93 results about "Propellant tank" patented technology

The invention relates to a welding seam reinforcing structure for a stainless steel storage tank, and belongs to the technical field of aerospace carrier propellant storage tank structures. A first box bottom, a first short shell, a cylinder section, a second short shell and a second box bottom; the first box bottom and the second box bottom are each of a hemispherical shell structure. The first short shell, the second short shell and the barrel section are all of barrel section structures; the first short shell, the barrel section and the second short shell are horizontally and coaxially butted in sequence; the first box bottom is coaxially mounted at the axial outer end of the first short shell; the second box bottom is coaxially mounted at the axial outer end of the second short shell; the opening end of the first box bottom extends into the first short shell, and the outer wall of the first box bottom is in butt joint with the inner wall of the first short shell. The opening end of the second box bottom extends into the second short shell, and the outer wall of the second box bottom is in butt joint with the inner wall of the second short shell. The problem that the welding performance of reinforcing materials cannot be weakened through partition thickness design is solved, and the purposes of high performance, low cost manufacturing and the like of the stainless steel storage box are achieved.

A flame effect system for an unmanned aerial vehicle (UAV) includes a hopper configured to store a powdered fuel, a propellant tank configured to store a propellant, and a nozzle configured to expel the powdered fuel into an atmosphere. Furthermore, the flame effect system includes a fluid path extending from the propellant tank to the nozzle. The hopper is fluidly coupled to the fluid path at an intersection between the propellant tank and the nozzle, the hopper is configured to enable the powdered fuel to flow into the fluid path, and the propellant tank is configured to expel the propellant through the fluid path to fluidize the powdered fuel within the fluid path and to drive the fluidized powdered fuel through the nozzle.

The invention relates to a multi-branch diffusion box bottom force transmission structure, and belongs to the technical field of aerospace vehicle propellant storage box structures, the multi-branch diffusion box bottom force transmission structure comprises a thrust ring base plate, a thrust ring outer ring, a thrust ring butt joint plate and a thrust ring inner ring, the thrust ring base plate is of a flat plate structure with a hole in the center, the thrust ring outer ring is of a cylindrical structure, and the inner diameters of the upper end face and the lower end face are different; the thrust ring butt joint plate is arranged on the upper end face of the thrust ring outer ring and used for being connected with an engine rack; the thrust ring inner ring is of a cylindrical structure with the upper end face and the lower end face being the same in inner diameter and is arranged in a cavity of the thrust ring outer ring, one end of the thrust ring inner ring is connected with the thrust ring butt-joint plate, and the other end of the thrust ring inner ring is connected with the thrust ring base plate which is connected with the storage tank bottom. The thrust ring structure is additionally arranged on the outer side of the bottom of the storage box, the thrust ring is connected with the bottom of the storage box through multiple branches, thrust transmitted by an engine rack is diffused to a main structure of the bottom of the storage box through the thrust ring, and the effects of saving an engine force transmission section and improving the engine force transmission efficiency are achieved.

The utility model discloses an ignition agent filling or discharging tool for an ignition agent storage box, and belongs to the technical field of ignition agent filling or discharging of ignition agent storage boxes. Comprising two fixing nuts, two threaded sleeves, two first check rings, two first sealing rings, two second check rings, two second sealing rings, two rotating shafts, two shells, two outer sleeve nuts and a storage box, the fixing nuts are in threaded connection with one ends of the rotating shafts, one ends of the fixing nuts are in lap joint with the threaded sleeves, the rotating shafts are sleeved with the threaded sleeves, and the storage box is connected with the outer sleeve nuts. The screw sleeve is in threaded connection outside the shell; according to the utility model, the ignition agent storage tank can be filled or discharged in an aerobic environment, the problem of high skill requirements on operators is solved, and meanwhile, compared with the traditional technology, the operation efficiency is greatly improved, the safety and reliability of the operation process are ensured, the operation is simple, and the practicability is high. And safety accidents caused by combustion of the ignition agent in contact with the air are avoided.

