266 results about "Ion thruster" patented technology

A plasma source for processing or imaging a substrate, for ion source for proton therapy, for ion thrusters, or for high energy particle accelerators includes a coolant circuit passing adjacent to a plasma ion reactor chamber and RF antenna coils. In a method for operating the plasma ion source having an induction coil adjacent to a reaction chamber for inductively coupling power into the plasma from a radio frequency power source, the method comprises pumping a dielectric fluid into contact with induction coils of the plasma ion source along the coolant circuit. Use of the dielectric fluid both electrically insulates the plasma chamber, so that it can be biased to 30 kV and up, and efficiently transfers heat away from the plasma chamber.

A rotatable thruster aircraft includes a fixed wing; rotatable thruster assemblies, each including first and second thrusters that provide rotation via differential thrust; and vertical lift thrusters, optionally connected via an elongated member; and an aircraft control unit, including a processor, a non-transitory memory, an input / output component, and a power manager that controls specific power applied to the thrusters.

The invention discloses a near-field plume mass-spectroscopic diagnostic E*B probe based on the Faraday cup and belongs to the technical field of plasma mass-spectroscopic diagnosis. The probe mainly applied to measuring near-field plumes of an ion thruster and of a Hall thruster comprises a central frame, ferrite permanent magnets, a flat electrode plate, an electrode plate holder, a collimator tube, a drift tube, a Faraday cup, six carbon steel shells and an anti-sputtering heat-insulating layer. According to the connectional relation, the central frame is used as a core part, the ferrite permanent magnets are distributed on upper and lower surfaces of the central frame, the electrode plate is fixed in the central frame, and an orthogonal electromagnetic field area is formed. The six carbon steel shells are used for packaging, and the front ends of the shells are coated with an anti-sputtering heat-insulating layer. The collimator tube of stainless steel and the drift tube are fitly fixed to the centers of two ends of the central frame through shaft holes. Ions different in valence are screened by adjusting voltage among the electrode plates, univalent and bivalent ion currents are acquired with the Faraday cup of aluminum, and the ratio of near-field plum bivalent ions is acquired by analytical computing.

A high density plasma generated by microwave injection using a windowless electrodeless rectangular slotted antenna waveguide plasma source has been demonstrated. Plasma probe measurements indicate that the source could be applicable for low power ion thruster applications, ion implantation, and related applications. This slotted antenna plasma source invention operates on the principle of electron cyclotron resonance (ECR). It employs no window and it is completely electrodeless and therefore its operation lifetime is long, being limited only by either the microwave generator itself or charged particle extraction grids if used. The high density plasma source can also be used to extract an electron beam that can be used as a plasma cathode neutralizer for ion source beam neutralization applications.

The invention discloses lifetime ground test equipment for an ion thruster. By employing the equipment, special requirements of high pumping ability, small return sputtering sediment, high reliability, and high equipment integration of test equipment for lifetime tests of the ion thruster can be satisfied. The lifetime ground test equipment for the ion thruster comprises a main cabin vacuum container, a gate valve, an auxiliary cabin vacuum container, a vacuum pumping system, an ion beam target, an anti-sputtering screen, a thruster mobile mechanism, a quartz crystal microbalance QCM, a thruster divergence angle measuring system, a grid corrosion on-line monitoring system, a ground test power supply system, an equipment control system, a xenon supply system, a repressing system, a cooling water system, a pneumatic element air supply system, a liquid nitrogen storage and supply system, and a shooting illumination system. By employing the equipment, real results of the lifetime of the ion thruster can be obtained, and real and reliable test data can be provided for the evaluation of the lifetime and the reliability of the thruster.

The invention discloses a radiation electromagnetic compatibility test method for an electric propulsion system and on-star radio frequency equipment. Firstly, an ion thruster is arranged in a wave transmitting secondary cabin; the wave transmitting secondary cabin is connected with a main vacuum tank; and electromagnetic radiation interference produced by a plume of the ion thruster is emitted out of the main vacuum tank through the wave transmitting secondary cabin. Each equipment and antenna are connected. When the electric propulsion system is started to normally work, both uplink level and downlink level of a responder are weak levels; conditions of uplink instruction reception and downlink remote metering error code are checked; if the instruction reception is correct, the code ratemeets the requirement and the responder is not interfered; remote metering parameters for a GPS (Global Positioning System) receiver are observed, and absolute positioning condition and relative positioning condition are determined; and if the remote metering is correct and the positioning is normal, the GPS receiver is not interfered.

