31 results about "Turbomachinery" patented technology

The present disclosure provides a kind of blade for inhibiting the blade of leakage vortex break, blade assembly, impeller machinery and blade shaping method.The blade profile is defined according to the cross section profile of the height direction accumulation, the cross section profile has the ratio of the height of the cross section profile and the total height of the blade, the dimensionless blade height H=1 is the blade tip, the dimensionless blade height H=0 is the blade root;The cross section profile also has the back bending angle δ;At the dimensionless blade height H=1, the back bending angle δ=13 °, and has the tolerance range of the numerical ratio ±20%.

Graphene-based sliding varnish and its use in turbomachinery manufacturing. The present invention relates to a sliding varnish formulation comprising a water-based carrier suspension, an inorganic binder, and graphene. This varnish formulation is used to protect turbomachine parts from contact wear in assemblies, particularly in the blades of the hot section of a reactor. Figure for the abstract: Fig. 3

Turbomachine arrangement (10) comprising an electric machine (19) having a shaft (20), at least one compressor section (12a, 12b) for increasing the pressure of a working medium, wherein each compressor section (12a, 12b) has a compressor shaft (13a, 13b), and / or at least one expander section for expanding the working medium, wherein each expander section has an expander shaft, wherein the electric machine (19), the at least one compressor section (12a, 12b) and / or the at least one expander section are arranged in a common, hermetically sealed housing (21), wherein the shafts, i.e., the shaft (20) of the electric machine (19), the respective compressor shaft (13a, 13b) and / or the respective expander shaft, are all completely enclosed within the hermetically sealed housing (21) and are coaxial with each other, and wherein the shafts in the hermetically sealed housing (21) are supported by active magnetic bearings (22) are stored,with at least one electrical current source (24a, 24b) arranged outside the hermetically sealed housing (21), which is configured to electrically energize the windings of the electrical machine (19) and / or the magnetic bearings (22) for heating an interior space (23) bounded by the hermetically sealed housing (21) when the electrical machine (19) is stationary, wherein the same has at least one first electrical current source (24a) for energizing the windings of the electrical machine (19), and wherein the same has at least one second electrical current source (24b) for energizing the magnetic bearings (22), wherein the at least one first electrical current source (24a) is configured to electrically energize the windings of the electrical machine (19) such that a first phase (U) of the electric current supplied by the respective first electrical current source (24a) has a first current amplitude,a second phase (V) of the electric current supplied by the respective first electric current source (24a),

A turbomachinery installation for an offshore platform is disclosed. The turbomachinery installation comprises a train system having at least one preassembled power module, and a torque tube baseplate, on which the train system is installed. A supporting baseplate of the power module is coupled with the torque tube baseplate. Also disclosed is a method of assembling a turbomachinery installation.

The application discloses a kind of aerodynamic damping analysis method based on influence coefficient method and vibration phase correction, it is related to the technical field of aerodynamic damping analysis of turbomachinery.The technical solution points include the following steps: according to the aerodynamic profile and structural model of blade, fluid domain model and solid domain model are established respectively;Carry out steady flow field analysis on fluid domain model, and provide initial field for non-steady flow field calculation;Modal analysis is carried out on solid domain model, and the natural frequency and modal shape of blade under rotating condition are obtained;Designate intermediate blade to vibrate according to the modal shape and natural frequency obtained, and carry out non-steady flow field analysis with steady flow field result as initial field;After calculation convergence, extract the aerodynamic force on blade surface, and calculate the aerodynamic work on blade under different pitch diameters.The application can give the situation under all pitch diameters by once non-steady flow field calculation, without multiple calculation of non-steady flow field under different pitch diameters, more quickly and efficiently, save calculation resources.

This invention relates to the field of turbomachinery technology, and provides a geometric construction method and system for a two-dimensional blade profile in tandem, comprising: using the pressure surface profile of the front row of blades as a reference, at the starting point... P End point of front pressure surface T PS An arc length coordinate system is established between the points, and a channel width function is defined. The channel generation segment profile of the rear blade suction surface is generated by offsetting along the outer side of the normal at each point. A free-form segment profile is then constructed that smoothly connects to the channel generation segment, resulting in the complete rear blade suction surface. Subsequently, the rear pressure surface profile is generated by offsetting along the inner side of the suction surface normal according to a preset thickness distribution function, and smooth leading and trailing edge connection curves are constructed respectively, completing the two-dimensional airfoil structure of the rear blade. This invention combines the slit channel width distribution with the rear blade suction surface generation process, ensuring the geometric continuity of the channel segment while improving the freedom, parameterization, and geometric adjustability of the airfoil design, facilitating aerodynamic design, optimization design, and engineering applications.