The present application relates to a kind of weld reinforcement structure for stainless steel tank, belong to aerospace vehicle propellant tank structure technical field;First tank bottom, first short shell, barrel section, second short shell and second tank bottom;Wherein, first tank bottom and second tank bottom are hemispherical shell structure;First short shell, second short shell, barrel section are barrel section structure;First short shell, barrel section and second short shell are in turn horizontal coaxial butt joint;First tank bottom is coaxially installed at the axial outer end of first short shell;Second tank bottom is coaxially installed at the axial outer end of second short shell;The open end of first tank bottom is inserted into first short shell, and the outer wall of first tank bottom is butt joint with the inner wall of first short shell;The open end of second tank bottom is inserted into second short shell, and the outer wall of second tank bottom is butt joint with the inner wall of second short shell;The present application solves the problem that the welding performance of reinforcing material cannot be reinforced by partition thickness design, realizes the purposes such as high performance, low cost manufacturing of stainless steel tank.

The application discloses a capillary type ion liquid electric spray thruster with electrically controlled thrust, and belongs to the technical field of micro-nano satellite micro propulsion. The application comprises a propellant tank, an upper shell, a constant force spring, a limiting plate, a capillary and an extraction electrode. The inner and outer walls of the capillary are covered with a dielectric layer with a micron level thickness, and the inner wall is covered with a hydrophobic layer with a nanometer level thickness. The extraction electrode is fixed through an insulating screw; and the upper shell is fixed with the propellant tank. The application is based on the principle of electric wetting, uses an electric field to control the active supply of ion liquid and simultaneously adjusts the thrust size of the thruster. The application can not only save complicated fluid control components such as high-precision pumps, valves and pipelines, but also can couple the supply process and the emission process of the ion liquid, simplify the operation process of the thruster system and facilitate the miniaturization of the thruster system. The application is suitable for resistance compensation, position keeping and attitude control of low-power micro-nano satellites.

This application provides a horizontal assembly and welding method for ultra-long cylindrical sections with circumferential seams, comprising: hoisting the long cylindrical section into a horizontal fixture and controlling the horizontality of the long cylindrical section and the coaxiality of the cylindrical section axis with the axis of the support mechanism; adjusting the roundness of the assembled cylindrical section by controlling the support tension of the inner support fixture and using an auxiliary local external pressure method; eliminating local misalignment after cylindrical section assembly by local positioning welding and local correction; and eliminating local gaps after cylindrical section assembly by in-situ milling of the cylindrical section end face and full positioning seam sealing welding, wherein the milling includes rough milling and fine milling. This application enables the horizontal assembly and welding of ultra-long cylindrical section propellant tanks for next-generation medium-sized launch vehicles with circumferential seams, solving the problem of deformation during horizontal assembly of thin-walled ultra-long weak rigid cylindrical section products, reducing product production costs, and improving product assembly and welding efficiency, providing effective technical support for a high-efficiency, mass production model for rocket propellant tank products.

The invention provides an engine support and a carrier rocket. The engine support is applied to a carrier rocket and comprises a support body and a mounting frame, the support body is in a table shape with a closed first end and an open second end, and the diameter of the first end of the support body is smaller than that of the second end; the first end of the support body is provided with an engine installation face used for being connected with a two-stage rocket engine, the second end of the support body is used for being connected with the bottom of a propellant storage box, and the outer wall face of the support body is used for installing engine equipment. The mounting frame is arranged on the outer wall face of the support body and used for being detachably connected with a servo swing mechanism of the two-stage rocket engine. The support body is arranged to be table-shaped and arranged at the bottom of the propellant storage box, the size and weight of the engine support can be reduced, and the weight of a carrier rocket can be reduced.

The invention concerns a spacecraft propellant tank designed for storing and supplying fuel to the propulsion system, as well as its application method, which is related to space technology and particularly to the design of tanks for storing liquid propellants within rocket engines. The tank in question comprises a body featuring at least one convex edge that creates a corner capillary and an intake port positioned on one of these convex edges. Additionally, the tank's interior surface is treated to be hydrophilic. The objective of this invention is to develop a spacecraft propellant tank characterized by its lightweight and straightforward design. It facilitates the creation of various propellant tank shapes while minimizing the quantity of internal components that reduce useful volume. It also ensures consistent and reliable functionality of the propellant system with continuous supply of propellant to the intake port, even in the absence of gravity and during multiple activations of the propulsion system. The claimed invention achieves its objective by implementing technical results that include enhancing the efficiency of propellant supply within the tank to the intake port, improving reliability, augmenting the useful internal volume, and offering a variety of tank shapes. The stated objectives are met in particular through the hydrophilic properties of the propellant tank's inner surface, the design of the propellant tank shape featuring at least one convex edge that forms a corner capillary, and positioning of the intake port on one of the convex edges that forms the corner capillary. The application method for the spacecraft propellant tank involves initially filling the tank with liquid propellant and pressurizing gas through the intake port prior to spacecraft launch. Subsequently, once the spacecraft is in orbit, propellant is conveyed from the tank to the propulsion system through the intake port. The movement of liquid propellant towards the intake port is facilitated by the pressure of the pressurizing gas and a localized pressure reduction generated during propellant discharge. This reduction occurs in the corner capillary near the intake port relative to the pressure of the propellant in other sections of at least one corner capillary, thus establishing a flow of propellant through the corner capillaries towards the intake port area.