An engine for providing rotary power about an output shaft with a high power-to-weight ratio includes a plurality of flow guiding blades mounted on the inner surface of an annular thruster base. The flow guiding blades cooperate with the peripheral surface of a rotor for forming a plurality of ramjet-like thrusters. The configuration of the flow guiding blades allows for optimization of the number of thrusters. The centrifugal forces generated by the rotating components is compensated by an annular reinforcement wall made with high strength materials allowing for downsizing of the rotor and associated components.

The invention discloses an ion thruster discharge chamber magnetic pole structure which comprises a lower magnetic pole (1), a middle magnetic pole (2), an upper magnetic pole (3), a lower pole shoe (4), a middle pole shoe (5) and an upper pole shoe (6). The lower pole shoe (4), the middle pole shoe (5) and the upper pole shoe (6) are correspondingly connected with the lower magnetic pole (1), the middle magnetic pole (2) and the upper magnetic pole (3). The structure further comprises permanent magnets (9 and 10). A magnetic force line loop is formed among the magnetic poles through the pole shoes (4, 5 and 6) and the permanent magnets (9 and 10). A loop tip cusped magnetic field is formed in a discharge chamber. The structure is characterized in that positive pole cylinders (7 and 8) of the discharge chamber are arranged in the cusped magnetic field, all the magnetic poles extend to the inner surfaces of the positive pole cylinders (7 and 8) and are charged with negative electricity relative to the positive pole cylinders (7 and 8), and a discharge chamber negative pole (11) installed on the lower pole shoe (4) directly extends into the discharge chamber under the circumstance without the positive pole cylinders ahead. The invention further discloses a design method for the ion thruster discharge chamber magnetic pole structure. When the structure is used, the primary electron utilization rate can be increased, and the discharging efficiency and the beam uniformity are improved.

The invention relates to an assembly type hollow cathode heater comprising a cathode pipe, a front positioning ring, a ceramic framework, a heater strip, a ceramic jacket, a rear positioning ring, a lead wire pipe and an emitting body, wherein the ceramic framework is sleeved at the excircle of the cathode pipe in a tightly matching way; the emitting body is sleeved in the cathode pipe; the ceramic jacket is sleeved on the ceramic framework wound with the heater strip; the front positioning ring and the rear positioning ring are welded on both ends of the ceramic framework and the ceramic jacket; one end of the heater strip is welded on the cathode pipe, and the other end of the heater strip is sleeved on the lead wire pipe. When the assembly type hollow cathode heater starts, voltage of 10-15V is applied between the lead wire pipe and the cathode pipe so that the heater strip generates current of 8-12 A and lasts for 3 minutes-5 minutes to heat the emitting body to 1500 K-1800 K, thus the emitting body has electronic emission ability. The assembly type hollow cathode heater has simple structure, easy manufacture and high finished product ratio and reliability and can be used for a xenon ion thruster, a Hoare thruster and a plasma source hollow cathode assembly.

The invention discloses a method and device for clearing an ion thruster grid of contaminants. A typical method includes applying a power supply to a grid to clear contaminants, monitoring an energy applied to the grid of the applied power supply and suspending application of the power supply to the grid after the monitored energy substantially reaches a predetermined value. A typical device includes a controller for applying a power supply to a grid to clear the grid of contaminants and a timer for monitoring an energy applied to the grid of the applied power supply and suspending application of the power supply to the grid after the monitored energy substantially reaches a predetermined value.

The present invention is an apparatus and method for producing very large area and large volume plasmas. The invention utilizes electron cylcotron resonances in conjunction with permanent magnets to produce dense, uniform plasmas for long life ion thruster applications or for plasma processing applications such as etching, deposition, ion milling and ion implantation. The large area source is at least five times larger than the 12-inch wafers being processed to date. Its rectangular shape makes it easier to accommodate to materials processing than sources that are circular in shape. The source itself represents the largest ECR ion source built to date. It is electrodeless and does not utilize electromagnets to generate the ECR magnetic circuit, nor does it make use of windows.

