128 results about "Turbopump" patented technology

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 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.

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.

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.

The utility model provides a variable-cross-section special-shaped gas exhaust pipe, a liquid rocket engine and a liquid rocket. The variable-cross-section special-shaped gas exhaust pipe comprises an inlet flange arranged at an inlet, an outlet flange arranged at an outlet and a barrel section arranged between the inlet flange and the outlet flange. The inlet and the outlet are eccentrically arranged; the inlet flange is connected with a turbine, and the outlet flange is connected with the thrust chamber gas collecting ring; high-temperature fuel gas enters from the inlet, flows out from the outlet, enters the thrust chamber gas collection ring and then is introduced into the spray pipe extension section. The exhaust pipe is connected between the turbine and the thrust chamber gas collecting ring, the eccentric layout design is adopted, and the problems that in the structural layout of an existing liquid rocket engine, the center distance between a turbine pump and a thrust chamber is large, the radial envelope of the whole machine is large, and the distance between a swing hose and the swing center is large are solved.

A vacuum pump is proposed that has favorable back pressure performance even in a case where an intake gas is a hydrogen gas or the like. A vacuum pump 100 includes an exhaust portion 114 constituted of a turbine pump portion 115 that includes a plurality of stages of rotor blades 102 and a plurality of stages of stator blades 123, and a drag pump portion 116 that is located on a further toward exhaust downstream side than the turbine pump portion 115, exhausts from an outlet port 133 through a gas flow passage FP1 an intake gas sucked from an inlet port 101 by the exhaust portion 114, and includes an introduction flow passage FP3 through which an intermediate introduction gas having viscosity higher than that of the intake gas is introduced, and the introduction flow passage FP3 is connected to the gas flow passage FP1 on the further toward exhaust downstream side than the rotor blade 102, from among the plurality of stages of the rotor blades 102, that is closest to the inlet port 101 in the exhaust portion 114.

The utility model relates to oil detection technical field, concretely relates to a kind of oilfield development rapid sampling detection device, including support plate, hydraulic cylinder output end is fixedly connected with connecting plate, motor output end is fixedly connected with gear one through shaft coupling, gear one outside is equipped with chain, drilling rod side wall is fixedly connected with gear two, connecting pipe one intercommunication turbine pump output end with filter inside, connecting pipe two connects sampling tank one with sampling tank two inside.The utility model is when entering sampling tank one after petroleum after filter, start ultrasonic homogenizer, make oil, water, solid particle etc. in sample fully mix by high-frequency vibration, form uniform stable suspension or emulsion, ensure the representativeness and accuracy of subsequent detection, sensor carries out detection to petroleum sample, and the preliminary detection parameter of sample in sampling tank two inside can be quickly viewed parameter display screen.

The invention discloses a supporting and lubricating device for a dynamic test of an engine turbine pump rotor. The supporting and lubricating device comprises a first supporting seat, a second supporting seat, an embedded oil spraying pipe, a quick-insertion oil supply connecting nozzle, a quick-insertion oil spraying bent pipe, a radial quick-insertion connector, a first swing frame and a second swing frame. The first supporting seat is arranged on the first swing frame, and the second supporting seat is arranged on the second swing frame; the first supporting seat and the second supporting seat are positioned on the same horizontal plane and are arranged in parallel; the first supporting seat is an independent rotary body, a first supporting seat axial oil supply hole is formed in the axial direction of the first supporting seat, and a first supporting seat radial oil supply hole and a first supporting seat through oil return hole which are communicated with the first supporting seat axial oil supply hole are formed in the radial direction of the first supporting seat; the high-efficiency, high-reliability and low-cost dynamic test requirements of the ultra-high-speed flexible rotor can be met, a supporting effect can be provided, efficient lubricating and cooling of paired bearings can be achieved in a small gap and a narrow space, and the matching precision with a dynamic test system is high.