Aerospace systems, and systems and methods for assembling an aerospace system or another system or structure, are disclosed. A representative system can include a propellant tank with a dome-shaped head; attachment features on the dome-shaped head; and a thrust structure with legs attached to the attachment features. A representative method can include forming shims to be positioned between the legs and the attachment features. The shims can be formed based on digital characterizations of mounting faces on the attachment features within a reference coordinate system. The method can include positioning a drill jig within the reference coordinate system and using the drill jig to make holes in the shims and / or the attachment features for fastening the legs to the attachment features. A drill jig can include plates that are movable relative to each other. At least one of the plates can include a mounting hole pattern.

A rocket propellant tank and layout (15) is disclosed. The tank (15) includes a housing operationally coupled to a rocket engine (20). The housing comprises a plurality of compartments (17a, 17b, 17c, 17d, 17e, 17f) arranged longitudinally along the housing and is adapted to store a propellant. A plurality of separator walls (24, 41, 42, 32, 33, 34) are disposed adjacent to each of the plurality of compartments (17a, 17b, 17c, 17d, 17e, 17f). A set of control valves (45) are coupled to each of the plurality of compartments (17a, 17b, 17c, 17d, 17e, 17f). The set of control valves (45) are adapted to facilitate a controlled transfer of the propellant from each of the plurality of compartments (17a, 17b, 17c, 17d, 17e, 17f) to the rocket engine (20).

The present invention relates to a multifunctional aerospace heat pipe loop system, which belongs to the technical fields of aerospace thermal control, aerospace power, and heat exchanger; the system comprises a propulsion system, which comprises a propellant storage tank and a thruster, and propellant is transported between the propellant storage tank and the thruster through a propellant pipeline assembly; the propellant pipeline assembly comprises a propulsion bypass pipeline, and the output end of the propulsion bypass pipeline is connected to the input port of the thruster; the system also comprises a heat dissipation condensation pipeline, a heat absorption evaporation pipeline and a three-way valve; the input end of the heat absorption evaporation pipeline is connected to the output end of the propellant storage tank, and the output end is connected to the three-way valve; the heat absorption evaporation pipeline passes through the heat absorption evaporation area; the input end of the heat dissipation condensation pipeline is connected to the three-way valve, and the output end is connected to the heat absorption evaporation pipeline; the heat dissipation condensation pipeline passes through the heat dissipation condensation area; the three-way valve is connected to the input end of the propulsion bypass pipeline; the propellant can be used as a heat flow carrier medium, and the active heating of the propellant pipeline can be reduced, thereby enhancing the gasification efficiency of the propellant.

The application discloses a low-temperature liquid rocket engine immersion precooling system and method, and belongs to the technical field of liquid rocket engine precooling, which comprises a low-temperature propellant tank, a conveying pipeline, a low-temperature pump, a post-pump pipeline and low-temperature valves; the application only utilizes necessary heat exchange components of the engine to realize immersion precooling, and does not need to set a discharge valve, a discharge pipeline, a reflux pipeline and a reflux valve, so that the system structure is greatly simplified; the precooling process does not need propellant discharge, effectively reduces the propellant consumption, and improves the rocket carrying capacity; in combination with an overall layout of an inclined downward direction, a guide inner wall and an inclined upward design of the post-pump pipeline, a continuous exhaust passage is constructed, gas is ensured to smoothly float up, and the precooling efficiency is greatly improved; by limiting the relative relationship between the inclination degrees of the guide inner wall and the post-pump pipeline, the gas floating rate can be flexibly optimized according to different working conditions, the best balance between the exhaust efficiency and the flow stability is realized, and the system working condition adaptability and the precooling effect controllability are remarkably improved.

