298 results about "Turbopump" patented technology

The invention discloses a fault diagnosis method for a turbine pump bearing of a liquid rocket engine based on dual adaptive filtering and related equipment, and the method comprises the steps: processing an original vibration signal through an adaptive line spectrum intensifier, and obtaining a vibration signal after the interference of a first-order cyclostationary signal is eliminated; blind cyclic frequency detection is carried out on the vibration signals after the first-order cyclostationary signal interference is eliminated, and a cyclic frequency set is obtained; taking the vibration signal after eliminating the interference of the first-order cyclostationary signal as the input of a multi-channel adaptive FRESH filter, taking the frequency shift component used by each channel as the frequency in a cyclic frequency set, and outputting the filtered vibration signal; and analyzing the filtered vibration signal by using cyclostationary analysis to obtain a diagnosis result. According to the method, a pure second-order cyclostationary signal related to a fault can be accurately extracted from a fault liquid rocket engine turbine pump bearing vibration signal, the filtering process can adapt to the change of a system state, and a reliable pretreatment result can be provided for a bearing fault diagnosis method based on cyclostationary analysis; and finally, accurate diagnosis of the turbine pump bearing fault of the liquid-propellant rocket engine is realized.

The invention relates to a multifunctional pump rear swing liquid rocket engine single-machine rack which is characterized in that two main force bearing beams and two auxiliary force bearing beams are welded to form a lower force transmission frame; the four upper connectors and the four upper connecting beams are welded to form an upper force transmission frame. The four main force bearing rods are welded to the upper force transmission frame and the lower force transmission frame respectively. The two servo supports are arranged to serve as upper fulcrums of the servo mechanism, and the bidirectional swing requirement of the liquid rocket engine can be met. And a pair of turbine pump supporting seats are symmetrically arranged with one of the servo supports. And a pair of turbine pump auxiliary support lugs is arranged on the same side of the turbine pump support. The turbine pump support and the turbine pump auxiliary supporting lug are jointly responsible for being connected with the turbine pump, and stable fixed supporting of the turbine pump is achieved. The main force bearing beam is connected with the lower end force transmission assembly, the upper connector is connected with a rocket body or a multi-machine rack, and the thrust transmission function of the engine is achieved.

The invention provides an axial force balance flow path for a multistage hydrogen turbine pump of a liquid rocket engine. Comprising a balance piston gap, a floating ring gap, a third-stage impeller backflow hole, a shaft diameter hole, a backflow groove, an axis hole, a second-stage impeller backflow hole I, a second-stage impeller backflow hole II, a first-stage impeller backflow hole, a pump end bearing rear sealing gap, a front sealing gap, a first-stage and second-stage rear sealing gap, a third-stage rear sealing gap and an inter-stage sealing gap. According to the downstream containing cavity of the balance piston, backflow grooves are formed in the shaft and the impeller, backflow holes of the second-stage impeller and the third-stage impeller are formed in the shaft and the impeller, water is directly drained to an inlet of the second-stage impeller, the balance capacity of the balance piston is improved in a large range, and the balance piston is suitable for axial force balance of a large hydrogen turbine pump of a high-thrust afterburning circulating liquid rocket engine.

The invention discloses a liquid rocket engine turbine pump rotating speed determination method and related device, and the method comprises the steps that a vibration signal and a rotating speed signal of a turbine pump are obtained, and the rotating speed signal is obtained through calculation of an original key phase signal of the turbine pump; processing the vibration signal by combining a Vold-Kalman filter according to the rotating speed signal to obtain a first-order signal of the vibration signal; performing Hilbert demodulation on the first-order signal of the vibration signal to obtain phase information of the first-order signal of the vibration signal; and deriving the phase information of the first-order signal of the vibration signal with respect to time, and calculating the instantaneous rotating speed of the turbine pump according to the derivation result. The invention aims to realize accurate determination of the rotating speed of the turbine pump, so that the accuracy and the reliability of fault diagnosis of the turbine pump are improved.

