Propulsion and a satellite system working on atmospheric particles

Abstract

Disclosed is an electric propulsion system (110, 210) utilizing particles collected from atmosphere, the electric propulsion system having a first end (112) facing, when in use, incoming flow of particles (102, 202) and an exit aperture (252) at a second end (114) to eject accelerated particles to generate thrust, the electric propulsion system comprising an active intake module (220, 320A, 320B, 320N) having an opening (222, 322A, 322B, 322N), an active compression mechanism (224) and a first output (226); a guiding channel (230, 330) comprising a splitter (332, 432); the ion thruster; and the cathode.

Description

[0001] PROPULSION AND A SATELLITE SYSTEM WORKING ON ATMOSPHERIC PARTICLES

[0002] TECHNICAL FIELD

[0003] The present disclosure relates to electric propulsion systems, specifically those that utilize atmospheric particles for propulsion. The system is applicable to spacecraft and artificial satellites operating in low earth orbits or other planetary atmospheres.

[0004] BACKGROUND

[0005] Satellites operating in ultra or very low Earth orbits (ULEO and VLEO) or similar low-altitude orbits around other planetary bodies are often subjected to significant atmospheric drag due to the presence of atmospheric particles. Over time, this drag causes the orbit of satellites to decay, reducing their operational lifespan.

[0006] Typically, traditional electric propulsion systems require onboard fuel or propellant to generate thrust and counteract this drag. These systems rely on stored propellant, which imposes limitations on mission duration and adds significant mass to the satellite. Once the stored propellant is depleted, the satellite can no longer maintain its orbit and eventually deorbits. This causes short lifetime of the satellites at these orbits which makes ULEO / VLEO satellites impractical. However, the emerging technology of atmosphere-breathing electric propulsion where ion engines can be operated using the atmospheric particles provide a solution for continuous satellite refueling and long-term operation in ultra-low orbits.

[0007] However to date there are no successful demonstration of this type of engine due to various challenges. The use of conventional Gridded Ion Engines (GIE), Hall Effect Thrusters (HET) or similar ion propulsion systems often faces the challenges of cathode-neutralizer degradation due to exposure to atomic and molecular oxygen as well as low particle density at these altitudes.

[0008] Thus, due to exposure to the oxygen particles, existing configurations of Gridded Ion Engines and their typically used hollow cathode-neutralizers are not the right candidates to operate under the oxygen-rich environment of ultra-low orbits.

[0009] Moreover, the air-intake systems which collect atmospheric particles are passive intake systems, which rely solely on the satellite's structure to collect atmospheric particles, have limited efficiency. These systems often fail to collect a sufficient number of particles at higher altitudes and compress them, leading to insufficient pressures in the ionization chamber for plasma ignition. As a result, satellites equipped with such systems must operate at extremely low altitudes, where atmospheric drag is much greater, further exacerbating the problem.

[0010] Currently, over 99% of all gridded ion engines and hall effect thrusters which have been flown in space used a hollow cathode to neutralize the ion flow. While hollow cathodes have high performance, they employ thermionic electron emitters that are highly sensitive to oxygen, for example Lanthanum Hexaboride (LaB6) and Barium Oxide Impregnated Tungsten (BaO-W) are designed for use with inert gases of over 99% purity, however even minimal oxygen content will rapidly oxidise them and cause them to lose their electron emission properties. This possesses a significant challenge in operation in oxygen rich-environment of VLEO / ULEO.

[0011] Therefore, in light of the foregoing discussion, there is a need to overcome the aforementioned limitations to enhance the efficiency and operational lifespan of satellites by using atmospheric particles as a propellant, without suffering from rapid degradation of propulsion system components.

[0012] SUMMARY

[0013] The aim of the present disclosure is to provide an electric propulsion system configuration capable of efficient thrust generation by utilizing atmospheric particles collected from the environment and which will not suffer from severe degradation. The aim of the disclosure is achieved by an electric propulsion system which collects and compresses particles from the atmosphere of ultra or very low orbits, accelerates the particles, and ejects those to generate thrust, as defined in the appended independent claims to which reference is made. Specifically, this performance is achieved by using a Radio-Frequency oscillation in a cathode-neutralizer and thruster to ensure long-term operation in an oxygen-rich environment. Advantageous features are set out in the appended dependent claims. In addition, a further aim of the disclosure is achieved by an artificial satellite using an electric propulsion system which collects and compresses particles from the atmosphere, ejects those to accelerate the artificial satellite, as defined in the appended independent claims to which reference is made. Advantageous features are set out in the appended dependent claims.

[0014] The embodiments of the present disclosure substantially enable to improve the operational efficiency and lifespan of artificial satellites in ultra-low or very low Earth orbit (or other ultra-low or low-level orbits of one or more planets and / or a moon of the one or more planets) by providing a continuous source of propellant from atmospheric particles and reducing degradation of propulsion components. In other words, the electric propulsion system, through a synergistic operation of its components namely, an active intake module, a guiding channel, an ion thruster, and a cathode, forming a self-contained propulsion architecture that is capable of operating indefinitely in low orbit without stored fuel. Additional aspects, advantages, features, and objects of the present disclosure would be made apparent from the drawings and the detailed description of the illustrative embodiments constructed in conjunction with the appended claims that follow.

[0015] Throughout the description and claims of this specification, the words "comprise" , "include", "have", and "contain" and variations of these words, for example "comprising" and "comprises" , mean "including but not limited to" , and do not exclude other components, items, integers or steps not explicitly disclosed also to be present. Moreover, the singular encompasses the plural unless the context otherwise requires. In particular, where the indefinite article is used, the specification is to be understood as contemplating plurality as well as singularity, unless the context requires otherwise.

[0016] BRIEF DESCRIPTION OF THE DRAWINGS FIG. 1 is a schematic illustration of an artificial satellite equipped with an embedded electric propulsion system, in accordance with an embodiment of the present disclosure;

[0017] FIG. 2 is a schematic diagram of the electric propulsion system, in accordance with an embodiment of the present disclosure;

[0018] FIG. 3 is a detailed illustration of the electric propulsion system, showing additional components, in accordance with an embodiment of the present disclosure; and

[0019] FIG. 4 is a detailed illustration of the ion thruster and cathode, in accordance with an embodiment of the present disclosure.