A number of techniques, structures, and materials for thermally insulating a cryogenic propellant tank barrel section are presented. Thermal insulation for a tank, such as tanks for use in space flight missions, may be a dual wall vacuum jacket system where one wall is the propellant tank wall and the other wall is a metallic foil that is welded to portions of the propellant tank wall. The metallic foil may act as a vapor barrier and a vacuum seal. A structural and breathable insulation that is durable and functional in a range from cold to relatively high temperatures may be used to maintain a standoff distance between the two walls. The volume between the propellant tank wall and the metallic foil may be purged with a gas that has a liquefaction temperature greater than that of the cryogenic propellant that will be in the propellant tank during operations.

This invention provides a multi-channel refueling system for measuring the volume of a launch vehicle's propellant tank, comprising: a single-channel refueling system and a newly developed data acquisition and control system; the single-channel refueling system includes a test station and a flow meter; the newly developed data acquisition and control system includes a pulse distributor, a PLC module, an industrial computer, and control components; the pulse distributor is connected to the test station, the flow meter, and the PLC module respectively; the PLC module is connected to the industrial computer and the control components respectively. This invention possesses multi-channel refueling capability and multi-channel test data processing capability, achieving a several-fold increase in the efficiency of launch vehicle propellant tank volume measurement.

This invention relates to the field of aerospace propulsion technology, specifically disclosing a high-frequency induction heating space propulsion system and method. The system includes a Laval nozzle, a high-frequency induction heating chamber, a flange joint, and a working propellant delivery pipeline. The high-frequency induction heating chamber includes an induction coil, a heating chamber shell, and a metal workpiece to be heated. The induction coil is wound around the outside of the heating chamber shell. The induction coil generates a high-frequency alternating magnetic field powered by a high-frequency AC power supply. This magnetic field induces eddy currents in the metal workpiece, thereby heating the heating chamber shell. The working propellant delivery pipeline injects propellant from the propellant tank into the high-frequency induction heating chamber for heating. The resulting high-temperature, high-pressure gas is ejected from the Laval nozzle, generating thrust. Utilizing the principle of electromagnetic induction, eddy currents are generated inside the thruster shell, directly heating the propellant, offering advantages such as high heating efficiency and fast response speed.

This utility model relates to a satellite electric propulsion system, addressing the problems of high cost, poor control precision, and low efficiency and reliability of existing chemical thrusters for micro and nanosatellites, as well as the issues of low efficiency and low reliability of resistance-heated thrusters. The system includes a propellant tank, a power supply and control module, a dual-path pressure regulation module, and a dual-thrust module. The propellant tank uses a full TC4 titanium alloy spherical structure to store liquid ammonia. The power supply and control module controls the high-voltage arc heating of the working fluid via a CAN bus. The dual-path pressure regulation module achieves redundant pressure reduction through solenoid valves, preheaters, and pressure reducing valves. The dual thrusters are arranged vertically at their centers of mass, providing mutual cold backup. This design reduces size and weight through modular integration, improves electrothermal conversion efficiency, orbit climbing mission success rate, and working fluid utilization efficiency, effectively solving the technical bottlenecks of existing propulsion systems and is suitable for high-precision orbit control of micro and nanosatellites.

Thermal protection systems for aerospace vehicles, such as hypersonic vehicles or space vehicles, are disclosed herein. In some embodiments, the disclosed thermal protection systems can be used to protect propellant tanks (e.g., composite propellant tanks) and vehicle primary structures (e.g., composite primary structures) from exceeding upper design temperature limits during vehicle ascent, reentry into Earth's atmosphere, and / or high velocity flight through an atmosphere. In some embodiments, a thermal protection system can include a framework of elongated composite members forming a plurality of cells. One or more layers of insulation material such as aerogel can be positioned in each of the cells. The cells can be covered by an external skin of the vehicle. In some embodiments, the framework is attached to a propellant tank and the insulation material is in contact with the external surfaced of the propellant tank.