The invention discloses a flexible insulator for a grid component of an ion thruster. The insulator comprises a first screw rod, a second screw rod, a spring and insulating ceramic, preferably the insulator further comprises a shielding hood, wherein the shielding hood is of a cavity structure, an opening is formed in one side of the cavity structure, the insulating ceramic is placed in the cavity of the shielding hood, and threaded holes are formed in both end surfaces of the insulating ceramic; the both ends of the spring are respectively connected with the first screw rod and the second screw rod in a welding manner, and the second screw rod penetrates through the bottom surface of the shielding hood and is connected with the threaded hole of one end surface of the insulating ceramic by screw threads, so that the shielding hood and the insulating ceramic are connected together to form the flexible insulator; the threaded hole of the other end surface of the insulating ceramic is used for being in threaded connection with a support ring of a piece of grid through the screw rods, and the first screw rod is used for being in screw joint with the other piece of grid by screw threads. Through the utilization of the grid component of the ion thruster, which is assembled by the flexible insulator, the changes of the spacing between the two grid pieces of the grid component can be effectively reduced, the spacing and the insulation of the two grid pieces are ensured, and the reliability of the ion thruster and the stability of the thrust are ensured.

The invention discloses a service life ground test method for an ion thruster. By using the ground test method, a ground service life test for the iron thruster can be performed for a long time. According to the ground test method, service life ground test equipment is built first of all, the performance of the ion thruster after working for a long time is tested on the basis of the simulation space condition of the test equipment, and accordingly actual results of the service life of the ion thruster can be obtained, and real and reliable test data are provided for evaluation of the service life and the reliability of the thruster. Main performance parameters of the ion thruster in the whole service life can be obtained through substantive tests in the test process, and the change rule of the performance parameters of the thruster along with accumulative working time is further obtained; the basis for long-service-life working and application of the ion thruster is provided, and essential data can be obtained for more effective acceleration tests of the iron thruster in future. The test method is effective and reliable, and the economic cost and the time cost of the service life test are greatly lowered.

A system and a method for an ambient atmosphere ion thruster having a pair of permeable electrical members for accelerating ambient atmosphere ions which enter the thruster due to a craft's relative velocity. Neutral ambient atmosphere molecules in the intake mass flux may also be ionized and subsequently accelerated by the pair of permeable electrical members. Acceleration of any entering mass comprises an exhaust mass flux which then produces a net thrust. Such net thrust is then vectored by configuring and orienting the thrusters for imparting axial and lateral accelerations as well as pitch, yaw, and roll controls. The present invention is operational in proximity to any celestial body having a sensible atmosphere during a portion of the free trajectory or during orbiting at altitude. Useful net thrust is thereby produced without need for carrying reaction mass onboard a spacecraft.

The invention discloses an underwater mechanical bionic flapping wing thruster which comprises the following five parts: a horizontal movement mechanism, a vertical movement mechanism, a tilting movement mechanism, a bionic flapping wing and a fixed base plate. A mechanical motion separating method is used in the thruster, for imitating the flapping wing three-dimensional motion by using a simple mechanical transmission mechanism from the perspective of Kinematics. The whole motion process of the mechanism is as follows: an offset crank block mechanism is driven to move by a motor so as to drive a horizontal slide block to move left and right in a horizontal slide block motion guide rail; meanwhile, a cam push rod mechanism is driven by the motor so as to drive a vertical slide block to move vertically relative to the horizontal slide block; the two motions are superposed for realizing the fitting of the tail motion locus of the flapping wing; and at the same time, the tilting moving mechanism is driven to move so as to realize the tilting motion of the flapping wing. The underwater mechanical bionic flapping wing thruster has the characteristics of simple and reliable structure, capability of generating a stable periodic propelling force, high energy utilization rate and wide market development space.

A helicopter includes a rotor with short blades and tip thrusters. The blades extend from an annular fuel tank that rotates with the blades. Centrifugal force propels fuel from the interior compartment of the annular tank, through conduits extending through the blades from root to tip, to the tip thrusters. Valves (e.g., solenoid valves and / or a flyweight governor) regulate fuel flow to achieve and maintain a determined steady rotational speed. A fuselage covered by a canopy is mounted atop a bearing on the annular fuel tank. The pitch of each blade may be adjusted collectively and cyclically using a swash ring and hydraulic linear actuators.

The invention discloses an ion and Hall hybrid type electric thruster. By means of the ion and Hall hybrid type electric thruster, an ion thruster and a Hall thruster can be effectively combined, the integral design is formed, and the thruster can fully inherit the advantages of the ion thruster and the advantages of the Hall thruster. The ion and Hall hybrid type electric thruster comprises an inner ion ring, an outer Hall ring, a cathode and a magnetic shielding structure, wherein a direct current ion thruster structure is adopted in the inner ion ring; a steady state plasma Hall thruster structure is adopted in the outer Hall ring; the inner ion ring and the outer Hall ring are concentrically arranged, and the outer Hall ring is embedded in the outer side of the inner ion ring; the annular magnetic shielding structure made of permanent magnet materials is arranged between the inner ion ring and the outer Hall ring. The cathode is of a hollow cathode structure and is located on the outer side of the outer Hall ring.