The invention discloses a double-thrust-chamber pump rear-swinging liquid rocket engine, relates to the technical field of rocket engines, and provides a technical scheme capable of effectively making up for the defects of a traditional single-turbine pump and single-thrust-chamber structural layout in a high-thrust or heavy-duty liquid rocket engine. The invention discloses a double-thrust-chamber pump rear swing liquid rocket engine which comprises a rack, two fuel gas swing devices, two thrust chambers and a fuel gas assembly. The fuel gas assembly communicates with the corresponding thrust chambers through the two gas swing devices and is used for providing fuel for the two thrust chambers correspondingly. The first ends of the two fuel gas swing devices are fixedly arranged on the two sides of the rack correspondingly, the second ends of the two fuel gas swing devices are fixedly connected with the corresponding thrust chambers correspondingly, and the two fuel gas swing devices and the two thrust chambers are symmetrically distributed along the center line of the rack. The fuel gas assembly is fixedly arranged in the middle area of the rack.

Turbopump assemblies and pump turbine assemblies utilized in rocket propulsion systems or other systems are disclosed. The turbopump assembly may include a rotor with at least one integral pump impeller and at least one integral turbine impeller, a housing containing the rotor, and first and second bearings supporting the rotor. The first and second bearings are supported by the housing. The rotor is disposed between the first and second bearings such that the at least one integral pump impeller and the at least one integral turbine impeller are supported between the first and second bearings. At least one integral pump impeller and at least one integral turbine impeller are provided between all bearings supporting the pump and turbine in the housing.

The invention discloses a low-temperature medium rapid supercharging test operation system and method, and belongs to the field of application of seals for turbine pumps of liquid rocket engines. The system comprises a rapid pressurization gas distribution unit, a medium rapid supply control unit and a pressure feedback rotating speed regulation driving control unit, through double-path pressurization of a large-flow regulating valve and a pore plate and by utilizing the characteristics of the dual-path pressurization of the large-flow regulating valve and the pore plate, accurate pressure stabilization is achieved through the large-flow regulating valve after rapid pre-pressurization is achieved through the pore plate, and rapid supply and accurate regulation of test pressure are achieved; meanwhile, based on the rapid pressurization technology, the rotating speed of the sealing product in the starting process changes along with pressure in a pressure feedback rotating speed adjusting mode, and the pressure is matched with the rotating speed.

This invention belongs to the field of liquid rocket engines, specifically disclosing a cryogenic sealing tester for turbopump end faces, comprising a shell and a rotor. The shell includes a first support, a second support, and an outer shell. The first support is provided with a tooling end face sealing stationary ring, and the second support is provided with a test piece end face sealing stationary ring. The outer side of the first support is provided with a transparent cover, and the outer side of the second support is provided with a blind cover. The outer shell is provided with a liquid inlet and a vent. The first and second supports are respectively provided with a first liquid outlet and a second liquid outlet. The rotor includes a first bearing, a second bearing, and a shaft. The outer side of the first bearing is provided with a tooling end face sealing dynamic ring, and the outer side of the second bearing is provided with a test piece end face sealing dynamic ring. There is a dynamic seal between the tooling end face sealing dynamic ring and the tooling end face sealing stationary ring, and a dynamic seal between the test piece end face sealing dynamic ring and the test piece end face sealing stationary ring. This invention can meet the requirements of high speed, high precision, and suitability for ultra-low temperature environments for turbopump end face cryogenic sealing testers.

A representative seal system (such as a seal system for a turbopump of a rocket engine) automatically adjusts a balance ratio of a seal. The system can include a ring element encircling an axis. A front side of the ring element contacts a revolving surface to form a seal with the revolving surface. The front side can include a stepped surface having two or more steps. Each step includes a sealing surface configured to contact the revolving surface to form a sealing area that is different from a sealing area of each other sealing surface. Each step is positioned and configured to wear away during operation of the machine to expose an underlying surface to the revolving surface, to change the sealing area and the balance ratio of the seal. A representative method of operating a turbomachinery system includes changing the balance ratio of a seal while rotating a rotor.