A flame effect system for an unmanned aerial vehicle (UAV) includes a hopper configured to store a powdered fuel, a propellant tank configured to store a propellant, and a nozzle configured to expel the powdered fuel into an atmosphere. Furthermore, the flame effect system includes a fluid path extending from the propellant tank to the nozzle. The hopper is fluidly coupled to the fluid path at an intersection between the propellant tank and the nozzle, the hopper is configured to enable the powdered fuel to flow into the fluid path, and the propellant tank is configured to expel the propellant through the fluid path to fluidize the powdered fuel within the fluid path and to drive the fluidized powdered fuel through the nozzle.

The invention relates to a cryogenic propellant tank for a spacecraft engine, the tank comprising a wall that defines a storage volume for the cryogenic propellant, the wall being made of a composite material and comprising a fibrous reinforcement in the form of a fibrous stack (110) which is densified by an organic matrix, the fibrous stack comprising at least in succession and in this order in the direction of the thickness (E) of the wall: (a) a first ply (113) having unidirectional fibres oriented at an angle of –α; (b) a second ply (115) having unidirectional fibres oriented at an angle of -β°; (c) a third ply (117) having unidirectional fibres oriented at an angle of +α; and (d) a fourth ply (119) having unidirectional fibres oriented at an angle of +β°, with α ranging from 5° to 20° and β ranging from 40° to 50°, wherein each angle is taken with respect to a longitudinal axis of the tank.

Dual mode engine for propelling spacecraft, including combustion chamber having flange end, open nozzle end, and enclosed chamber portion extending therebetween, propellant tank in fluidic communication with combustion chamber, electronic controller, power source operationally connected to electronic controller, and fluid flow motivator operationally connected to electronic controller and connected in fluidic communication with propellant tank. Engine has chemical propulsion portion with propellant inlet port operationally connected to combustion chamber and disposed adjacent flange end, ignition trigger electrode positioned in combustion chamber adjacent propellant inlet port and operationally connected to electronic controller and operationally connected to power source propellant inlet port fluidically connected to tank electric propulsion portion with two electrodes ionizing propellant positioned in combustion chamber adjacent nozzle end, plurality of attitude control thrusters operationally connected to electronic controller and in fluidic communication with propellant tank, and plurality of valves, each fluidically connected between attitude control thruster and propellant tank.

A debris removal satellite includes a capture device, a thruster of a chemical propulsion method, and a propellant tank to store chemical fuel. A solar array wing is operable in an orbit at an orbital altitude higher than a congested orbit region congested with satellites forming a satellite constellation. The debris removal satellite is built in advance for future use as a satellite to be launched, and when a debris intrusion alarm to give a warning about intrusion of debris into the congested orbit region is issued, propellant is loaded into the propellant tank and the debris removal satellite is launched by a rocket built in advance for future use as a launch rocket. The debris removal satellite captures capture-target debris at an orbital altitude higher than the congested orbit region, and operates a propulsion device with the capture-target debris being captured.

A debris removal satellite includes a capture device, a thruster of a chemical propulsion method, and a propellant tank to store chemical fuel. A solar array wing is operable in an orbit at an orbital altitude higher than a congested orbit region congested with satellites forming a satellite constellation. The debris removal satellite is built in advance for future use as a satellite to be launched, and when a debris intrusion alarm to give a warning about intrusion of debris into the congested orbit region is issued, propellant is loaded into the propellant tank and the debris removal satellite is launched by a rocket built in advance for future use as a launch rocket. The debris removal satellite captures capture-target debris at an orbital altitude higher than the congested orbit region, and operates a propulsion device with the capture-target debris being captured.

This invention provides a rocket propellant tank bracket positioning system and method based on database and machine vision, including the following steps: S1: establishing a database of external brackets for various propellant tank models; S2: retrieving propellant tank data from the external bracket database according to the propellant tank model; S3: the automated positioning system for the external brackets corrects the downward displacement of each position of the propellant tank model under gravity in the external bracket model database based on measurement results; S4: rotating the reference for the external bracket positioning and assembly of the propellant tank to a suitable position; S5: generating assembly and marking schemes for each external bracket based on machine vision results; S6: operators welding each bracket sequentially at the bracket welding station according to the marked positions determined by each reference; and S7: grinding away the weld seams. This application features rapid positioning technology that enables one-time accurate marking of batch propellant tank brackets, reducing repeated adjustments and improving production line consistency.