An integrated gear-driven turbomachinery system (200) includes a wheel gear (90) configured to rotate about a rotation axis (R) and at least one pair of pinion shafts (10, 20) mechanically coupled to the wheel gear (90). A first pinion shaft (10) is configured to be mechanically coupled to the wheel gear (90) and to rotate about a first axis (X) parallel to the rotation axis (R) at a first rotational speed, and a second pinion shaft (20) is configured to be mechanically coupled to the first pinion shaft (10) and to rotate about a second axis (Y) at a second rotational speed; the first axis (X) and the second axis (Y) are non-parallel.

This utility model discloses a blade structure with aerodynamic damping function in turbomachinery, belonging to the technical field of blade structure technology. It includes a blade, a swing groove, a damping block, a connecting hole, and damping air holes. The swing groove is located at the root of the blade, and the blade has a matching connector for the turbomachinery. The connecting hole is located within the connector. The damping block includes a connecting part, a swing part, and a damping part, all made of elastic material. The connecting part is inserted into the connecting hole, and the damping part is located within the swing groove. The connecting part and the damping part are connected by the swinging part, and the top of the damping part is fixed to the top of the swing groove. Multiple damping air holes are provided through the blade at positions corresponding to the swing groove. This utility model dissipates energy through the elastic swing of the damping block and the damping air holes, without affecting the structural integrity of the overall impeller in the turbomachinery, and suppresses high-order modal vibrations.

The application discloses a kind of based on rough contact interface turbomachinery center pull rod rotor modeling analysis method, belong to turbomachinery high temperature rotor technical field.The present application includes calling pre-trained contact stiffness prediction model, input current operating condition roughness parameter and load parameter, obtain the predicted stiffness of contact interface;The three-dimensional finite element model of rotor is established, the mesh division is carried out to contact interface area, and nonlinear spring damping unit is set between the corresponding nodes of contact interface;The predicted stiffness obtained is given to the nonlinear spring damping unit, and the reduced degrees of freedom are carried out to the finite element model after stiffness assignment using hybrid modal reduction method, and the reduced model is solved for vibration response, the fidelity of the model is ensured, the full model solving degree of freedom is greatly compressed by hybrid reduction method, while not affecting the accuracy of the model, reduce the amount of calculation, facilitate the analysis and solution of model.

This invention relates to the field of aero-engine technology, and discloses an adjustable dual-blade tip clearance simulation test device and method for rotor elastic line reconstruction and fault identification tests. The invention employs a dual independent drive assembly design. During rotor rotation, the axial displacement of the first and second auxiliary bearings is controlled by the first and second drive assemblies, respectively. The axial movement of the first and second auxiliary bearings is converted into radial extension and retraction of the blades via a connecting rod structure, allowing both adjustable blades to extend or retract simultaneously, achieving synchronous adjustment of the tip clearance of both blades. Alternatively, the tip clearance of a single adjustable blade can be adjusted by extending or retracting it individually. This invention can meet the dynamic adjustment requirements of tip clearance in rotor elastic line reconstruction and fault identification tests, and can support rotor dynamics research in rotating turbomachinery.

The invention relates to a system of epicyclic gear trains for a turbomachine, comprising: i) a first epicyclic gear train (10) intended to provide a mechanical reduction function between a first shaft (9) connected to a turbine and a second shaft (14) connected to a fan, and comprising: - a sun gear (10a) connected to the first shaft, - a ring gear (10b) connected to the second shaft, - several planet gears (10c) meshing with the sun gear and the ring gear, ii) a second epicyclic gear train (20) comprising electrical machines (22) each having a rotor portion (22b) and a stator portion (22a), the rotor portions constituting a first assembly and the stator portions constituting a second assembly, one of the two assemblies being connected to the ring gear (10b) of the first gear train while the other of the two assemblies is intended to be connected to a fixed part (18b) of the turbomachine. Fig. 2.

This application discloses a method, apparatus, storage medium, and electronic device for testing the vibration of a turbine blade. The method includes: controlling an excitation device to apply electromagnetic excitation to a rotating blade; acquiring a key phase pulse timing sequence and actual arrival time data of the blade passing each non-contact blade tip timing sensor using a non-contact blade tip timing sensor, the key phase pulse timing sequence being used to indicate the phase reference position during blade rotation; constructing a measured displacement vector based on the key phase pulse timing sequence and actual arrival time data; constructing a frequency and pitch diameter number search grid based on the theoretical mode range of blade vibration; determining the vibration frequency and pitch diameter number of the blade based on the measured displacement vector and the frequency and pitch diameter number search grid; and performing data fitting on the measured displacement vector based on the vibration modes corresponding to the vibration frequency and pitch diameter number to extract the amplitude and phase of the blade vibration under the vibration modes, thus completing the turbine blade vibration test.