The invention relates to the technical field of exhaust emission, in particular to an external exhaust emission device, a test system and an installation method. A waste gas external connection discharging device comprises a discharging elbow pipe which is suitable for being arranged in a test workshop and comprises an air inlet elbow and a horizontal air outlet section, and the air inlet elbow is suitable for being connected with the exhaust end of a turbine pump; the discharge pipe is connected with the discharge elbow pipe in an inserted mode, the end, away from the discharge elbow pipe, of the discharge pipe is arranged outside the test workshop, part of the horizontal air outlet section is arranged on the periphery of the discharge pipe in a sleeving mode, or part of the discharge pipe is arranged on the periphery of the horizontal air outlet section in a sleeving mode, and a gap is reserved between the horizontal air outlet section and the discharge pipe; and the buffer layer is arranged in the gap. The invention provides an external exhaust gas emission device, a test system and an installation method, and aims to solve the problem that potential safety hazards exist due to resonance between a turbine pump and an emission pipeline caused by the fact that the end part of the emission pipeline is influenced by vibration of the turbine pump when a gas emission device is connected with a gas outlet to emit gas outdoors.

The invention discloses a reusable high-reliability floating ring sealing device for a turbopump, belongs to the technical field of sealing structures of turbopumps of liquid rocket engines, and aims to avoid the problem of axial interference caused by eccentricity of a floating ring during blind mounting of a mounting shell and improve the assembly manufacturability of a product through a conical sealing surface; meanwhile, the floating ring forms a conical liquid film sealing structure with self-adjusting capability in the working process; when the turbine pump operates, a medium generates axial and radial compound hydraulic power in the conical sealing gap, so that the floating ring has automatic floating and force balancing capabilities in the axial direction, and the problem that a traditional unthreaded hole type floating ring is prone to eccentric wear or failure due to unbalanced axial stress is solved; the problems that in the prior art, a floating ring sealing assembly does not have the axial force adjusting and balancing function, the liquid floating capacity is insufficient, the requirement for assembly precision of a fixed inner diameter gap structure is high, and manufacturability is poor are solved, and the floating ring sealing assembly is suitable for repeated disassembly and assembly and long-term reliable use.

The invention discloses a high-thrust extrusion type rocket manufactured based on a mature industrial system. The high-thrust extrusion type rocket comprises a fuel storage tank, an oxidizing agent storage tank, a liquid nitrogen storage tank, an exhaust pipe, an oxidizing agent injector, a fuel injector and a thrust chamber. The nitrogen turbine pump is only used for pressurizing inert gas, an oxidizing agent and fuel are directly extruded into the thrust chamber, the risk of explosion caused by rapid oxidation or leakage is avoided, meanwhile, the turbine is driven by high-temperature fuel gas extracted from the thrust chamber, a combustor is not needed, the structure of the turbine pump is simplified, and the cost is reduced. In addition, the nitrogen turbine pump is driven by high-temperature gas and used for pressurizing nitrogen, batch production or direct purchase of the nitrogen turbine pump can be achieved depending on an existing mature industrial system, and the manufacturing cost of the nitrogen turbine pump is remarkably reduced; in addition, under the pressurization effect of the nitrogen turbine pump, the propellant is effectively extruded to enter the thrust chamber in a large-flow mode, a high-pressure gas cylinder is not needed, the three storage boxes do not need to be provided with high-pressure thick walls, and the overall structure manufacturing cost of the rocket is remarkably reduced.