[0020] DETAILED DESCRIPTION OF EMBODIMENTS

[0021] The following detailed description illustrates embodiments of the present disclosure and ways in which they can be implemented. The present disclosure provides an electric propulsion system that effectively uses atmospheric particles as propellant to generate thrust, significantly improving the lifespan and efficiency of artificial satellites in ultra or very low Earth orbit. By using atmospheric particles, the system ensures a continuous supply of propellant, reducing the need for stored onboard fuel and limiting the degradation of propulsion components over time. In a first aspect, the present disclosure provides an electric propulsion system. The electric propulsion system utilizes, when in use, particles collected from the atmosphere of ultra or very low orbits, the electric propulsion system having a first end facing incoming flow of particles and an exit aperture at a second end to eject accelerated particles to generate thrust. The electric propulsion system comprises an active intake module arranged at the first end, the active intake module having an opening, an active compression mechanism to collect and compress particles entering the active intake module via the opening, and a first output to provide the compressed particles. A guiding channel to provide a path for particles to flow from the first output to a first input of an ion thruster and a second input of a cathode, wherein the guiding channel comprises a splitter configured to direct a first flow portion of the compressed particles to the first input and to direct a second flow portion of the compressed particles to the second input. The ion thruster creates an accelerated ion beam of the compressed particles received via the first input and propels the accelerated ion beam from the exit aperture to generate thrust. The ion thruster is one of a radio frequency gridded ion thruster or microwave thruster, and the cathode neutralizes the ion beam, wherein the cathode is one of a radio frequency or microwave cathode.

[0022] The system operates by first using the active intake module to efficiently collect atmospheric particles as the artificial satellite moves through its orbit. The air-intake module includes mechanisms such as pumps, cryopumps, surfaces with moving liquids, compressors, moving parts or turbines that ensure particles are compressed to a sufficiently high density before being transferred through the guiding channel. Once the particles are compressed (i.e., particle density per cubic meter is increased), they are divided between the ion thruster and the cathode, where the ion thruster accelerates them, forming an ion beam to generate thrust, and the cathode neutralizes the ion beam, preventing any charge imbalance. In other words, the active compression mechanism of the active intake module increases the particle density per unit volume at the first output beyond ambient atmospheric conditions, enabling conditions suitable for ionization and plasma discharge. The use of radio frequency or microwave thrusters and cathodes ensures resistance to the oxygen-rich environment in the lower Earth orbit, preventing rapid degradation of key components, unlike conventional systems employing hollow cathodes. A synergistic combination of the active intake module, the guiding channel with the splitter, the ion thruster, and the cathode within a single electric propulsion system facilitates system-level efficiency, minimizing energy loss and ensuring synchronized particle processing across the thrust generation and neutralization subsystems.

[0023] The benefit of this aspect is the ability to maintain long-term artificial satellite operation by using a constant supply of atmospheric particles as propellant. This eliminates the limitations posed by conventional propulsion systems that rely on finite fuel stores. Additionally, the choice of specific thrusters and cathodes reduces wear and tear on the propulsion components, resulting in a longer operational lifespan for the artificial satellite.

[0024] The propulsion system utilizes preferably either a radio frequency (R.F) or microwave gridded ion thrusters and cathodes. These components are designed to resist degradation caused by exposure to oxygen-rich atmospheric particles, particularly at low Earth orbit altitudes where oxygen is a major component of the atmosphere. Conventional electric propulsion systems, such as those using hollow cathodes, degrade quickly in these conditions, significantly reducing their operational lifespan. The RF or microwave thrusters and cathodes used in the present disclosure, however, are much more robust, capable of long-term operation without substantial wear. This synergy between the RF / microwave ion thruster and cathode allows the system to maintain extremely low rate of erosion and have high efficiency over several years of operation, making this combination as the only suitable configuration for long term operation in the environment of ultra-low orbits.

[0025] According to embodiments of present disclosure a configuration which utilizes cathode and thruster which based on RF or Microwave discharge, enabling a long-term operation using atmospheric particles is provided. This in combination with RF Gridded Ion Thruster and active collection of atmospheric particles allows artificial satellites to keep the orbit in ultralow earth orbit without significant degradation.

[0026] A term "ultra or very low orbit satellite" refers to an artificial satellite that operates in a region of space known as Low Earth Orbit (LEO) or similar low-altitude orbits around other planetary bodies. This orbit typically ranges from approximately 150 kilometres to 2,000 kilometres above the surface of the Earth or other planetary atmospheres. Artificial satellites in low orbit experience more significant atmospheric drag compared to those in higher orbits, requiring effective propulsion systems to maintain their orbit.

[0027] "Particles" in the context of the present disclosure refer to the small atmospheric constituents collected by the artificial satellite's active intake module. These particles are typically composed of gases such as nitrogen, oxygen, and other trace elements found in the atmosphere. The particles are used as the propellant in the electric propulsion system, where they are ionized and accelerated to generate thrust. Indeed, particles which are present at atmosphere are considered in the present disclosure to refer to gas molecules and atoms present at altitudes associated with ultra and very low Earth orbit satellites. The gases could be considered to be mostly as rarefied atmospheric gases.

[0028] "Compression" refers to the process of increasing the density and pressure of the collected atmospheric particles within the active intake module. In this system, compression involves mechanisms such as pumps, turbines, or other moving parts that reduce the volume of the particles, thereby increasing their pressure to a level sufficient for use in the ion thruster and cathode. The goal of compression is to achieve a particle density that allows for effective ionization and thrust generation. Results of the compression is "compressed particles" i.e., particle density which is higher in comparison to "free space" around the operating environment of the electric propulsion system (and thus the artificial satellite). Term compressed particles does not, thus refer, that actual individual particle size (such as size of oxygen atom) would be reduced. "Active compression" refers to the process of actively collecting and compressing atmospheric particles using mechanical or fluidic components within the propulsion system. Unlike passive systems that rely solely on the structure of the intake to decelerate and collect particles, active compression involves the use of moving parts or dynamic systems such as pumps, turbines, pistons, or liquid-cooled surfaces. These components work together to increase the density and pressure of the collected particles to levels necessary for efficient propulsion. The active nature of the compression process allows for higher efficiency and control over particle intake and flow, improving the performance of the propulsion system.

[0029] The "guiding channel to provide a path for particles to flow" refers to tube system or channel via which particles can be transferred. In essence guiding channel is channel for the gas. The guiding channel connects fluidly the first output together with the first input. Optionally, the present disclosure provides an electric propulsion system wherein the active compression mechanism comprises at least one of the following: a liquid-cooled surface, movable surfaces, a piston, a pump, a turbine, and / or a servo motor. According to an embodiment, the active intake module can include one or more of these components to enhance the particle collection and compression process. For example, a pump or turbine may be used to generate a higher compression ratio, while liquid-cooled surfaces can reduce the temperature of the collected particles, increasing their density before compression. These components provide flexibility in optimizing the collection process depending on atmospheric conditions, thereby ensuring a stable supply of compressed particles to the ion thruster and cathode.

