more can be selected as drive thrusters and one or more can be selected as auxiliary thrusters. The auxiliary thrusters can be selected to provide only electrons and thus not use their thrust capabilities. The advantage of this selection is that, due to the additional electron source, such as an auxiliary thruster, performing a simpler function, the risk of accidental functional failure (reset or other failure) is smaller.

[0085] Figure 9 A timing diagram is shown that illustrates control signals applied to an electron and an ion source to achieve an operating sequence according to some embodiments of the present invention. In Figure 9 this, the auxiliary electron source is represented as a backup neutralizer and the drive thrusters are represented as modules 1, 2, 3. It can be seen that the control signals applied to the electron source and the ion source rise and fall over a period of time to limit the generation of electrons or ions to gradually rise from an on state to an off state and vice versa. This is contrary to an on / off switch that switches from on to off state faster than the item being controlled switches from on to off or vice versa, so the rate of change of the flow is only limited by the item's hardware.

[0086] Generally, the sequence has the following new features, which can be implemented individually or in any combination:

[0087] - Operating at least one electron source of each <!--8-->drive thruster to generate an ion flow to propel the spacecraft while operating the electron source Front <!--9--> Inject electrons into the plasma around the spacecraft - it can be seen that for each of the modules 1, 2, 3, The neutral beam current is turned on before the "thrust" or ion flow;

[0088] - Continuing to operate at least one electron source of each <!--10-->drive thruster after the operation of the ion source of the same drive thruster has stopped After <!--11-->, injecting electrons into the plasma - it can be seen that for each of the modules 1, 2, 3, the neutral beam current It is turned off after the thrust is turned off;

[0089] - Before starting to operate the ion source of any drive thruster <!--13-->, starting to operate All <!--12-->that drive the thrusters One electron source (which can operate successively) - it can be seen that before any of the modules 1, 2, 3 thrusters are turned on, all the beams of modules 1, 2, 3 are turned on, and the beam currents are turned on successively before the thrusters are turned on in sequence;

[0090] - Continuing to operate at least one electron source <;14>; of the drive thruster <;15>; after <; / 15>; the operation of all ion sources has stopped, optionally stopping the operation one by one - it can be seen that all thrusters of modules 1, 2, 3 are turned off before the neutralizer beam of modules 1, 2, 3 is turned off, and the thrusters are turned off in sequence before the neutralizer beam is turned off in sequence;

[0091] - Starting to operate sequentially Ion <!--16-->Source - Module 1, 2, 3 beam currents start sequentially;

[0092] - Stop running in sequence Ion <!--17-->Source - The beam currents of modules 1, 2, and 3 are stopped in sequence;

[0093] - Select at least one thruster as an auxiliary thruster and operate Auxiliary thruster The ion source; start before any ion or electron source driving the thruster runs and / or stop after all ion or electron sources driving the thruster stop running - the standby neutralizer is turned on before any module 1, 2, 3 beam current or thrust and turned off after any module 1, 2, 3 beam current or thrust.

[0094] According to Figure 9 the order, switch the control signals of the beam current respectively, that is, the start and stop of the operation, rise or fall each module and the backup neutralizer during the TN period. The ramp-up or ramp-down time required can be calculated using the standard current model. In Figure 9 this, it is shown that the switching of all neutralizer beam currents is the same, but this may not always be the case, and it may vary between on and off and from one neutralizer to another. Similarly, the switching of the thrust (ion source) is on / off during a period T I for each module, but it may vary between on and off and from one electron source to another. According to some embodiments, the duration of the rise and / or fall of the ion source or electron source can be greater than 1 second, optionally greater than 5 seconds or even greater than 10 seconds. In principle, there is no upper limit to the rise or fall period of the ion source or electron source, because the slower the speed to reach this, the better. For practical purposes, the upper limit can be 10 minutes, or it can be 2 minutes or 1 minute. In Figure 9 the embodiment shown, the ramp period of the ion source is shorter than that of the electron source.

[0095] In Figure 9 it can be seen that during a period D, all modules or drive thrusters and all neutralizers including the backup are in the on state. D represents the nominal thrust time, which is determined by the flight time required by the spacecraft.

[0096] As pointed out elsewhere, some embodiments of the present invention relate to selecting one or more thrusters as auxiliary thrusters. It only works as an electron source and does not emit ions. When needed, the selection of the auxiliary thruster may vary depending on the drive operation, for example, to evenly distribute the wear amount of the thrusters. Additionally, the auxiliary thruster can be selected based on its proximity to known or discovered components that get hot, such as spacecraft subsystems. The thruster operating as an auxiliary thruster generates less heat and thus does not produce additional unwanted heat.

[0097] In embodiments of the present invention, the orbit or orbital path of the satellite is not restricted and may, for example, include any geostationary orbit (GEO), low Earth orbit (LEO), medium Earth orbit (MEO), polar orbit, and sun-synchronous orbit (SSO), transfer orbit, and geostationary transfer orbit (GTO), as well as Lagrangian points (L-points). However, in some embodiments of the present invention, those skilled in the art will readily understand that it may be necessary to test the satellite in a laboratory (e.g., a test laboratory on the ground) before launching it into space.

[0098] Embodiments of the present invention can be used to mitigate damage to various parts of the satellite, not just to the specific components mentioned above. Any of the components discussed below may also be vulnerable to damage, thus limiting the functionality of the satellite.

[0099] As Figures 2 - 4 shown, the antenna or antenna array 180 on the satellite 140 is configured to receive or transmit radio frequency signals sent to or from them through a radio frequency connector (not shown). These may be standard radio frequency connectors, such as snap-on or push connectors. The connector connects the antenna or antenna array to one or more amplifiers (not shown) on the circuit board, where the radio frequency signals are amplified. In the receive mode, the amplified signal can be output by the amplifier to the computing system. The sensor signal can be encoded in the computing system or communication system for transmission.