The utility model provides a turbine pump assembly and a chemical vapor deposition machine table. The turbine pump assembly comprises a lifting driving mechanism, at least two guide rails and a turbine pump, the lifting driving mechanism supports the at least two guide rails and controls the lifting of the at least two guide rails; the at least two guide rails are connected with the base in a sliding manner; the base is provided with a containing cavity for containing the bottom of the turbine pump. According to the turbine pump assembly, the base for containing the turbine pump is connected with the sliding rail, meanwhile, the sliding rail is arranged on the lifting driving mechanism, and therefore when the turbine pump of the chemical vapor deposition machine table breaks down and needs to be replaced or overhauled, the turbine pump can be conveniently and safely disassembled or assembled, manpower is reduced, and the equipment maintenance cost is reduced.

This application relates to the field of liquid rocket engine technology and discloses a rocket engine turbopump fixing device, comprising: a connector, at least two connecting plates, a first telescopic rod, and a second telescopic rod. The connector has a notch adapted to avoid the turbopump and is adapted to connect to the rocket engine. The two connecting plates are adapted to be fixed to the turbopump, and the portions of the two connecting plates corresponding to the periphery of the notch are detachably connected. One end of the first telescopic rod is hinged to the connector, and the other end is adapted to connect to the turbopump. One end of the second telescopic rod is hinged to the connector, and the other end is adapted to connect to the turbopump. The first and second telescopic rods are located on opposite sides of the connecting plates along a first direction. This rocket engine turbopump fixing device discloses a solution to or improvement of the problems of complex assembly and disassembly and difficult positioning of turbopumps.

The utility model discloses a turbine pump casing, including outer casing and inner casing, and the inner casing inserts the inside of outer casing, and the outer casing includes the liquid inlet and liquid outlet of being located at the outside, and the liquid inlet, liquid chamber, liquid chamber and liquid outlet are communicated, which are formed between the outer casing and the inner casing. The utility model discloses the mode of the outer casing wrapping the inner casing, can realize the cooling flow passage design of casing inside, need not to adopt the external flow channel, greatly reduces the volume of turbine pump whole.

The embodiment of the present application provides an inducer and a turbopump, which are used for further improving the cavitation erosion resistance of the inducer. The inducer is suitable for a liquid rocket engine, and the inducer comprises a rotating shaft and blades fixed to the rotating shaft; a first development surface corresponding to the blades comprises a first region, and a part of the first development surface located in the first region is in a curve shape; the first development surface is a planar pattern obtained by unfolding a first section surface of the blades which is cut by a cylindrical surface of the rotating shaft on a shaft radius. The inducer and the turbopump provided by the embodiment of the present application can further improve the cavitation erosion resistance of the inducer.

A turbo pump low-temperature bearing full-size working condition test system and method belong to the technical field of turbo pump test, the tested bearing is enclosed in a test main body, the test system comprises a liquid nitrogen main line and a nitrogen pressurizing system; the liquid nitrogen main line comprises a liquid nitrogen main storage tank, a liquid nitrogen high-pressure storage tank, a liquid nitrogen subcooler, a liquid supply and return valve group pipeline, the test main body, a liquid nitrogen recovery storage tank and a liquid nitrogen recovery pump group connected in sequence, the liquid nitrogen in the liquid nitrogen main line flows from the liquid nitrogen main storage tank to the liquid nitrogen recovery pump group; the liquid nitrogen recovery pump group of the liquid nitrogen main line is connected with the liquid nitrogen high-pressure storage tank, forming a liquid nitrogen circulating loop; the liquid nitrogen recovery pump group of the liquid nitrogen main line is connected with the liquid nitrogen main storage tank, forming a liquid nitrogen recovery loop. The turbo pump low-temperature bearing full-size working condition test system and method have strong low-temperature simulation capability, strong compatibility of full-size working condition, long test duration, and can meet the low-temperature, full-size and large-flow test requirements of turbo pump low-temperature bearings.