[0030] The benefit of this embodiment is the increased versatility and efficiency of the propulsion system in varying atmospheric conditions. The use of such mechanical or thermally-managed components ensures that compression is not reliant on passive aerodynamic effects, which are inefficient and not applicable in the low particle density environment of low Earth orbit. Consequently, this allows for reliable and scalable compression performance under varying external atmospheric conditions, thereby improving the operational robustness and efficiency of the propulsion system. By allowing different mechanisms for compression, the system can also adapt to changes in particle density, temperature, and flow rate, maintaining optimal performance. This adaptability also contributes to a more consistent thrust output and extends the operational life of the artificial satellite.

[0031] It will be appreciated that the splitter ensures that the compressed particles are appropriately divided between the ion thruster and the cathode. As an example, the splitter can consist of various materials or structures, such as porous elements or channels of different cross-sectional areas, that control the flow rates to both the thruster and cathode. This configuration allows the system to maintain a balanced distribution of particles, ensuring both the thrust generation and the neutralization process occur efficiently. In other words, the splitter ensures controlled allocation of compressed particles to the ion thruster and cathode, maintaining independent yet coordinated flow rates essential for sustaining both thrust and ion beam neutralization functions. A technical benefit of this is that a precise control over the particle flow is achieved, which optimizes the performance of both the ion thruster and the cathode. By ensuring that the appropriate amount of particles reaches each component, the system can maintain continuous and stable thrust, thereby improving the artificial satellite's ability to counteract atmospheric drag over an extended period of time

[0032] Optionally, the splitter comprises at least one of the following: at least one porous material, a set of tubes.

[0033] Based on this embodiment, the splitter in the guiding channel can be constructed from porous materials that control the flow of particles through the system. As an example, the porous material allows particles to pass through at a regulated rate, ensuring the right amount reaches the ion thruster and cathode. Alternatively, a set of tubes with varying diameters can be used to direct the particles into the different inputs of the system. These tubes could be designed to handle different flow rates based on the system's operational requirements, improving the precision of particle distribution.

[0034] The benefit of this arrangement is that it offers flexibility in the design of the particle flow control mechanism, allowing for an efficient and reliable way to split the particle stream between the ion thruster and the cathode. This enhances the overall efficiency of the system, ensuring that each component receives the proper amount of particles for thrust generation and ion beam neutralization. This is because the use of porous materials allows for diffusion-based regulation, while the set of tubes may enable diameter-based or length-based tuning of flow resistance. As a result, this facilitates a robust and predictable distribution of the compressed particles to the ion thruster and the cathode.

[0035] Optionally, the present disclosure provides, according to an embodiment, an electric propulsion system wherein the guiding channel is at least one of the following: a pipe, a tube, or a duct. I.e., a channel which can move flow of particles (which is flowing as a rarefied gas) between components of the system.

[0036] According to an embodiment, the guiding channel connecting the output of the active intake module to the inputs of the ion thruster and the cathode can be constructed using various configurations such as pipes, tubes, or ducts. As an example, a pipe may be used to provide a straight, smooth pathway for the compressed particles, minimizing resistance and preserving the particle flow rate. Tubes or ducts could be selected based on the specific requirements of the propulsion system, such as the desired particle velocity or the need to navigate through different structural components of the artificial satellite.

[0037] The benefit of this embodiment is the flexibility in designing the propulsion system to suit different operational environments. The use of pipes, tubes, or ducts allows for optimized flow control, reducing energy losses and ensuring that the compressed particles reach their respective destinations efficiently. In this way, the guiding channel ensures that a continuous and coordinated supply of compressed particles is maintained throughout operation, contributing to temporal stability of thrust and beam neutralization. This contributes to improved overall system performance and helps maintain consistent thrust levels.

[0038] According to additional embodiment, an electric propulsion system further comprises a gas damper arranged in the guiding channel between the first output and the splitter, employed to maintain a pre-determined particle flow rate to the cathode and the ion thruster.

[0039] According to the embodiment, the gas damper is arranged in the guiding channel to regulate the flow of particles between the active intake module and the splitter. As an example, the gas damper can act as a buffer that ensures the particle flow rate remains stable, preventing fluctuations that could negatively impact the performance of the ion thruster or the cathode. This allows the system to compensate for variations in atmospheric particle density, ensuring consistent thrust and ion beam neutralization.

[0040] The benefit of this embodiment is that it provides improved control over the particle flow, reducing the risk of performance disruptions due to sudden changes in atmospheric conditions. By maintaining a steady flow rate, the gas damper ensures that both the ion thruster and the cathode can function efficiently. As an example, ensuring that the flow of the particles is not too high or too low. This helps to keep thrust stable during usage. In other words, this addressed transient fluctuations in particle density or compressor performance, ensuring that downstream components receive a consistent mass flow rate. The inclusion of the gas damper thus stabilizes thrust generation and neutralization, improving temporal performance predictability and minimising thrust oscillations, which is critical for precise attitude or orbit control.

[0041] According to an embodiment, the gas damper further comprises a buffer volume. The gas damper in the guiding channel includes a buffer volume designed to store compressed particles temporarily. This buffer volume acts as a reservoir that regulates the particle flow towards the ion thruster and cathode. As an example, the buffer volume helps to absorb any fluctuations in particle collection from the active intake, ensuring a continuous and controlled flow of particles even when atmospheric conditions vary. The benefit of this embodiment is that the buffer volume provides a steady and reliable supply of particles, mitigating the effects of sudden changes in atmospheric density or particle intake efficiency. This ensures that the propulsion system can operate smoothly without interruptions, improving both the consistency of thrust and the overall efficiency of the system. In other words, the presence of the buffer volume facilitates temporal decoupling between particle intake and particle utilisation, thereby ensuring continuity in particle supply to the ion thruster and the cathode. As a result, the system maintains operational continuity and thrust stability even in the presence of environmental fluctuations, such as solar-induced atmospheric density changes.

[0042] According to one embodiment, the gas damper further comprises a first valve to control the release of particles from the buffer volume storage towards the ion thruster and the cathode. The valve system is integrated into the gas damper to regulate the flow of particles from the buffer volume. As an example, the first valve opens or closes depending on the particle demand of the ion thruster and the cathode, ensuring that the correct amount of particles is released from the buffer volume at any given time. This regulation is crucial in maintaining the desired flow rates and preventing either under- or over-supply of particles to the propulsion components.