The invention relates to the technical field of bearing tests, in particular to a loading device and a loading method for a low-temperature bearing test of a spaceflight turbine pump. Comprising a test shaft shell, a test bearing A, an axial loading gland, an axial loading piston cylinder shell, an axial loading piston, a variable-diameter piston cylinder shell, a variable-diameter piston cylinder piston, a vaporizer, a test bearing B, a liquid nitrogen pipeline, a nitrogen pipeline and a hydraulic oil pipeline. Liquid nitrogen enters a test shaft shell from a liquid nitrogen pipeline to be fully cooled and then is discharged, the liquid nitrogen is converted into nitrogen through heat exchange of a vaporizer, the nitrogen is introduced into a variable-diameter piston cylinder shell through a nitrogen pipeline to push a variable-diameter piston cylinder piston to compress hydraulic oil, and the hydraulic oil is introduced into an axial loading piston cylinder shell to generate oil pressure to push an axial loading piston to move. The axial loading piston generates thrust on the axial loading gland to promote the axial loading gland to complete thrust loading on the test bearing, the electric cylinder is adopted to directly apply push-pull force load to the variable-diameter piston rod, load fluctuation is compensated, and stability of the test load is guaranteed.

A multifunctional low-temperature high-pressure end face sealing test device for a liquid rocket engine turbine pump is characterized in that a to-be-tested product liquid sealing wheel is sleeved on a rotor assembly, and a low-temperature end face sealing assembly is sleeved on the rotor assembly according to a certain compression amount, is in contact with a moving ring and is fixed with a shell; the rotor assembly is sleeved with the process liquid sealing wheel, and the process liquid sealing wheel and the labyrinth sealing static ring fixed to the shell form an axial fit clearance. During test, the rotor assembly rotates, the liquid seal wheel product rotates to reduce the pressure before low-temperature end face sealing, and the process liquid seal wheel, a product to be tested and the end face seal form a seal cavity; the low-temperature end face sealing assembly makes contact with the movable ring to conduct movable sealing on the sealing cavity, and the working stability of low-temperature and high-pressure end face sealing is obtained by monitoring the inlet pressure and the leakage cavity flow. According to the invention, an examination test of one or two liquid seal wheels and high-pressure end face sealing can be completed at a time, and the working condition of high-pressure and low-temperature end face sealing in the turbine pump can be effectively simulated.

[Problem] To propose a vacuum pump that offers excellent back pressure performance even when the suction gas is hydrogen gas or the like. [Solution] A vacuum pump 100 comprising an exhaust part 114 that includes: a turbine pump section 115 with a plurality of stages of rotary blades 102 and fixed blades 123; and a drag pump section 116 positioned on the exhaust downstream side of the turbine pump section 115. Intake gas sucked in from an intake port 101 by the exhaust part 114 is discharged from an exhaust port 133 via a gas flow passage FP1. The vacuum pump 100 is characterized by comprising an introduction flow passage FP3 for introducing intermediate introduction gas having a higher viscosity than the intake gas. The vacuum pump 100 is also characterized in that the introduction flow passage FP3 is connected to the gas flow passage FP1 on the exhaust downstream side of the rotary blade 102 closest to the intake port 101 among the plurality of stages of rotary blades 102 in the exhaust part 114.

The invention relates to a space turbine pump bearing low-temperature medium simulation test device, and belongs to the technical field of bearing test devices.The test device divides a test shafting into a low-temperature cooling section and a normal-temperature lubricating section, through the sectional design, the test cost is effectively reduced, and the test efficiency is improved; through a preloading spring and a compensation air inlet structure on the axial loading piston, the loading accuracy of axial force is improved, through the design of an axial guiding linear bearing, the phenomena of low-temperature clamping stagnation and high-speed slipping are prevented, and through the design of an axis liquid inlet flow channel and a spiral auxiliary liquid inlet, an inner ring of a test bearing is effectively cooled, so that the service life of the test bearing is prolonged. By dividing the cooling space in the test cavity, the flowability and cooling efficiency of the cooling medium are effectively improved, and the whole test device has high reliability, stability and safety, is suitable for the simulation test of the low-temperature bearing, and can reduce the test cost and difficulty and improve the test efficiency.