The invention is suitable for the field of spaceflight carrier rockets, and provides a vertical recovery carrier rocket propellant storage box which comprises an upper cabin, a lower cabin, a partition plate assembly, a flow accumulator, an anti-shaking plate and an anti-vortex and anti-collapse device. The storage box is divided into an upper cabin and a lower cabin through the partition plate assembly. The lower cabin is provided with a flow accumulator, and an outlet of the flow accumulator is provided with an anti-vortex and anti-collapse device; and the anti-shaking plates are unevenly distributed on the upper cabin and the lower cabin along the axis of the storage box. According to the invention, the management problem of vertical recovery carrier rocket propellants is solved, and the ignition condition of multiple starting of an engine can be met; a special bottom sinking system is omitted, and compared with an existing scheme adopting a bottom sinking engine, the system complexity is reduced, the mass of a rocket body is reduced, and the reliability of first-sublevel recycling work is improved.

A thruster includes a current loop, a magnetic circuit within the current loop, a channel, a propellant reservoir and an ejection outlet. The channel forms an ejection path from an output of the propellant reservoir to the ejection outlet. The current loop is driven by a tuned current density based on at least a selected propellant. The selected propellant is ejected down the channel due to an applied electric field.

The present invention discloses a POGO vibration suppression system and a POGO vibration suppression method, belonging to the field of the liquid launch vehicle power system, which includes an accumulator and an injection gas pipeline; the accumulator is a sealed pressure vessel with an accommodation cavity inside; the propellant tank delivery pipe is communicated with the accommodation cavity; the accumulator is configured to rely on the propellant delivered to the accommodation cavity by the propellant tank delivery pipe, and an air cushion for suppressing POGO vibration is formed in the upper part of the accommodation cavity; the inlet end of the injection gas pipeline is connected to the propellant pressurization pipeline, and the outlet end of the injection gas pipeline is communicated with the accommodation cavity above the accumulator; the injection gas pipeline is configured to realize the volume control of the air cushion by filling the accommodation cavity with the volume of the vaporized propellant. The present invention does not need to carry an additional helium gas source, nor does it need to be inflated before launch, which greatly reduces the complexity of the on-board power system and improves the reliability of the launch.

A method and apparatus for measuring a level for calculating a flow of a propellant tank are provided. The method of measuring a level for calculating a flow of a propellant tank includes, in response to an instruction to measure a level in a propellant tank storing a liquid fuel, generating propagated power for level measurement at a specified generation intensity inside the propellant tank through a transmission antenna, receiving the propagated power for level measurement passing through the liquid fuel through a reception antenna, calculating a degree of attenuation between the generation intensity and a reception intensity of the propagated power for level measurement, and measuring a height level of the liquid fuel by considering the degree of attenuation.

The present invention belongs to the field of detonation engines, and specifically relates to a metal-based gel propellant detonation engine, comprising an air inlet, a rotating detonation combustion chamber, and a propellant supply system; the propellant supply system comprises a gel propellant tank disposed within a central cylinder, a gel propulsion structure connected to the gel propellant tank, and an atomizing nozzle, the atomizing nozzle being mounted on the propellant injection structure; the propellant supply system also comprises an atomizing gas tank disposed within the combustion chamber cylinder, the atomizing gas tank being connected to the atomizing nozzle via an atomizing gas delivery channel. The present invention proposes the application of metal-based gel propellant to a rotating detonation engine, which has an energy density superior to that of existing gas and liquid propellant rotating detonation engines, is highly safe, and has controllable thrust output. Furthermore, the propellant supply system does not occupy any space outside the engine, achieving a highly effective gel propellant rotating detonation engine without increasing the engine size.

The application relates to the field of rocket power systems, in particular to an upper stage annular propellant tank which comprises an upper tank shell, a lower tank shell, an upper tank management device and a lower tank management device; the upper tank shell and the lower tank shell are annular tank bodies, and the upper tank shell and the lower tank shell are stacked in layers; the upper tank management device is arranged in the upper tank shell, and a plurality of upper tank collectors of the upper tank management device are net collectors, and one upper tank collector is communicated to an upper tank liquid outlet; the lower tank management device is arranged in the lower tank shell, and a plurality of lower tank collectors of the lower tank management device are net collectors, and one upper tank collector is communicated to a lower tank liquid outlet. The application can be used for normal-temperature propellant upper stages and low-temperature propellant upper stages, can improve the structural efficiency, space utilization and propellant utilization, and can realize long-term on-orbit sliding and propellant supply requirements for multiple startings of main engines.