[0100] In an Earth observation satellite, the sensor may include a radar antenna. As is known to those skilled in the art, the sensor may also include any image capture device, temperature sensor, etc. The power supply may include power storage, for example, in the form of one or more batteries, the purpose of which is to enable the satellite to operate under low sunlight conditions. This is useful, for example, if it is necessary to continuously monitor the equipment carried by the satellite.

[0101] The satellite according to some embodiments of the present invention may also include systems not further described herein, such as but not limited to a thermal control system, an attitude control system, to ensure that the satellite points in the correct direction.

[0102] The term "computing system" is used herein to refer to any device or group of devices having processing capabilities such that it can execute instructions. Those skilled in the art will recognize that such processing capabilities are incorporated into many different devices, and thus the term "computing system" used herein may include personal computers, servers, and many other devices.

[0103] Unless otherwise stated, the components described herein are not necessarily physically separated from each other, and the functions of the components shown in the figures may be distributed or shared among different or the same physical devices. For example, certain functions of the communication system may be performed by the computing system, and vice versa.

[0104] It will be appreciated that the benefits and advantages described above may relate to one embodiment or several embodiments. Embodiments are not limited to those that solve any or all of the stated problems or have any or all of the stated benefits and advantages.

[0105] Any reference to "a" item refers to one or more of these items. The term "comprising" is used herein to mean including the identified method steps or elements, but these steps or elements do not comprise an exclusive list, and the method or apparatus may comprise additional steps or elements.

[0106] As used herein, the terms "component" and "system" may include a computer-readable data storage that is configured with computer-executable instructions that, when executed by a processor, cause certain functions to be performed. The computer-executable instructions may include routines, functions, etc. It should also be understood that a component or system may be localized on a single device or distributed across multiple devices.

[0107] Furthermore, to the extent that the term "comprising" is used in the detailed description or claims, that term is intended to be inclusive in a manner similar to the term "including" as that term is interpreted when used as a transitional word in claims.

[0108] Example methods are illustrated in the figures. Although the methods are shown and described as a series of acts performed in a particular sequence, it is to be understood and appreciated that the methods are not limited by the order of the sequence unless otherwise stated. For example, certain acts may occur in a different order than described herein. Additionally, one act may occur concurrently with another act. Further, in some instances, not all acts are required to implement the methods described herein.

[0109] It will be understood that the description of the above embodiments is given by way of example only, and that various modifications may be made by those skilled in the art. The examples described above include one or more embodiments. Of course, it is not possible to describe every conceivable modification and change of the above-described apparatus or method for the purposes of describing the above aspects, but those of ordinary skill in the art will recognize that many further modifications and permutations of the various aspects are possible. Accordingly, the aspects described are intended to include all such alterations, modifications, and variations that fall within the scope of the appended claims.

Claims

1. A method of operating a spacecraft propulsion system, the spacecraft propulsion system including at least one thruster, the at least one thruster including an ion source for generating an ion stream to propel the spacecraft, and an electron source for emitting electrons from the spacecraft to neutralize the ion stream, The method includes operating the at least one thruster by controlling the electron source to limit the generation of electrons to perform one or both of the following: Gradually rising from a closed state of the electron source to an open state to gradually increase the electron stream; and Gradually descending from an open state of the electron source to a closed state to gradually reduce the electron stream.

2. The method according to claim 1 includes one or both of the following: Controlling the ion source of the at least one thruster to start generating ions after the electron source is in an open state, and Controlling the electron source to start descending after the ion source is in a closed state.

3. The method according to claim 2 includes controlling the ion source to limit the generation of ions to perform one or both of the following: Gradually rising from a closed state of the ion source to an open state, and Gradually descending from an open state of the ion source to a closed state.

4. The method according to any one of claims 1-3, wherein the duration of the upward slope and / or the downward slope is greater than 1 second.

5. The method according to any one of claims 1-3, wherein the duration of the upward slope and / or the downward slope is less than 10 minutes.

6. The method according to any one of claims 1-3, wherein the propulsion system includes a plurality of thrusters, each thruster including an ion source for supplying ions to propel the spacecraft and at least one electron source.

7. The method according to claim 6 includes controlling the operation of the ion source such that it starts after the electron sources of all thrusters are in an open state.

8. The method according to claim 7 includes controlling the ion sources of the respective thrusters to start operating one by one.

9. The method according to claim 8 includes controlling the electron sources of the respective thrusters to start operating one by one.

10. The method according to claim 9 includes controlling the electron sources of all the driving thrusters to continue operating after all the ion sources stop operating.

11. The method according to claim 10 includes controlling the ion sources of the respective thrusters to stop operating one by one.

12. The method according to claim 11 includes operating an additional thruster to generate electrons without generating ions, and operating the electron source of the auxiliary thruster to perform one or both of the following: Starting before any of the ion or electron sources of the driving thrusters operate, and Stopping after all the ion or electron sources of the driving thrusters stop operating.

13. A spacecraft propulsion system, including at least one thruster, the at least one thruster including an ion source for generating an ion stream to propel the spacecraft, and at least one electron source for emitting electrons from the spacecraft to neutralize the ion stream, and a controller configured to operate the propulsion system according to the method of any of the preceding claims.

14. A computer-readable medium comprising instructions for implementation in a spacecraft propulsion system controller, the spacecraft propulsion system including at least one thruster, the at least one thruster including an ion source for generating an ion stream to propel the spacecraft and an electron source for emitting electrons from the spacecraft to neutralize the ion stream, wherein the instructions, when implemented in a processor of the spacecraft propulsion system controller, cause the system to operate according to the method of any one of claims 1 to 12.

Images