[0043] The benefit of this embodiment is that it allows precise control over the release of particles from the buffer volume, enhancing the stability and efficiency of the propulsion system. By regulating the flow in real time, the valve system ensures that the ion thruster and cathode receive a consistent supply of particles, which helps in maintaining continuous thrust and effective ion beam neutralization. This is because a controlled release mitigates the risks of over-supply or under-supply to the ion thruster and cathode, thereby maintaining flow consistency and enabling closed-loop flow regulation. As a result, the propulsion system achieves improved responsiveness and stability under dynamic operating conditions.

[0044] According to one embodiment, an emergency gas injection system is arranged in fluidic connection to the guiding channel to inject an additional flow of particles when the pre-determined particle flow rate to the cathode and ion thruster is less than IO19particles per cubic meter, wherein the emergency gas injection system comprises an auxiliary propellant storage and a second valve between the auxiliary propellant storage and the guiding channel. As an example, the flow rate may be between 1017, 1018, or IO19to 1018, or IO19particles per cubic meter. The emergency gas injection system is preferably arranged to provide additional flow of particles to portion on the guiding channel between the gas damper and the splitter.

[0045] Based the embodiment, the emergency gas injection system is designed to ensure that the propulsion system can continue operating when the atmospheric particle density falls below a critical threshold. As an example, the auxiliary propellant storage contains a reserve of particles that can be injected into the system via the second valve when the particle flow rate from the active intake becomes insufficient. This system activates automatically when the particle density drops below the predetermined threshold of IO19particles per cubic meter, ensuring that the ion thruster and cathode receive enough particles to continue generating thrust and neutralizing the ion beam. The emergency gas injection system can be controlled by having a pressure gauge arranged in the guiding channel to measure particle flow in the channel or for example to measure amount of particles (partial pressure in practice) in the storage of the gas damper.

[0046] The benefit of this embodiment is that it provides an additional layer of reliability, ensuring that the propulsion system can function even during periods of low atmospheric particle density. This emergency backup ensures that the artificial satellite remains operational, maintaining its orbit and thrust capabilities, even in challenging environmental conditions. An inclusion of the auxiliary propellant storage and the second valve enables a fallback mechanism to maintain minimum operational particle density. This ensures continued thrust generation and ion beam neutralization during transient environmental events, such as density dips due to solar activity. The technical effect is enhanced fault tolerance and system uptime, which is vital for long-duration satellite missions in volatile orbital environments.

[0047] Optionally, the present disclosure provides, based on one embodiment, that the active air intake module is configured to generate a compression ratio of at least 500 for the collected particles.

[0048] According to an embodiment, the active intake module is designed to achieve a significant compression ratio, where the collected atmospheric particles are compressed to a ratio of at least 500 times their initial volume or pressure. As an example, this high compression ratio allows the propulsion system to gather a sufficient number of particles from the low-density atmosphere in low Earth orbit, enabling the ion thruster to generate adequate thrust while using only ambient particles from the environment. As an example, compression ratio can be 500, 600, 700, 800, 900, 1000 or more. Compression of particles in present disclosure refers to concentrating a given number N of particles to a smaller volume. As an example, if initial particle density is 1016particles per cubic meter, a compression ratio 1000 would result to a local density of 1019particles per cubic meter.

[0049] The benefit of this embodiment is that the high compression ratio enhances the efficiency of the propulsion system by ensuring that the collected particles are dense enough for reliable plasma ignition and thrust generation. This feature is particularly advantageous for artificial satellites operating in very low Earth orbits, where the atmospheric particle density is lower, allowing for effective operation even in such challenging conditions. Accordingly, this implementation supports efficient plasma generation and sustained operation without reliance on stored propellant, resulting in a propulsion system optimized for low-altitude, long-duration satellite deployment.

[0050] According to one embodiment, the electric propulsion system further comprises an igniter for initiating cathode discharge, the igniter comprising a plurality of electrodes.

[0051] Based on one embodiment, the igniter is integrated into the propulsion system to assist in the initial discharge of the cathode. As an example, the igniter consists of multiple electrodes that create a high-voltage breakdown when activated, generating a small plasma within the cathode's discharge chamber. This initial plasma provides the necessary seed electrons to trigger a stable and sustained cathode discharge, enabling the cathode to neutralize the ion beam effectively.

[0052] The benefit of this embodiment is the reliable ignition of the cathode, ensuring that the propulsion system can operate consistently even under varying environmental conditions. The use of multiple electrodes to generate a high-voltage breakdown facilitates seed plasma formation for subsequent R.F or microwave excitation. By incorporating an igniter, the system can rapidly initiate the cathode discharge process, reducing startup time and improving the overall responsiveness and efficiency of the propulsion system under variable thermal and electrical conditions. Based on one embodiment, the cathode and the ion thruster are equipped with a magnetic system to facilitate helicon discharge.

[0053] The magnetic system can be integrated into both the cathode and the ion thruster to enhance the ionization process through helicon discharge. As an example, the magnetic field generated by the system helps to guide and accelerate the electrons and ions within the discharge chambers, making the ionization process more efficient. This additional magnetic system ensures that the particles are ionized effectively, leading to higher thrust generation and improved neutralization of the ion beam. According to one setup the magnetic system is arranged to surround a first discharge chambers of the ion thruster and related radio frequency coil / antenna. According to additional or alternative setup the magnetic system is arranged to surround a second discharge chambers of the cathode and related radio frequency coil / antenna. As an example, the magnetic system can be configured to generate an external magnetic field (in respect to ion thruster and / or cathode) in various configurations. One example a coil around the discharge chamber. Another example is more complex system with multiple magnetic poles arranged around the chambers. In one example a magnetic coil (magnetic system) around the discharge chamber creates an axial magnetic field.

[0054] The benefit of this embodiment is the increased efficiency in both the thrust generation and ion beam neutralization processes. The presence of such a magnetic system enhances electron confinement and plasma density within the discharge chambers, thereby increasing ionization efficiency. This facilitates higher thrust-to-power ratios and reduced ion loss to chamber walls. In other words, the magnetic system enables better control over the ionization, resulting in more stable and powerful thrust production. This reduces energy consumption and extends the operational lifespan of the propulsion system, making it more effective for long-term artificial satellite missions.

[0055] Optionally, the present disclosure provides, according to an embodiment, that the ion thruster comprises:

[0056] a first discharge chamber to which the first flow portion of the compressed particles is fed from the first input, wherein the first discharge chamber is at least partly surrounded with a first radio frequency coil or a first antenna element to ionize particles in the first discharge chamber; and an ion optic system comprising a screen grid electrode and an accelerator grid electrode, wherein an electrical voltage is applied from a power source between the electrodes to accelerate the ionized particles to generate thrust.