The invention relates to a rocket engine accessory, in particular to a turbopump isolation cavity heat exchange coefficient testing device and a thermal state analysis method, and aims to solve the problem that the thermal state of an internal structure of an existing turbopump isolation cavity is difficult to obtain through a test method. The device comprises a fuel cavity simulation piece, an isolation cavity simulation piece and an oxidant cavity simulation piece which are connected in sequence, the fuel cavity simulation part is a first annular cavity, the oxidant cavity simulation part is a third annular cavity, the isolation cavity simulation part is a circular truncated cone of which the large end is connected with one end of a second annular cavity, the small end of the circular truncated cone is connected with one end of the first annular cavity, and the other end of the second annular cavity is connected with one end of the third annular cavity; a first annular plate is arranged at the other end of the first annular cavity, and a third annular plate is arranged at the other end of the third annular cavity; isolation circular ring plates are arranged between the first circular ring cavity and the circular truncated cone and between the second circular ring cavity and the third circular ring cavity.

A multifunctional sealing test device for low temperature is characterized in that a to-be-tested product end face sealing assembly, a to-be-tested process end face sealing assembly and a to-be-tested leather cup sealing element are symmetrically installed on a rotor assembly and are fixed on a shell assembly, and when the three sealing elements are installed at the same time, the examination test of two product sealing elements can be completed at a time; when only the product end face sealing assembly and the process end face sealing assembly are installed, the sealing capability of the process end face sealing assembly on a gas-liquid two-phase medium is checked; when only a product end face sealing assembly and a leather cup sealing element are installed, the device is used for checking the sealing capability of the leather cup sealing element to a gas-liquid two-phase medium. When only the leather cup sealing element is mounted, the sealing capability of the leather cup sealing element on a liquid nitrogen medium is realized; when only the process end face sealing assembly is installed, the sealing capacity of the process end face sealing assembly to the liquid nitrogen medium can be achieved. According to the invention, multifunctional tests of low-temperature high-speed end face sealing of the turbine pump of the liquid rocket engine, test device process end face sealing and low-temperature packing cup sealing are realized.

The utility model relates to a turbo pump and a breathing machine. The turbine pump comprises an upper shell, a lower shell and a rotor, wherein the upper shell is provided with an air inlet; the lower shell is arranged below the upper shell, an installation space is defined by the lower shell and the upper shell, the lower shell is provided with an air outlet pipe communicated with the installation space, the air outlet pipe is spirally arranged on the surface, away from the upper shell, of the lower shell, and the longitudinal section area of the air outlet pipe is gradually increased in the air outlet direction; the impeller is rotatably arranged in the mounting space; and the brushless motor is located on the side, away from the upper shell, of the lower shell, and the output end of the brushless motor penetrates through the lower shell to be connected with the impeller so as to drive the impeller to rotate. In this way, the airflow can flow in the tangential direction of the spiral inner wall of the air outlet pipe, and stable conveying of the airflow is achieved. Meanwhile, the energy loss of the airflow in the air outlet pipe can be reduced, so that the gas flow and pressure of the airflow output by the turbine pump are increased, the response time of the turbine pump is guaranteed, and the use performance of the breathing machine is guaranteed.

The integrated structure of the turbine pump, the main valve and the fuel gas generator comprises the turbine pump, the turbine pump is of a barrel-shaped structure with the two ends open, the first end of the turbine pump is a first medium inlet, and the second end of the turbine pump is connected with a fuel gas generator flange; a movable valve element is installed in an inner channel between the turbine pump and the fuel gas generator. When the valve element moves to the first position, the inner channel is closed. When the valve element moves to the second position, the inner channel is opened; the structures of the turbine pump, the main valve and the fuel gas generator are improved, the improved turbine pump and the fuel gas generator are in flange connection, the main valve is installed in an internal channel of the turbine pump and the fuel gas generator, the integrated structure of the turbine pump, the main valve and the fuel gas generator is achieved, pipeline connection is avoided, and the space size and the assembly difficulty are reduced.

