166 results about "Turbopump" patented technology

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.

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 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 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 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 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 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.

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.

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.

To provide a system comprising a process module, a pump, and an adapter plate, the process module having an opening at a lower end and including a poppet valve above the opening to process a semiconductor substrate.SOLUTION: A pump 102 operates in conjunction with a poppet valve 110 to evacuate gas from a process module 100. An adapter plate 120 is arranged above the pump and below an opening at a lower end of the process module and includes an inner cavity coincident with the opening and an outer perimeter smaller than the lower end of the process module. The adapter plate includes a first set of bores arranged around the inner cavity and a second set of bores arranged along the outer perimeter. Bolts 122 fasten the pump to the adapter plate through the first set of bores. Bolts 124 fasten the adapter plate to the process module through the second set of bores.SELECTED DRAWING: Figure 2

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.

The invention discloses a supersonic axial flow turbine blade profile design method for a turbine pump, and belongs to the field of turbine blade design. According to the method, a Laval spray pipe is replaced by a supersonic static blade designed based on a characteristic line method, so that an accelerated supersonic flow surface can be matched with an annular channel of a movable blade inlet, and circumferential uniform airflow is provided for the turbine movable blade inlet; and therefore, the pneumatic loss generated by air inlet degree configuration of similar parts in the turbine moving blades is reduced. The blade profile of the ultrasonic quick-acting blade designed based on the characteristic line and the free vortex does not depend on design experience, and the stability of the performance of the designed movable blade can be guaranteed. The blade profiles of the supersonic static blades and the moving blades designed through the method remarkably improve the performance of a supersonic turbine in a turbine pump, the fuel carrying amount of a rocket is reduced, and then the thrust-weight ratio of the rocket is increased.

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 invention relates to a liquid rocket engine turbine disc pitch diameter vibration monitoring method, and belongs to the technical field of liquid rocket engine testing. According to the invention, high-frequency speed pressure sensors are additionally arranged at front and rear cavity parts at two sides of the turbine disc, and high-frequency pressure pulsation of fluid near the turbine disc is measured in real time, so that the pitch diameter vibration inherent frequency and the vibration magnitude of front and rear traveling wave frequencies of the monitored turbine disc in the high-temperature and high-rotating-speed operation process are obtained; therefore, a basis is provided for health judgment of the working state of the turbine pump after test run. The method can accurately reflect the pitch diameter vibration level of the thermal test rocket engine turbine disc, has high sensitivity, and solves the problem that the local vibration frequency of the rotor turbine disc is not easy to monitor during the vibration of a traditional turbine pump shell.

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 utility model discloses a bag type sealing structure and a space turbine, and belongs to the field of space engines. The bag type sealing structure comprises a structure body. The first baffles are arranged on the structure body in a protruding mode and arranged at intervals in the first direction, and the structure body and the first baffles are matched with each other to form a bag-shaped sealing structure; in the first direction, the two adjacent first baffles and the structure body jointly form a first cavity. And the second baffle is located in the first cavity and can prevent the gas from moving in the circumferential direction. The bag type sealing structure can seal the turbine pump, prevent gas from circumferentially moving in a sealing gap, resist fluid cross exciting force and effectively improve the stability of a rotor system of the turbine pump.

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.

The invention discloses a self-adaptive locking hydraulic coupler for a transmission chain of a wind turbine generator, and belongs to the technical field of wind power generation. The device comprises an electromagnetic valve, a guide wheel, a turbine, a pump wheel, a signal monitoring controller, an input shaft, a locking piston, a damping assembly and a sensor. The sensor monitors equipment vibration and temperature parameters in real time and transmits the parameters to the signal monitoring controller through a cable. The signal monitoring controller controls the displacement direction of the electromagnetic valve according to monitoring data, so that the joint state of the locking piston is adjusted, and self-adaptive switching between rigid transmission and flexible transmission is achieved; the hydraulic coupler alarm assembly receives an alarm signal of the signal monitoring controller through a cable, and real-time reporting of a fault is achieved. According to the device, the design that a low-viscosity synthetic oil medium is matched with a light-weight corrosion-resistant material shell is adopted, hydraulic transmission loss and impact damage are effectively reduced, and the service life is remarkably prolonged.