An engine (10) includes a rotor shaft, an air chamber, a lubricant chamber containing lubricant, and at least one sealing device (20) disposed between the chambers and including a radially annular seal (21) associated with a sealing track (31) carried by a dynamic sealing track support (30) fixed to the rotor shaft (12). The dynamic sealing track support (30) includes at least one lubricant distribution channel (37) configured to recover lubricant contained in the lubricant chamber and, in a region of the spray zone located axially outside the sealing track (31) relative to the lubricant chamber (15), to spray the lubricant onto an internal heat exchange surface (35a) by centrifugal force.

A turbomachinery power plant may be too large and heavy for transportation. Therefore, embodiments for separating the complete power plant into roadworthy turbine and generator trailers, as well as for realigning and reconnecting the trailers to operate the power plant at the desired destination, are disclosed.

A mechanical gearbox (10) for an aircraft turbomachine (1), this gearbox comprising: - a sun gear (11), - a ring gear (14), - satellites (12) guided in rotation by bearings (8) and meshed with the sun gear (11) and the ring gear (14), and - a satellite carrier (213) comprising a cage (220) in which the sun gear (11) and the satellites (12) and their bearings (8) are mounted, the cage (220) comprising an annular row of slots (300) extending radially and surrounded by an outer annular portion (400) of the cage (220). Figure for the abbreviation: Figure 7

An integrated combustor nozzle (100) includes a combustion liner (110) extending radially between an inner liner section (106) and an outer liner section (108), the combustion liner (110) including a forward end portion (112), an aft end portion, a first sidewall (116), and a second sidewall (118). The aft end portion (114) of the combustion liner (110) defines a turbine nozzle (120). The combustion liner (110) defines a cavity (126) forward of the turbine nozzle (120). The cavity (126) extends between the first sidewall (116) and the second sidewall (118). An impingement cooling apparatus is positioned within the cavity (126). The impingement cooling apparatus includes a flange (310, 311). The impingement cooling apparatus further includes a plurality of impingement members (302) configured to direct a coolant to impinge on the first sidewall (116) and the second sidewall (118). Each impingement member (302) extends from a respective inlet (313) defined within the flange (310, 311) to a respective closed end (312). A plurality of impingement holes (304) are defined on each impingement member (302).

The utility model provides a kind of recovery hydrogen compressor positive pressure device heating equipment, it is related to compressor accessory technical field.The recovery hydrogen compressor positive pressure device heating equipment includes the shell that is internally provided with first chamber and second chamber;Turbine mechanism is provided in first chamber, and turbine mechanism at least has impeller assembly and transmission shaft;Power generation mechanism is provided in second chamber, and power generation mechanism is connected with transmission shaft;And heating mechanism is provided on positive pressure device, and heating mechanism is electrically connected with power generation mechanism;Wherein, shell is provided with the air inlet pipe and gas outlet pipe of intercommunication first chamber;Air inlet pipe and instrument gas source end intercommunication, and motor instrument end intercommunication of compressor and gas outlet pipe;When air inlet pipe is inhaled instrument gas source wind, turbine mechanism can drive power generation mechanism to generate electric energy, and electric energy is supplied heating mechanism to heat positive pressure device.The utility model simplifies the overall structure of device, and reduces energy consumption.

The utility model provides a pipeline turbine flowmeter, including the pipe body, both sides of pipe body are equipped with the connecting sleeve, be equipped with the turbine mechanism in the pipe body, the fluid of turbine mechanism is driven to rotate by pipe body, the outside of pipe body is equipped with monitoring mechanism for monitoring the rotating speed of turbine mechanism, calculates the flow according to the rotating speed, be equipped with the thread in the connecting sleeve, one connecting sleeve is equipped with the connecting pipe, is equipped with the thread outside the connecting pipe. The utility model discloses the rotation of turbine mechanism under the push of liquefied air, cooperate hall sensor and turbine mechanism, can accurately measure the flow of ultralow temperature liquefied air, satisfy the monitoring demand of ultralow temperature liquefied air rock breaking technology, through the injection and replacement lubricant of lubricating mechanism, effectively reduced the wear and tear of bearing under the ultralow temperature environment, greatly prolonged the service life of bearing, and then improved the reliability and stability of flowmeter.

This invention relates to the field of turbomachinery technology and discloses a method for quantifying the uncertainty of aerodynamic performance due to geometric deviations caused by fouling on compressor blades. The method includes defining a sampling space for fouling variables and operating condition variables and obtaining an initial parameter sample set; generating a three-dimensional geometric model of the fouled blade based on the sample set; performing computational fluid dynamics simulation to extract flow rate, pressure ratio, and efficiency parameters to construct a sample database; training three independent fully connected neural network surrogate models to predict flow rate, pressure ratio, and efficiency using the database; sampling within the operating condition variable sampling space, inputting the flow rate surrogate model to obtain the predicted flow rate, and comparing the error with the target flow rate under a benchmark operating condition to obtain the optimal operating condition variable; inputting the parameter combination containing the optimal operating condition variable into the pressure ratio and efficiency surrogate models to output the predicted distribution, and completing the quantification of aerodynamic performance uncertainty through statistical processing. This invention reduces the computational cost of large-scale evaluations and improves the efficiency of quantifying aerodynamic performance degradation under complex fouling conditions.