The invention relates to a multi-mode spiral wave ion thruster and belongs to the field of aerospace electric propulsion. A quartz discharging chamber is placed in a Faraday barrel; an anode is fixedin an opening end of the quartz discharging chamber; an opening end of the quartz discharging chamber is in contact connection with a screen grid; the screen grid, an acceleration grid and a deceleration grid keep a distance; an antenna is mounted on an outer barrel wall of the quartz discharging chamber; a plurality of spiral coils are mounted on an outer wall of the Faraday barrel; a cathode neutralizer is placed at the opening end of the quartz discharging chamber and is used for neutralizing an electrostatic grid electrode to accelerate led-out ion beam flow; the cathode neutralizer and the anode generate an axial electric field, a route of electrons can be effectively increased and the ionization rate is improved. The thruster provided by the invention is provided with the neutralizerand the neutralizer can be used for neutralizing the ion beam flow led out from the electrostatic grid electrode; cathode and anode structures are arranged and the axial electric field can be provided, so that the ionization rate of plasmas is improved and the production cost of ions can be calculated; the designed spiral wave ion thruster has the advantages that an additional magnetic field hasan adjustable bit type and multi-mode working of the thruster can be realized.

A thrust producing system includes an RF ion thruster with a discharge chamber having a gas inlet and an outlet, and a coil about the discharge chamber. The system further includes an RF cathode proximate the discharge chamber outlet of the RF ion thruster for ion beam neutralization. The RF cathode includes a discharge chamber having a gas inlet and an outlet and a coil about the discharge chamber. A tank for containing iodine in solid form and a heater associated with said tank to produce iodine vapor. A feed subsystem fluidly couples the tank with the RF ion thruster discharge chamber gas inlet and with the RF cathode discharge chamber gas inlet.

The present invention discloses an ion thruster beam test system based on a Faraday probe array. According to the present invention, the Faraday probes are arranged in a surface array manner, all probe test signals are acquired synchronously, the data information amount is large, and the whole test and the data processing can be finished usually in only several minutes. A shielding outer sleeve is arranged at the outer side of an ion current catch tray of each Faraday probe and is used to shield the stray ions flying to the side surface of the ion current catch tray, thereby guaranteeing that the ions are only collected on the right side of the ion current catch tray. The shielding outer sleeve is directly and electrically connected with a disc-type metal support by the metal contact, so that a bias voltage can be supplied to the disc-type metal support by just one conducting wire, and the biggest advantages of the mode are that: the number of the probe leads is reduced substantially, the structure is simplified, and the reliability is improved.

The invention relates to a loading test apparatus for simulating a ship thruster. The loading test apparatus comprises a propeller simulation block, an axial static thrust loading device, a dynamic force loading device, torque loading devices and a three-component dynamic force measuring device. The propeller simulation block is installed at the tail end of a propeller shaft screw shaft; the axial static thrust loading device is connected with the propeller shaft to realize axial force transmission; the dynamic force loading device can load dynamic forces along axial and radio directions; the torque loading devices are arranged at the two sides of the propeller shaft symmetrically and transmission between the torque loading devices and the propeller shaft is realized by a narrow V-band mechanism; and the three-component dynamic force measuring device can measure practical dynamic force component force values transmitted to X, Y, and X directions of the propeller shaft. With the test apparatus provided by the invention, various loads in mutual interaction between the propeller and the propeller shaft during the ship operating process can be simulated comprehensively; and loading of an axial static thrust, a dynamic force, and a torque can be realized.

An ion thrustor minimum subsample reliability assessment method comprises the following steps that first, according to the failure and structure features of an ion thrustor, a service life distribution model is selected; second, according to the ground service life verification test results of the ion thrustor, a reliability analyzing method is selected, if a subsample is non-failure data, an ion thrustor non-failure data reliability analyzing method is selected, and if the subsample is minimum failure data, an ion thrustor minimum failure data reliability analyzing method is selected; third, the service life dispersibility of the ion thrustor non-failure data is determined; fourth, the parameter values of the ion thrustor minimum failure data service life distribution model are computed; and fifth, the reliability lower confidence limit of the given service life of the ion thrustor and the service life lower confidence limit of the given reliability are computed. The reliability assessment method for ion thrustor non-failure data and minimum failure data is established, reliability assessment accuracy is improved, and a foundation is provided for reliability assessment of the ion thrustor in the future.