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.

An apparatus for providing plasma processing is provided. A plasma processing chamber is provided. A first turbopump with an inlet is in fluid connection with the plasma processing chamber and an exhaust. A gas source provides gas to the plasma processing chamber. At least one gas line is in fluid connection between the gas source and the plasma processing chamber. At least one bleed line is in fluid connection with the at least one gas line. At least one gas line valve is on the at least one gas line located between, where the at least one bleed line is connected to the at least one gas line and the plasma processing chamber. At least one bypass valve is on the at least one bleed line.

The utility model relates to the technical field of reliability testing devices, and discloses a turbo pump reliability testing device, which comprises a testing stand main body, a three-branch inlet pipe and a heating box, the top of the testing stand main body is fixedly connected with a turbo pump body, a main body pipeline of the three-branch inlet pipe is fixedly connected to one side of the turbo pump body, and the three-branch inlet pipe is fixedly connected to the other side of the turbo pump body. Pipelines at the other two ends of the three-branch inlet pipe are fixedly connected with a first control valve and a second control valve respectively, one end, far away from the three-branch inlet pipe, of the first control valve is fixedly connected with a first pipeline, and the heating box is located at the bottom of the test bed main body. When the turbine pump body is used for gas (liquid) high-temperature corrosion resistance test, the first control valve opens the pipe orifice at one end of the three-branch inlet pipe, the turbine pump body is started to extract gas (liquid) in the storage tank, the temperature of the heating box is regulated and controlled at the same time, and the gas (liquid) flows into the heating pipe, so that the heater heats the heating pipe.

The utility model discloses a turbine and a liquid rocket engine turbine pump applying the turbine, and relates to the technical field of liquid rocket engine turbine pumps. The turbine comprises a turbine shell, a turbine disc, a guide ring and a turbine outlet pipe; the turbine outlet pipe is fixed to the turbine shell, and the turbine shell and the turbine outlet pipe are provided with mortise structures used for radial positioning. The guide ring is located in the turbine outlet pipe and fixed to the turbine shell. An outer spigot used for radial positioning is arranged between the guide ring and the turbine shell. The turbine disc is located in the guide ring, and a labyrinth sealing groove corresponding to the turbine disc is formed in the inner side of the guide ring. The turbine, the guide ring and the turbine shell are positioned through the outer seam allowance, the turbine shell and the turbine outlet are positioned through the mortise structure, mounting and dismounting are convenient, the labyrinth seal groove structure is arranged on the guide ring, the high-temperature gas leakage rate can be reduced, and the turbine efficiency is improved.

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.

The invention relates to a device and a method for solving excessive leakage rate of a mechanical seal of a turbine pump. The device comprises a turbine pump body, an air supply device, a diaphragm box type compensation mechanism, a rotating device and a leakage monitoring device, a mechanical sealing cavity is formed in the turbine pump body, a static ring and a rotor structure comprising a movable ring are arranged in the mechanical sealing cavity, the static ring comprises a first sealing face, and the movable ring comprises a second sealing face. The first sealing face and the second sealing face are attached and sealed. And the rotor structure rotates to drive the movable ring to rotate, so that the second sealing surface is in running-in with the first sealing surface. The gas supply device is used for introducing test gas into the mechanical seal cavity so as to form a gas pressure field in the cavity; the diaphragm box type compensation mechanism at least partially matched with the static ring is driven by a gas pressure field, and the first sealing face of the static ring is evenly pressed on the second sealing face of the movable ring. The rotating device is connected with the rotor structure and drives the rotor structure to rotate; the leakage monitoring device is used for detecting the leakage amount of the test gas in the mechanical sealing cavity.

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.