[0057] Based on the embodiment, the ion thruster contains a first discharge chamber that receives the first flow portion of the compressed particles from the active intake module. Surrounding the discharge chamber is a radio frequency coil or antenna element that ionizes the particles by applying an alternating electromagnetic field. Once ionized, the particles are directed into an ion optic system where the screen grid electrode and the accelerator grid electrode, powered by an external voltage source, accelerate the ionized particles to high velocities, thereby creating the ion beam that generates thrust. The power source can be arranged to provide voltage of about 2000V to the screen grid electrode and voltage of about -500V to the accelerator grid electrode. This way total potential of about 2500V is used to accelerate the ionized particles. Acceleration voltage can be from 1000, 1100, 1200, 1300, 1400, 1500, 1600, 1700, 1800, 1900, 2000, 2100, 2200, 2300, 2400, 2500, 2600, 2700, 2800, 2900, 3000, 3100, 3200, 3300, 3400, 3500, 3600, 3700, 3800, 3900V up to 1100, 1200, 1300, 1400, 1500, 1600, 1700, 1800, 1900, 2000, 2100, 2200, 2300, 2400, 2500, 2600, 2700, 2800, 2900, 3000, 3100, 3200, 3300, 3400, 3500, 3600, 3700, 3800, 3900 or 4000V.

[0058] The benefit of this embodiment is that the configuration of the discharge chamber and ion optic system allows for highly efficient ionization and acceleration of particles. By applying radio frequency energy and precise electrical control through the ion optics, the system achieves greater thrust while minimizing energy consumption. This efficient particle acceleration ensures that the artificial satellite can maintain its orbit with minimal degradation over extended periods of operation. This implementation ensures efficient conversion of neutral particles into a high-velocity ion beam through controlled plasma generation and electrostatic acceleration. The RF-based ionization improves component longevity in oxygen-rich environments, while the grid acceleration enables precise thrust vectoring and high specific impulse.

[0059] Bases on further embodiment, the cathode comprises a second discharge chamber to which the second flow portion of the compressed particles is fed, wherein the second discharge chamber is surrounded at least partly with a second radio frequency coil or a second antenna element to ionize particles in the second discharge chamber, an electron aperture to provide a path for electrons to neutralize ionized particles from the ion thruster, and an ion collector.

[0060] According to the embodiment, the cathode includes a second discharge chamber where a portion of the compressed particles is directed. This chamber is surrounded by a second radio frequency coil or antenna element, which ionizes the particles inside the chamber. Electrons are then emitted through an electron aperture, allowing them to neutralize the ion beam emitted by the ion thruster. Additionally, an ion collector is used to gather any excess ions, ensuring a balanced flow of electrons and ions, which is crucial for maintaining the stability of the ion beam. The ion collector is preferably electrically connected chassis (neutral I "ground") of the electric prolusion system (or an artificial satellite).

[0061] The benefit of this embodiment is the efficient neutralization of the ion beam, which prevents the accumulation of charge and ensures stable operation of the propulsion system. By using radio frequency coils or antenna elements, the cathode effectively ionizes the particles and facilitates the electron flow needed for neutralization, thus improving the overall efficiency and durability of the propulsion system during extended artificial satellite missions. Furthermore, this configuration facilitates oxygen-tolerant neutralization of the ion beam without reliance on erosion-prone hollow cathodes. The RF-based ionization chamber enables robust electron production, and the ion collector ensures system charge neutrality. The result is stable beam neutralization and extended cathode operational life in low Earth orbit.

[0062] Based on one embodiment, that a ratio between amount of particles in the first flow portion to amount of particles in the second flow portion is 10:1. As an example, the ratio can be from 3:1, 4:1, 5:1, 6:1, 7:1, 8:1, 9:1, 10:1, 11:1, 12:1, 13:1, 14:1, 15:1, 16:1, 17:1, 18:1, 19:1, 20:1, 21:1, 22:1, 23:1, 24:1, 25:1, 26:1, 27:1, 28:1, 29:1 up to 6:1, 7:1, 8:1, 9:1, 10:1, 11:1, 12:1, 13:1, 14:1, 15:1, 16:1, 17:1, 18:1, 19:1, 20:1, 21:1, 22:1, 23:1, 24:1, 25:1, 26:1, 27:1, 28:1, 29:1 or 30:1.

[0063] According to an embodiment, the system is configured to divide the compressed particles between the ion thruster and the cathode in a specific ratio, where the first flow portion directed to the ion thruster comprises ten part of the total particles, and the second flow portion directed to the cathode comprises one parts. This ratio ensures that the cathode receives a smaller of particles to maintain its neutralization function while the ion thruster generates thrust from a sufficient particle stream.

[0064] The benefit of this embodiment is that it optimizes the performance of both the ion thruster and the cathode by carefully controlling the particle distribution. This ensures that the ion thruster has enough particles to generate effective thrust, while the cathode receives an adequate flow to maintain neutralization without affecting overall system performance. The precise control over particle flow improves the stability and efficiency of the propulsion system. Moreover, the aforesaid ratio (namely, a predefined flow division) ensures that the ion thruster receives a majority share of the compressed particles to maximize thrust, while allocating a smaller but sufficient portion to the cathode for ion beam neutralization. The result is optimized resource utilization and stable system-level performance, particularly in environments where particle supply is limited and flow efficiency is critical to mission success.

[0065] In a further aspect, the present disclosure provides an artificial satellite comprising a body of the artificial satellite, and an electric propulsion system integrated with the body of the artificial satellite, the electric propulsion system having a first end facing the movement direction of the artificial satellite and an exit aperture at a second end to eject accelerated particles to generate thrust for the artificial satellite. The electric propulsion system comprises an active intake module having an opening, an active compression mechanism to collect and compress particles entering the active intake module via the opening, and a first output to provide the compressed particles. A guiding channel provide a path for particles to flow from the first output to a first input of an ion thruster and a second input of a cathode, wherein the guiding channel comprises a splitter configured to direct a first flow portion of the compressed particles to the first input and to direct a second flow portion of the compressed particles to the second input. I.e. the guiding channel fluidly connects the first output to the first input. The ion thruster creates an accelerated ion beam of the compressed particles received via the first input and propels the accelerated ion beam from the exit aperture to generate the thrust, wherein the ion thruster is one of a radio frequency gridded ion thruster or microwave thruster, and the cathode neutralizes the ion beam, wherein the cathode is one of a radio frequency or microwave cathode.

[0066] In this embodiment, the artificial satellite incorporates the electric propulsion system into its body, allowing it to utilize atmospheric particles collected during its orbit to generate thrust. The active intake module collects and compresses particles from the atmosphere, and the ion thruster accelerates these particles to produce thrust. The cathode neutralizes the ion beam, preventing charge imbalance and ensuring stable propulsion. This configuration allows the artificial satellite to operate in low Earth orbits without relying on stored propellant, significantly extending its operational life. The artificial satellite can be equipped with solar panels to generate electricity. The electricity can be stored in battery system of the artificial satellite which can be used to operate active intake module as well as ion thruster and cathode.