The invention discloses a vacuum pumping system of test equipment of ion thrusters. Through the adoption of the vacuum pumping system disclosed by the invention, a vacuum environment can be provided for the ground test equipment of the ion thrusters, and the special requirements that the ground test equipment of the ion thrusters is oil-free vacuum, is large in pumping capability, prevents sputtering, has a plasma-protection function, and is high in reliability are met. The vacuum pumping system disclosed by the invention comprises a vacuum pumping device of a main chamber vacuum container, and a vacuum pumping device of an accessory chamber vacuum container, wherein the vacuum pumping device of the main chamber vacuum container comprises a coarse pumping assembly of the main chamber vacuum container, a transition pumping assembly of the main chamber vacuum container, and a high vacuum sucking assembly of the main chamber vacuum container. Firstly, the main chamber vacuum container is pumped to 10Pa by using the coarse pumping assembly of the main chamber vacuum container; then, the main chamber vacuum container is pumped to 10-2Pa by using the transition pumping assembly of the main chamber vacuum container; finally, when the main chamber vacuum container is pumped to below 1.0*10-4Pa by using the high vacuum sucking assembly of the main chamber vacuum container, a test is started.

The invention discloses a time division multiplexing manner based small satellite orbit control method, and relates to the mall satellite orbit control field, which can solve the problem that small satellite orbits are high in cost and low in efficiency. The method includes: obliquely arranging four thrusters on a lower surface of a small satellite in a symmetrical manner, wherein the first thruster and the second thruster are symmetrical with respect to an X axis, the third thruster and the fourth thruster are symmetrical with respect to the X axis, the first thruster and the fourth thrusterare symmetrical with respect to a Y axis, and the second thruster and the third thruster are symmetrical with respect to the Y axis; keeping the inclined angles of the four thrusters in a range of -55degrees-+55 degrees, wherein the inclined angles and directions of the first thruster and the second thruster are identical, and the inclined angles and directions of the third thruster and the fourth thruster are identical; using a retroaction momentum wheel to control an attitude of the small satellite, and using attitude control pulse to unload the angular momentum of the satellite; and correcting a control coefficient by using an on-orbit calibration manner. The time division multiplexing manner based small satellite orbit control method is low in cost, can save resources, and is simple and reliable.

The invention provides an integrated electric propulsion satellite platform position holding electric thruster redundant configuration method. Firstly, a satellite ground board rectangular coordinate system is built; secondly, the side, close to a south board, of a satellite ground board is provided with three electric thrusters, and the side, close to a north board, of the satellite ground board is provided with three electric thrusters; thirdly, work areas of the electric thrusters are generated respectively under the condition that all the electric thrusters work or some electric thrusters break down, ignition is conducted according to the work areas of the electric thrusters, and the electric thruster redundant configuration method is implemented. By arranging the six electric thrusters to keep the south, north, east and west positions of the satellite on-orbit period and complete momentum unloading together, requirements of an integrated electric propulsion satellite for thrust strength and the service life of the thrusters can be effectively reduced, adaptability of the integrated electric propulsion satellite to the electric thrusters is improved, the problems that a single principal part thruster makes an application strategy complex if losing efficacy and propellant consumption demand is increased are solved, and the method has the advantage of being good in adaptability to failure of the single thruster.

The invention discloses an ion thruster operation life evaluation method based on grid corrosion. The method comprises the following steps of substituting a test measurement result and a numerical simulation calculation result into a semi-empirical analysis model; calculating to obtain the operation life of the ion thruster; evaluating the operation life of the ion thruster by using limited test data. Therefore, the most critical factors of the influence of the service life of the ion thruster can be quickly determined, the links of machining, manufacturing, assembling, testing, testing and the like of the thruster after failure are effectively reduced, the production cost of products is reduced, and the application range and the application value of the ion thruster can be predicted in advance.

A voltage tracking system for an ion thruster includes a discharge chamber, a screen grid, an accelerator grid, and an accelerator grid voltage controller. The discharge chamber contains plasma at a given potential. The screen grid is adjacent to the discharge chamber and is voltage biased relative to the plasma to form a plasma sheath that repels electrons and attracts ions from the discharge chamber plasma. The accelerator grid is adjacent to the screen grid and has a voltage for accelerating the ions to create thrust and prevent any electrons from backstreaming into the ion thruster from the beam plasma. The accelerator grid voltage controller supplies voltage to the accelerator grid. The accelerator grid voltage controller adjusts the magnitude of the accelerator grid voltage to minimize the amount of voltage required to prevent electron backstreaming into the ion thruster.