The invention provides a liquid rocket engine and a liquid rocket, and relates to the technical field of rocket engine equipment. In the engine, a turbine pump comprises a fuel pump, an oxidizing agent pump and a turbine, and an outlet of the fuel pump and an outlet of the oxidizing agent pump communicate with a thrust chamber and a fuel gas generator correspondingly; the fuel gas generator is provided with a first fuel gas passage connected with the turbine and a second fuel gas passage used for being communicated with the external atmosphere, and the fuel gas adjusting valve is installed on the second fuel gas passage. The reaction type stepping motor is connected with a valve element of the fuel gas adjusting valve through the transmission mechanism, and the controller is electrically connected with the reaction type stepping motor and used for sending pulse signals to the reaction type stepping motor at set control frequency and controlling the power-on time sequence of the reaction type stepping motor so as to drive the reaction type stepping motor to rotate forwards and backwards. And the opening degree of the gas adjusting valve is adjusted. Only one gas adjusting valve is involved in the thrust adjusting process, reliability is high, a reaction type stepping motor is adopted, large output torque can be generated under small current, and needed power is small.

The invention relates to a semi-closed performance test system for a hydraulic turbine pump of a liquid rocket engine. A turbine inlet of the hydraulic turbine pump is connected with an outlet of a centrifugal pump through a turbine inlet pipeline, and an inlet of the centrifugal pump is communicated with a water pool; a pump inlet of the hydraulic turbine pump is connected with an outlet of the water container through a pump inlet pipeline, and a pump outlet of the hydraulic turbine pump is connected with a liquid inlet of the water container through a pump outlet pipeline; the gas distribution table is connected with a gas inlet of the water container through a gas distribution pipeline; a liquid discharge port of the water container is connected with a discharge pipeline; the discharge pipeline, the pump outlet pipeline, the pump inlet pipeline and the turbine inlet pipeline are each provided with a flowmeter, a stop valve and an adjusting valve. And pressure sensors are arranged at a turbine inlet, a pump inlet and a pump outlet of the hydraulic turbine pump. The device has the capabilities of conveying a high-pressure water medium of a liquid flow test system, testing the performance of a hydraulic test system, maintaining the liquid level of a discharge system to be constant and the like, so that the overall performance test requirement of the hydraulic turbine pump can be met.

The invention discloses a liquid rocket engine turbine pump combined seal test device and system, which combines high-pressure low-temperature end face sealing and normal-temperature circumferential segmented sealing together, and adds a nozzle on a gas medium supply system for throttling and pressure reduction before segmented sealing and simulation of a real airflow state. The real working conditions of low-temperature and high-pressure end face sealing and normal-temperature circumferential segmented sealing of a liquid rocket engine turbine pump in a low-temperature and normal-temperature mixed medium can be simulated, the real working conditions of different sealing pressures can be simulated, and meanwhile multiple dynamic sealing products can be completed at a time. And reliable test data support is provided for the development of the engine while the test efficiency is improved.

The invention relates to a high-speed driving control system and a double-loop redundancy control method, a speed driving unit is electrically connected with a motor unit, an Ethernet and hard wire double-loop communication interface is designed, and the speed regulation of the motor unit is realized; the motor unit outputs specified driving power according to a driving instruction of the speed driving unit; the gear speed increasing unit is directly connected with the output end of the motor, is used for transmitting and increasing the rotating speed of the motor and providing high-speed rotating driving power for the system, and has vibration speed signal and gear bearing temperature detection functions; the lubricating and cooling unit is used for lubricating and cooling a running gear bearing in the gear speed increasing unit and monitoring the lubricating oil temperature, the oil pressure, the oil way blockage, the oil liquid level and the oil pump running in real time; the control unit integrates control logics of all the units and controls the whole high-speed driving control system to operate according to the high-speed rotation test driving requirements of turbine pump components. Feature identification of early fault symptoms of the system is realized, and the problem of lagging response of a slowly-changing parameter to a high-speed running fault is solved.