[0067] The benefit of this aspect is that it provides a sustainable propulsion method for artificial satellites, utilizing ambient atmospheric particles to generate thrust. This eliminates the need for onboard fuel, reducing the mass of the artificial satellite and extending its operational lifespan. Additionally, the use of radio frequency or microwave thrusters and cathodes minimizes component degradation, ensuring long-term reliability and performance.

[0068] When the artificial satellite integrates the electric propulsion system, benefits from all features previously discussed, including the active intake module, guiding channel, ion thruster, cathode, and various optional elements like the gas damper, buffer volume, and emergency gas injection system, are achieved by implementing the artificial satellite. As an example, the artificial satellite could include a propulsion system with a compression ratio of at least 500, ensuring that sufficient atmospheric particles are collected to generate the necessary thrust for long-term operation in low Earth orbit.

[0069] The benefit of this embodiment is the comprehensive integration of all the advantageous features described in the earlier embodiments related to the electric propulsion system into the artificial satellite. This ensures that the artificial satellite operates with maximum efficiency and longevity, benefiting from continuous propellant supply from atmospheric particles, while minimizing degradation of the propulsion components. The synergistic operation of the active intake module, the guiding channel with the splitter, the ion thruster, and the cathode within a single electric propulsion system facilitates system-level efficiency, minimizing energy loss and ensuring synchronized particle processing across the thrust generation and neutralization subsystems. This allows the artificial satellite to maintain its orbit effectively, even in challenging atmospheric conditions, without the need for traditional fuel storage. The electric propulsion system, by utilising particles collected from atmosphere and eliminating dependency on stored propellant, enables reduction in satellite mass and increases payload capacity or functional lifespan for satellites operating in low Earth orbits.

[0070] In some embodiments of the present disclosure, the electric propulsion system leverages an active air intake module to achieve high particle collection efficiency and compression ratios. Unlike passive intake systems, which rely solely on their structural design to decelerate and collect particles, active intakes incorporate mechanical components such as pumps, turbines, or moving surfaces to actively capture and compress atmospheric particles. Passive intakes typically achieve low compression ratios, often below 100, and a particle collection efficiency of approximately 30%, making them impractical for effective use in low orbits where atmospheric drag is significant. In contrast, the active air intake system of the present disclosure is capable of generating a compression ratio of 500 exceeding even 1000, significantly enhancing the density of collected particles for use in the ion thruster. This ensures that the system can maintain efficient propulsion in low Earth orbit or similar environments, where atmospheric particle densities are relatively low.

[0071] Additionally, atmospheric particle density can vary dramatically due to external factors, such as solar activity. During solar maximums or fluctuations in solar radiation, atmospheric density can increase or decrease by a factor of ten or more, as observed in the GOCE (Gravity field and steady-state Ocean Circulation Explorer of European Space Agency) mission. To mitigate the effects of these fluctuations, the propulsion system, according to embodiments of present disclosure, incorporates a buffer system that includes a damping mechanism. This system ensures a steady flow of particles to the ion thruster and cathode, even when the atmospheric particle density suddenly drops. In such cases, the emergency gas injection system can activate to supplement the particle flow from the auxiliary propellant storage, ensuring that the particle density remains above the critical threshold of 1019particles per cubic meter required for stable operation of the ion thruster and cathode. The buffer and emergency systems thus ensure that the propulsion system operates reliably, even in variable atmospheric conditions. This configuration allows stable propulsion performance even in the presence of fluctuations in atmospheric particle density, due to its active collection and compression mechanism delivering particles at a regulated density to downstream components.

[0072] In some embodiments, the ion thruster is configured to achieve a specific impulse sufficient to counteract the drag imposed by atmospheric particles on the artificial satellite. Atmospheric drag, caused by particles moving at orbital velocities (approximately 8 km / s), exerts a continuous force on the artificial satellite. The ion thruster must therefore accelerate particles to velocities that exceed the orbital speed by at least a factor of three— approximately 24,000 km / s— to generate enough thrust to offset this drag. The specific impulse of the propulsion system in the present disclosure can reach up to 10,000 seconds, far exceeding the capabilities of conventional Hall-effect thrusters, which typically have specific impulses below 2,000 seconds. This high specific impulse enables the system to generate sufficient thrust while consuming less power, thereby improving the overall efficiency of the artificial satellite. In some cases, the artificial satellite structure itself contributes to atmospheric drag, particularly through solar panels and lateral surfaces. The configuration of the propulsion system takes this into account, compensating for the drag created by non-ideal orientations of the artificial satellite relative to the atmospheric particle flow. Computational models, such as Direct Simulation Monte Carlo (DSMC) simulations, indicate that up to 40% of the total atmospheric drag can come from the artificial satellite's structure, particularly from exposed surfaces such as solar panels. By adjusting the thrust and exhaust velocity of the ion thruster, the present disclosure ensures that the propulsion system generates a net thrust capable of overcoming both the drag from the artificial satellite's solar panels and other structural elements, in addition to atmospheric drag.

[0073] The electric propulsion system of the present disclosure can be also referred as an atmosphere-breathing electric propulsion system. The configuration provides efficient thrust generation by utilizing atmospheric particles collected from the environment and which will not suffer from severe degradation. This is achieved by an electric propulsion system which collects and compresses particles from the atmosphere, the particles are ionized and then the ionized particles are accelerated. The created accelerated ions are further neutralizes as those are ejected to generate thrust. Specifically, this performance is achieved by using a Radio-Frequency oscillation in a cathode-neutralizer and thruster to ensure long-term operation in an oxygen-rich environment.

[0074] According to one aspect an electric propulsion system utilizing particles collected, during a use, from atmosphere of ultra or very low orbits is provided. The electric propulsion system having a first end facing, when in use, incoming flow of particles and an exit aperture at a second end to eject accelerated particles to generate thrust, the electric propulsion system comprising: an active intake module arranged at the first end of the electric propulsion system, the active intake module having an opening, an active compression mechanism to collect particles entering the active intake module via the opening, wherein a particle density per cubic meter is a first particle density, and a first output to provide a particle flow having an increased particle density of a second particle density;

[0075] a guiding channel to provide a path for particles to flow from the first output to a first input of an ion thruster and a second input of a cathode , wherein the guiding channel comprises a splitter, configured to direct a first flow portion of the particle flow having the second particle density to the first input and to direct a second flow portion of the particle flow having the second particle density the second input;

[0076] the ion thruster to create an accelerated ion beam of the particle flow having the second particle density received via the first input and to propel the accelerated ion beam from the exit aperture to generate the thrust, wherein the ion thruster is one of radio frequency gridded ion thruster or microwave thruster; and

[0077] the cathode, to neutralize the ion beam, wherein the cathode is one of radio frequency or microwave cathode.