The invention discloses an umbrella-shaped folding ground effect aircraft which comprises an umbrella-shaped folding propeller made of flexible materials or rigid materials, a main driving device is installed on propeller blades, and blade connecting and fixing rings supported by elastic frameworks are arranged on the outer sides of the propeller blades, or rigid multi-wing centrifugal machine fan-shaped blade connecting and fixing rings are arranged on the inner sides and the outer sides of the propeller blades. The connecting and fixing ring blade is provided with a groove type jet engine nozzle or an electric rotor wing type auxiliary driving device, the multi-wing centrifugal machine fan-shaped blade connecting and fixing ring is conical, the conical centrifugal machine fan-shaped blade connecting and fixing ring is connected with a turbine pump or a center shaft, and the propeller blade is connected with an upper tray and a lower tray of the center shaft through an umbrella type framework. A turbine pump or other energy supply devices are installed on the center shaft, a safety air bag is arranged on the lower portion of the lower tray, the center shaft is connected with a load cabin, a tail vane of the load cabin is a telescopic rod, and a pushing device for controlling the direction such as a jet engine or an electric rotor wing is installed at the tail of the telescopic rod.

The invention relates to low-temperature displacement sensor test equipment for a liquid rocket engine, a displacement sensor is installed on a loading mechanism, a measuring target is installed on a test device main shaft, and the test device main shaft is connected with a driving device through a coupler; the driving device drives the main shaft of the testing device to rotate at a high speed, and the testing device is filled with low-temperature media to simulate the real working conditions of the measurement target and the displacement sensor. Wherein the loading mechanism drives the displacement sensor to move axially, real-time dynamic adjustment of the axial distance between the displacement sensor and the measuring target is achieved, and testing of the displacement sensor in a low-temperature state is achieved through displacement signals of the displacement sensor and the contact type displacement sensor on the loading mechanism. According to the invention, the real working environment and rotating speed of the displacement sensor and the target piece in the turbine pump are effectively simulated, and the working reliability of the researched and developed low-temperature displacement sensor is verified through early-stage rapid ground tests, so that a basic support is provided for development, test run, emission and the like of an engine.

Cryogenic turbopump feed line, comprising a main channel able to transport a cryogenic fluid, a plurality of secondary channels parallel to and disposed around the main channel in which the plurality of secondary channels are disposed on the periphery of an external perimeter of the main channel.

This invention relates to the field of aerospace propulsion technology and discloses a gas generator circulating liquid rocket engine system with an electric bypass. This invention constructs an electric bypass supply channel by reusing the discharge pipeline, introducing the originally directly discharged high-pressure propellant into the fuel electric pump and oxidizer electric pump for secondary pressurization, and then injecting it into the main thrust chamber via a separately set bypass injector. This solves the problems of traditional open gas generator circulating liquid engines, such as the main turbopump easily entering the critical speed range and causing mechanical resonance under deep throttling conditions, and the deterioration of atomization mixing effect due to insufficient pressure drop of the main injector. It achieves stable operation of the main turbopump within the rated operating speed range and stable high-pressure atomization and combustion of propellant under extremely low thrust conditions, expands the lower limit of the engine's deep throttling capability, and improves propellant energy utilization efficiency and system operational reliability. Simultaneously, the electric bypass supply system enables rapid response and precise adjustment under low thrust conditions.

The invention relates to a turbopump combined supporting flexible rotor structure of a liquid-propellant rocket engine, and belongs to the technical field of liquid-propellant rocket engines. The invention relates to a flexible rotor structure of a liquid rocket engine with a secondary impeller and a secondary turbine disc. A rotor support is a combined structure of a hydrostatic bearing and a roller bearing; main radial loads are borne by the fluid hydrostatic bearing, the hydrostatic bearing is not abraded when working, the combined support of the fluid hydrostatic bearing and the roller bearing adapts to the working conditions of high rotating speed and high bearing DN value, the service life of the roller bearing is prolonged, the reuse frequency of the turbine pump is increased, and the manufacturing cost and the maintenance cost of the turbine pump are reduced. The required bearing capacity and rigidity range are obtained through design factors such as static pressure bearing structural design, pressure arrangement and gap design, the design is more flexible, the adaptability is wider, the structure is simpler, the economical efficiency is better, and the use maintainability is better.