The invention discloses a digital twinborn body and a digital twinborn method for a hydraulic test system of a turbine pump of a liquid rocket engine, and solves the technical problem of high production and test resource consumption caused by research and development of engine products. The basic architecture of the digital twinning body comprises a digital twinning engine, a physical entity and a digital virtual body, and dynamic unification of digital virtual body prediction performance and object actual measurement performance is realized through virtual-real interaction and model training and optimization; finally, the display effect that a digital virtual body completely replaces a liquid rocket engine turbine pump hydraulic test system entity can be achieved, and the contradiction between increasing turbine pump hydraulic test tasks and limited physical test resources is greatly relieved.

The invention provides a gas collection structure, a turbine pump and a rocket engine. The gas collecting structure comprises a gas collecting ring and a nozzle disc, the gas collecting ring comprises an annular body and a gas inlet formed in the outer wall of the annular body, and a gas collecting groove communicated with the gas inlet is formed in one side of the annular body; the nozzle disc covers a notch of the gas collecting groove, a plurality of nozzles are arranged on the nozzle disc, inlet ends of the nozzles are communicated with the gas inlet, outlet ends of the nozzles are configured to be communicated with the turbine, the nozzles comprise a group of first nozzles and a group of second nozzles, and the second nozzles are far away from the gas inlet relative to the first nozzles; the cross-sectional area of the inlet end of the first nozzle is smaller than that of the inlet end of the second nozzle, and the cross-sectional area of the outlet end of the first nozzle is smaller than that of the outlet end of the second nozzle, so that the airflow attack angles of the first nozzle and the second nozzle at the outlet ends are equal or the difference value of the airflow attack angles is smaller than a design threshold value. The circumferential uniformity of airflow sprayed out of the nozzle disc can be improved, and the turbine efficiency is improved.

The utility model relates to a rocket engine vertical transportation packing box, which comprises a buffer base, a shell outer cover, a plurality of groups of pull rod components, a support component and a clamp component, an engine is vertically placed at the upper end of the buffer base, and the pull rod components are respectively connected between the middle part of a main body of the engine and the buffer base in an obliquely upward manner. The clamping hoop assembly is arranged on the throat portion of the engine in a clamped mode and connected with a turbine pump of the engine through the connecting rod assembly extending upwards in an inclined mode, at least one supporting assembly is connected between the buffering base and the upper portion of the throat portion of the engine in an inclined mode, and the supporting assemblies are connected with the supporting assemblies in an inclined mode. At least one supporting assembly is connected between the buffering base and a turbine pump of the engine in an obliquely upward mode, the engine is covered with the shell outer cover, and the lower end of the shell outer cover is connected and assembled with the buffering base. The device has the advantages of being simple in structure, stable, reliable, small in occupied space, high in disassembly and assembly efficiency, high in universality and the like, and the requirements for vertical transportation and storage of rocket engines can be met.

The invention discloses a liquid rocket engine starting device and method. The liquid rocket engine starting device comprises a gas cylinder set, a starting system and a pressurization system. High-pressure gas used for starting the turbine and pressurizing the propellant in the fuel storage tank is stored in the gas cylinder group; the starting system comprises a starting controller and a plurality of starting branches, and the pressure and the flow entering the turbine pump are adjusted by controlling the opening and closing number and the opening and closing time of the starting branches. The pressurization system comprises a pressurization controller and a plurality of pressurization branches, and the pressure and the flow in the gas cavity of the fuel storage tank are adjusted by controlling the opening and closing number and the opening and closing time of the pressurization branches. According to the invention, multiple starting of the liquid rocket engine can be realized, the advantages of low development difficulty and low development cost of gas cylinder starting are exerted, and meanwhile, the gas source utilization rate of traditional gas cylinder starting is optimized.

The invention discloses a filtering and circulating device for excimer laser gas, and belongs to the technical field of excimer lasers. The filter comprises a hollow cylindrical outer shell, an air inlet pipeline and an air conveying pipeline are installed at the upper end of the outer shell, and a purification box body is installed in the outer shell; the turbine pump is rotationally arranged outside the rectangular cavity; the cross flow fan is rotationally arranged in the rectangular cavity; and the output end of the double-shaft output motor is respectively connected with the turbine pump and the cross flow fan. The flow of working gas in the backflow discharge cavity is increased, meanwhile, the flowing speed of the working gas is increased, the filtering efficiency of the filter on discharge dust particles is improved, meanwhile, the filter screen rotates all the time and can be self-cleaned, and the effective filtering area of the filter screen can change along with the speed of the working gas. And the filtering effect is ensured while the filtering efficiency is ensured.