[0078] The disclosed artificial satellite uses the same an atmosphere-breathing electric propulsion system which collects and compresses particles from the atmosphere, ejects those to accelerate the artificial satellite and effectively neutralizes exhausted ions.

[0079] According to one aspect an artificial satellite comprising

[0080] a body of the artificial satellite; an electric propulsion system integrated with the body of the artificial satellite, wherein the electric propulsion system is as described above. Said artificial satellite preferably harvested energy to for operations using solar panels integrated on the artificial satellite.

[0081] DETAILED DESCRIPTION OF THE DRAWINGS FIG. 1 is a schematic illustration of an artificial satellite 100 equipped with an embedded electric propulsion system 110, in accordance with an embodiment of the present disclosure. The artificial satellite 100 includes a body 111 that houses the electric propulsion system 110. The electric propulsion system 110 is oriented such that a first end 112 faces the incoming flow of atmospheric particles 102, while a second end 114 is configured to eject accelerated particles 104 from an exit aperture to generate thrust. As the artificial satellite 100 moves through its orbit, the electric propulsion system 110 continuously collects particles from the atmosphere and propels them via the exit aperture, maintaining the artificial satellite's orbit and compensating for atmospheric drag. The artificial satellite 100 comprises a set of solar panels 106 to collect energy (electricity) for the operations of the artificial satellite and the electric propulsion system.

[0082] FIG. 2 is a schematic diagram of the electric propulsion system 210, in accordance with an embodiment of the present disclosure. The electric propulsion system 210 includes an active intake module 220 located at the first end of the electric propulsion system. This module comprises an opening 222 through which atmospheric particles 202 enter. The active intake module 220 also features an active compression mechanism 224 designed to collect and compress (increase density of the particles in other words) the incoming particles. Once compressed, the particles are delivered to the system's guiding channel 230, which fluidly connects the first output 226 of the active intake module to both the ion thruster 250 (via a first input 252) and the cathode 240 (via a second input 242). The ion thruster 250 is responsible for accelerating the particles to generate thrust, while the cathode 240 neutralizes the ion beam, ensuring that the ionized particles 204 do not cause any charge imbalance.

[0083] FIG. 3 is a detailed illustration of the electric propulsion system 310, showing additional components, in accordance with an embodiment of the present disclosure. Herein, FIG. 3 shows an arrangement of the additional components within the electric propulsion system 310. According to an embodiment there can be 1 to N active intake modules. As an example, 1, 2, 3, 4, 5, 6, 7, 8, 9, 10, 15, 20, 30 or more. In the FIG. 3 number of active intake modules is three. The active intake modules 320A, 320B, and 320N are positioned at the first end of the electric propulsion system, each with respective openings 322A, 322B, and 322N for collecting atmospheric particles 302. The intake modules 320A, 320B, and 320N include an active compression mechanism 324 that compresses the particles before directing them to the guiding channel 330.

[0084] The guiding channel 330 is configured to provide a path for particles from the output of the active intake module 320 to both the ion thruster 350 and the cathode 340. Notably, the guiding channel 330 contains a splitter 332 that divides the compressed particle flow into a first flow portion 336, directed to the ion thruster 350, and a second flow portion 334, directed to the cathode 340. The splitter 332 ensures that both the ion thruster 350 and the cathode 340 receive the appropriate amount (flow) of particles for efficient operation. Additionally, optionally, the propulsion system 310 features a gas damper 370, positioned within the guiding channel 330 between the output of the intake module 320 and the splitter 332. The gas damper 370 is employed to regulate the flow of particles, ensuring a steady and controlled particle flow rate to both the ion thruster 350 and the cathode 340. The gas damper 370 includes a buffer volume 372 that temporarily stores compressed particles, mitigating any fluctuations in particle flow caused by variations in atmospheric conditions. This buffer volume ensures a continuous and stable supply of particles to the propulsion components.

[0085] Further, the gas damper 370 is equipped with a first valve 374 that controls the release of particles from the buffer volume 372 toward the ion thruster 350 and cathode 340. This valve system helps regulate particle flow, ensuring that both components receive the right amount of particles (flow of particles) for sustained operation.

[0086] An emergency gas injection system 380 is also integrated into the guiding channel 330. This system comes into play when the particle flow rate (or pressure measured by the pressure gauge 376) drops below a critical threshold. It consists of an auxiliary propellant storage 382 and a second valve 384, which injects additional particles into the system when the flow rate is insufficient, ensuring continuous operation even in challenging conditions. A pressure gauge 376 is arranged to measure pressure of the buffer volume 372. (Pressure is equivalent to number of particles per cubic meter as an example). This information can be used to control the first valve 374 as well as a second valve 384.

[0087] FIG. 4 is a detailed illustration of the ion thruster 450 and cathode 440, in accordance with an embodiment of the present disclosure. Herein FIG.

[0088] 4 shows specific components and their electrical connections. The ion thruster 450 includes a first discharge chamber 454, which receives a particle flow from the first input 452. Surrounding the discharge chamber 454 is a first radio frequency coil 456 (or antenna), which ionizes the particles which are inside the first discharge chamber 454. Once the particles are ionized, they are accelerated through an ion optic system composed of two grid electrodes: a screen grid electrode 458A and an accelerator grid electrode 458B.

[0089] The screen grid electrode 458A and the accelerator grid electrode 458B are powered by two separate voltage sources., A first power source 460 applies a positive voltage of+2000V to the screen grid electrode 458A, while a second power source 462 applies a negative voltage of -500V to the accelerator grid electrode 458B. The difference in potential (in this example 2500V) between the two grids accelerates the ionized particles, creating an ion beam that exits the thruster, generating thrust for the artificial satellite. This precise voltage control allows the system to achieve high-efficiency particle acceleration and stable thrust generation. The screen grid electrode 458A is arranged to be in contact with the plasma and the discharge chamber. The accelerator grid 458B is arranged in parallel to the screen grid electrode 458A at a distance d away from the discharge chamber.

[0090] The cathode 440 also plays a crucial role in the propulsion system by neutralizing the ion beam emitted by the ion thruster 450. The cathode 440 includes a second discharge chamber 444, where particles from the second flow portion are ionized using a second radio frequency coil (or antenna) 446. After ionization, electrons are emitted through an electron aperture 449, which neutralizes the ion beam. Additionally, the cathode is equipped with an ion collector 448, which is electrically connected to the ground / chassis to collect excess ions and maintain charge balance within the system. This connection ensures that any stray ions are safely neutralized, preventing any charge buildup that could interfere with the proper functioning of the propulsion system.