A guide vane arrangement for use in a turbopump having a first guide vane and a second guide vane. The second guide vane is arranged adjacent to the first guide vane such that a flow channel is defined between a leading surface of the first guide vane and a trailing surface of the second guide vane. The trailing surface of the second guide vane has a trailing portion arranged adjacent to a trailing edge of the second guide vane and which is arranged at a first angle with respect to a trailing edge of the first guide vane and the trailing edge of the second guide vane, a leading portion which is arranged adjacent to a leading edge of the second guide vane and which is arranged at a second angle with respect to the trailing edge of the first guide vane and the trailing edge of the second guide vane, and an intermediate portion arranged between the trailing portion and the leading portion and which is arranged at a third angle. The second angle is larger than each of the first and third angles.

The application relates to a turbine pump automatic overturning device and an assembling process, which comprises the following steps: a servo motor is arranged in a base plate to drive a roller to realize the movement of the automatic overturning device; a turnover assembly is used for clamping and fixing a turbine pump with an eccentric structure; one end of the turnover assembly is connected with a first support through a rotary bearing, and the other end of the turnover assembly is connected with a second support through an RV precision speed reducer; the first support and the second support are fixedly arranged at the two ends of the base plate and are used for supporting the turnover assembly; the RV precision speed reducer is a compact speed reducer, integrates the servo motor and the speed reduction, can flexibly control the speed of the turnover assembly, has the arbitrary angle self-locking function, thereby realizes the fixation of the turnover assembly at an arbitrary rotating angle, and is convenient for the turbine pump assembling. The RV precision speed reducer is used for driving the turbine pump to overturn, the assembling quality and the working efficiency can be effectively improved, and the labor intensity of operators is reduced.

The application relates to a manufacturing method of a turbine pump accessory and the turbine pump accessory, the accessory comprising a turbine end bearing seat, a pump end bearing seat, a positioning shaft, a pump inlet shell and a pump outlet shell; the manufacturing method of the turbine pump accessory comprises the following steps: machining the turbine end bearing seat so that the turbine end bearing seat has first preset parameters; the turbine end bearing seat has a first shaft hole; machining the pump end bearing seat so that the pump end bearing seat has second preset parameters; the pump end bearing seat has a second shaft hole; one of the size precision, the shape and position tolerance and the surface roughness is included in the first preset parameters and the second preset parameters; the turbine end bearing seat is assembled to the pump outlet shell, and the coaxiality and the straightness of the first shaft hole of the turbine end bearing seat are aligned; the positioning shaft is assembled to the turbine end bearing seat; the pump inlet shell is sleeved on the positioning shaft and assembled to the pump outlet shell; and the pump end bearing seat is sleeved on the positioning shaft through the second shaft hole and assembled to the pump inlet shell.

The invention discloses a method and a system for calculating axial force applied to an impeller by a balancing device in a turbine pump, and belongs to the technical field of axial force calculation and axial force balancing of turbine pumps of liquid rocket engines. The method mainly comprises the steps that a rotational flow ratio k and sealing flow resistance characteristic coupling calculation model is established, and pressure distribution in the balancing device and axial force acting on a centrifugal impeller disc are calculated based on novel balancing device structure parameters, turbine pump working condition parameters and fluid physical property parameters. According to the invention, the model and the method for calculating the axial force applied to the impeller disc by the novel balancing device with the vortex breaker are provided for the first time, and the pressure distribution in the novel balancing device and the axial force applied to the impeller disc by the novel balancing device can be quickly and accurately calculated through the one-dimensional static characteristic model; and the capability of predicting, calculating and balancing the axial force of the rotor-impeller-balancing device of the turbopump in engineering design and engineering application is improved.