The application relates to the technical field of bearing testing, in particular to a space turbine pump low-temperature bearing test loading device and a loading method, which comprises a test shaft shell, a test bearing A, an axial loading gland, an axial loading piston cylinder shell, an axial loading piston, a variable-diameter piston cylinder shell, a variable-diameter piston cylinder piston, a vaporizer, a test bearing B, a liquid nitrogen pipeline, a nitrogen pipeline and a hydraulic oil pipeline; liquid nitrogen enters the test shaft shell, is fully cooled and is discharged, is converted into nitrogen gas through heat exchange of the vaporizer, is introduced into the variable-diameter piston cylinder shell through the nitrogen pipeline, generates compression on the hydraulic oil, the hydraulic oil is introduced into the inside of the axial loading piston cylinder shell to generate oil pressure and push the axial loading piston to move, the axial loading piston generates thrust on the axial loading gland, the axial loading gland completes thrust loading on the test bearing, an electric cylinder is directly used to apply push-pull force load on the variable-diameter piston rod, load fluctuation is compensated, and the test load is stable.

The invention discloses a reciprocating displacement pump fluid end for conveying trace super media, which comprises a pump body, a motor arranged on the pump body, a worm connected to the output end of the motor, a turbine rotatably connected to the inside of the pump body, a plunger slidably connected to the pump body, an accommodating cavity, a liquid suction one-way valve and a liquid discharge one-way valve arranged on the pump body, and a lantern ring connected to the pump body, an axial flow channel is arranged in the lantern ring; the pump body is connected with a fan-shaped valve seat, a fan-shaped cavity for heat conduction liquid to flow is formed in the fan-shaped valve seat, an inlet flow channel is formed in one end of the fan-shaped valve seat, an outlet flow channel is formed in the other end of the fan-shaped valve seat, and an electric heating wire for heating the heat conduction liquid is wound around the inlet flow channel. One end of the axial flow channel communicates with the inlet flow channel, and the other end communicates with the outlet flow channel; a swing plate is rotationally connected into the fan-shaped valve seat. A cylinder body is arranged outside the fan-shaped valve seat, a piston is connected into the cylinder body in a sliding mode, and an intercepting plate is connected to the piston. The reciprocating type displacement pump fluid end for conveying the trace super-temporary medium is suitable for low-temperature occasions, the phase state of the medium is constant, and conveying is more stable.

The utility model relates to a liquid rocket engine turbopump bearing axial clearance measuring device, which comprises a cylindrical shell, a gland, a pressure loading device, a pressure loading piece and a dynamometer, a step ring groove matched with a bearing to be measured is arranged at the upper part of the inner wall of the shell, a bearing outer ring of the bearing to be measured is embedded and supported at the groove bottom of the step ring groove, and the pressure loading device is arranged on the shell. The gland is detachably installed at the upper end of the shell, the dynamometer is arranged at the bottom in the shell, the pressure loading piece penetrates through an inner ring of a bearing to be tested, the bottom of the pressure loading piece makes contact with the dynamometer, the upper end of the pressure loading piece is provided with a force loading part attached to the upper end of the inner ring of the bearing to be tested, and the pressure loading device is installed on the gland and connected with the upper end of the pressure loading piece. And the gland is used for applying downward pressure to the pressure loading piece and is provided with a vertically through measuring hole. The device has the advantages of being simple and reasonable in structural design, reliable in work, low in operation labor intensity and capable of enabling the measurement of the axial clearance of the bearing to be more accurate and convenient.