[0091] By utilizing the power sources 460 and 462 to control the potential difference across the grids 458A and 458B, and grounding the ion collector 448, the system ensures efficient ion acceleration and neutralization, enabling the propulsion system to generate continuous and stable thrust while preventing charge imbalances. In an embodiment the first power source is a first battery and a second power source is a second battery. The batteries are connected in a way that the positive terminal of the first battery is connected to the screen grid electrode 458A and the negative terminal of the first battery to ground / chassis. Further the negative terminal of the second battery is connected to the accelerator grid electrode 458B and positive terminal of the first battery to ground / chassis.

[0092] Provided figures are merely examples, which should not unduly limit the scope of the claims herein. A person skilled in the art will recognize many variations, alternatives, and modifications of embodiments of the present disclosure.

Claims

CLAIMS1. An electric propulsion system (110 210) utilizing particles collected from atmosphere of ultra or very low orbits, the electric propulsion system having a first end (112) facing, when in use, incoming flow of particles (102, 202) and an exit aperture (252) at a second end (114) to eject accelerated particles to generate thrust, the electric propulsion system comprising:an active intake module (220, 320A, 320B, 320N) arranged at the first end of the electric propulsion system, the active intake module having an opening (222, 322A, 322B, 322N), an active compression mechanism (224) to collect and compress particles entering the active intake module via the opening, and a first output (226) to provide the compressed particles;a guiding channel (230, 330) to provide a path for particles to flow from the first output to a first input (252, 452) of an ion thruster (250, 350, 450) and a second input (242, 442) of a cathode (240, 340, 440), wherein the guiding channel comprises a splitter (332, 432), configured to direct a first flow portion (336) of the compressed particles to the first input and to direct a second flow portion (334) of the compressed particles to the second input;the ion thruster to create an accelerated ion beam (104, 204) of the compressed particles received via the first input and to propel the accelerated ion beam from the exit aperture to generate the thrust, wherein the ion thruster is one of radio frequency gridded ion thruster or microwave thruster; andthe cathode, to neutralize the ion beam, wherein the cathode is one of radio frequency or microwave cathode.

2. The electric propulsion system according to claim 1, wherein the active compression mechanism comprises at least one of: a liquid cooledsurface, a movable surfaces, a piston, a pump, a turbine and / or a servo motor.

3. The electric propulsion system of claim 1 or 2, wherein the splitter comprises at least one of the following: at least one porous material, a set of tubes.

4. The electric propulsion system according to any of the preceding claims, wherein the guiding channel is at least one of the following: a pipe, a tube, a duct.

5. The electric propulsion system according to any of the preceding claims, wherein the electric propulsion system further comprises a gas damper (370) arranged in the guiding channel between the first output and the splitter, employed to maintain a pre-determined particle flow rate to the cathode and the ion thruster.

6. The electric propulsion system of claim 5, wherein the gas damper comprises a buffer volume (372).

7. The electric propulsion system of claim 6, wherein the gas damper further comprises a first valve (374), to control the release of particles from the buffer volume storage towards the ion thruster and the cathode.

8. The electric propulsion system according to any of the preceding claims, wherein an emergency gas injection system (380) is arranged in fluidic connection to the guiding channel to inject an additional flow of particles when the pre-determined particle flow rate to the cathode and ion thruster is less than 1019particles per cubic meter, wherein the emergency gas injection system comprises an auxiliary propellant storage (382) and a second valve (384) between the auxiliary propellant storage and the guiding channel.

9. The electric propulsion system according to any of the preceding claims, wherein the active air intake module is configured to generate a compression ratio of at least 500 for the collected particles.

10. The electric propulsion system according to any of the preceding claims, further comprising an igniter (442) for initiating cathode discharge, the igniter comprising a plurality of electrodes.

11. The electric propulsion system according to any of the preceding claims, wherein the cathode and the ion thruster are equipped with a magnetic system to facilitate helicon discharge.

12. The electric propulsion system according to any of the preceding claims, wherein the ion thruster comprises:a first discharge chamber (454) to which the first flow portion of the compressed particles are fed from the first input, wherein the first discharge chamber is at least partly surrounded with a first radio frequency coil (456) or a first antenna element to ionize particles in the first discharge chamber; andan ion optic system comprising a screen grid electrode (458A) and accelerator grid electrode (458A), wherein an electrical voltage is applied from a power source (460, 462) between the electrodes to accelerate the ionized particles to generate thrust.

13. The electrical propulsion system according to any of the preceding claims, wherein the cathode comprises a second discharge chamber (444) to which the second flow portion of the compressed particles are fed, wherein the second discharge chamber is surrounded at least partly with a second radio frequency coil (446) or a second antenna element to ionize particles in the second discharge chamber, an electron aperture (449) to provide a path for electrodes to neutralize ionized particles from the ion thruster and an ion collector (448).

14. The electric propulsion system according to any of the preceding claims, where in a ratio between amount of particles in the first flow portion to amount of particles in the second flow portion is 10:1.

15. An artificial satellite (100) comprising:a body of the artificial satellite (111);an electric propulsion system (110, 210) integrated with the body of the artificial satellite, the electric propulsion system having a first end (112) facing, when in use, movement of direction of the artificial satellite, and an exit aperture (252) at a second end (121) to eject accelerated particles to generate thrust for the artificial satellite, the electric propulsion system comprising:the active intake module (220, 320A, 320B, 320N) having an opening (222, 322A, 322B, 322N), an active compression mechanism (224) to collect and compress particles entering the active intake module via the opening, and a first output (226) to provide the compressed particles;a guiding channel (230, 330) to provide a path for particles to flow from the first output to a first input (242) of an ion thruster (250) and a second input (242) of a cathode (240), wherein the guiding channel comprises a splitter (332, 432), configured to direct a first flow portion (336) of the compressed particles to the first input and to direct a second flow portion (334) of the compressed particles to the second input;the ion thruster to create an accelerated ion beam of the compressed particles received via the first input and to propel the accelerated ion beam (204) from the exit aperture to generate the thrust, wherein the ion thruster is one of radio frequency gridded ion thruster or microwave thruster andthe cathode (240, 340), to neutralize the ion beam, wherein the cathode is one of radio frequency or microwave cathode.

16. An artificial satellite (100) comprisinga body of the artificial satellite;an electric propulsion system (110, 210) integrated with the body of the artificial satellite, wherein the electric propulsion system is according to any of the claims 1-14.

Images