This invention discloses a design method and inducer for a liquid rocket engine turbopump. Based on the overall design requirements of the liquid rocket engine, the operating conditions of the inducer are determined. According to the design requirements and empirical formulas, the basic parameters of the inducer are initially determined. Based on the range of core parameters, different combinations of core parameters are used to establish multiple initial three-dimensional models of the inducer. Simulations are performed on each of the established initial three-dimensional models to obtain the cavitation performance of the inducer. Based on the flow analysis results, the inducer with the optimal blade profile among all parameter combinations is obtained. The selected inducer with the optimal blade profile is then optimized and adjusted until the liquid rocket engine turbopump inducer with the best cavitation performance is obtained. This invention can design a liquid rocket engine turbopump inducer with good cavitation performance.

The invention relates to a vertical turbine pump (1) with a riser pipe (5) extending along an axis (4), a motor shaft (6) arranged in the riser pipe (5), a manifold housing (2) with a manifold (3) arranged therein and connected to the riser pipe (5) on the pressure side, a lantern housing (11) arranged on the manifold housing (2) and receiving the motor shaft (6), and a motor (12) arranged on the lantern housing (11) and driving the motor shaft (6), wherein the manifold housing (2) and / or the lantern housing (11) is designed without a specific shape and has a rectangular cross-section.

This invention discloses a method and apparatus for determining the residual axial force of an engine turbopump rotor system, relating to the field of engine turbopump technology, to solve the problem in the prior art that it is impossible to evaluate and predict the residual axial force of an engine turbopump during the design stage. The method includes: obtaining the structural parameters and target operating condition parameters of the target engine turbopump; performing one-dimensional simulation calculations on the target structure in the target engine turbopump based on the target structure and operating condition parameters to obtain the axial force of the target structure; determining the axial force of the balance cavity in the dual-clearance balancing system based on the initial axial clearance and the pressure within the balance cavity; and determining the residual axial force of the target engine turbopump rotor system based on the axial force of the target structure and the axial force of the balance cavity in the dual-clearance balancing system. This achieves efficient and accurate evaluation and prediction of the residual axial force of the engine turbopump rotor system during the design stage.

A bearing arrangement for guiding a motor shaft for supporting axial and / or radial forces of a vertical turbine pump. The bearing arrangement includes a bearing comprising an inner ring which can be non-rotatably connected to the motor shaft and an outer ring, an oil sump filled with an oil, within which the bearing is arranged at least partially immersed in the oil, and an inner oil deflector which is non-rotatably connected to the inner ring and deflects oil conveyed between the inner ring and the outer ring back into the oil sump as a result of rotation of the bearing.

This invention discloses a low-temperature, high-speed, dual-station end-face seal testing device and method for hydrogen-oxygen engine turbopumps. The device includes a housing, a main shaft, rotating ring I, rotating ring II, stationary ring assembly I, and stationary ring assembly II. Stationary ring assembly I and stationary ring assembly II are symmetrically mounted on the housing and respectively contact the two rotating rings mounted on the main shaft to form a dual-station end-face seal. This testing device and method are applied to running-in tests, sampling tests, and research tests of low-temperature, high-speed end-face seals for hydrogen-oxygen engine turbopumps. It can simultaneously and quantitatively measure the leakage of two sets of end-face seals, evaluate the sealing performance of the two sets of end-face seal products, double the testing efficiency, and halve the testing cost.

The present invention discloses a multi-stage self-priming centrifugal vortex pump, comprising a water pool, characterized in that: a pipeline one is arranged above the water pool, a turbine pump one is arranged at one end of the pipeline one, an air pump assembly is arranged on one side of the turbine pump one, an output shaft is arranged in the middle of the turbine pump one, a motor is arranged on one side of the output shaft, a support column is arranged below the outer shell of the turbine pump one, a base is arranged at the bottom of the support column, the support column and the base are fixed structures, and the base is fixed to the ground, a protective shell is arranged above the turbine pump one, a turbine pump two is arranged inside the protective shell, an air pump channel is arranged on one side of the turbine pump two, a pipeline two is arranged above the pipeline one, a ball valve is arranged above the pipeline two, and a hydraulic one is arranged inside the ball valve. The present invention has the characteristics of strong practicality and convenient operation.