The present invention relates to a fastening device (10) for fastening a turbomachinery, and more particularly a turbocharger, to a base (11). The fastening device (10) comprises at least one leg (12) that can be fastened to the base (11). The fastening device (10) also comprises a first interconnection element (13) positioned on a first side surface (10A) of the fastening device (10) to establish interconnection and rotational fixed coupling with the corresponding interconnection element (42) of the first casing (41) of the turbomachinery without the use of additional fastening means. The fastening device (10) also comprises a plurality of fastening means receiving portions (14) positioned on the first side surface (10A) for receiving fastening means (18) for fastening the first casing (41) to the fastening device (10). In addition, the fastening device (10) includes a second interconnection element (15) positioned on a second side surface (10B) of the fastening device (10) to establish an interconnection that prevents vertical movement of the second casing (31) of the turbomachinery relative to the corresponding interconnection element (32) without the use of additional fastening means. The fastening device is formed as a single piece.

A method for evaluating turbomachinery blade flutter using a single-channel harmonic balance model includes the following steps: Step 1: Mesh the turbomachinery blades to obtain the vibration modes, vibration frequencies, and vibration pitch diameters of the vibrating blades; Step 2: Mesh the turbomachinery blades into a fluid mesh to obtain the mesh coordinates in a cylindrical coordinate system; Step 3: Align the coordinates of the structural mesh and the fluid mesh to obtain the vibration modes of the fluid mesh; Step 4: Utilize the Doppler frequency shift effect to obtain the blade vibration frequencies, vibration pitch diameters, and inter-blade phase angles of blades in other blade rows besides the vibrating blade; Step 5: Enable the mutual transfer of upstream and downstream information at the rotor-stationary interface; Step 6: Perform single-channel unsteady harmonic balance analysis; Step 7: Evaluate the stability of the turbomachinery blade flutter system. This invention enables the evaluation of flutter in multi-row turbomachinery blades, featuring low computational cost and high evaluation accuracy.

The application discloses a kind of magnetic bearing high-efficiency model predictive control method and device, it is related to magnetic suspension bearing control technical field.For the problem that traditional model predictive control is large in online calculation, real-time is poor in magnetic bearing system, it is difficult to meet engineering application demand, the application is solved by constructing linear state space model of magnetic bearing, and efficient model predictive control strategy of combination of offline-online is designed: solving state feedback control gain and stable attractor in offline stage, only optimizing free variable in online stage, greatly reduce calculation complexity.The application is ensured by augmenting state space, designing Lyapunov function and using Schur complement theorem, stability and dynamic performance of system, expand state attractor, applicable to high-precision control of magnetic bearing in high-speed motor, turbomachinery and other scenarios, with the advantages of high calculation efficiency, strong robustness, good engineering realizability.

This invention provides a turbomachinery design method, system, computer equipment, and medium based on a large language model and multiple agents, relating to the field of intelligent design and engineering technology. The method includes: receiving design requirements information for the turbomachinery in natural language form; at the agent layer, invoking a large language model to parse the design requirements information into multiple task nodes and generating empirical initial design parameter values ​​corresponding to each task node; for each task node, at the agent layer, different agents invoke the large language model and a knowledge database to query and interact, generating interpretable design suggestions for each agent to execute based on the query results, thereby completing the design tasks of each task node; during the design process, iterative optimization of the design is performed based on a hierarchical strategy of first performing local three-dimensional optimization, followed by one-dimensional and two-dimensional redesign. This scheme improves the intelligence and efficiency of compressor design.

The reducer is provided with a ring comprising at least one cylindrical body (21) having a first lateral face (23), a second lateral face (24), and an external peripheral face (25) connecting said lateral faces, and at least one retaining ring (22) received in an annular groove (28) formed in said external peripheral face. Figure for the abbreviation: Fig 3

The present application relates to constant temperature and humidity equipment field, especially to constant temperature and humidity equipment for photoetching machine processing, the body is internally configured with cold water coil pipe, hot water coil pipe, equipment box, electric heating plate and humidifier are sequentially configured above the cold water coil pipe. In the present application, only needs to concentrate in the heating cabin after heating, the boiling clean water circulating in the cabin circulates among the equipment box, hot water coil pipe and preheating coil pipe under the action of water pump assembly, so as to improve the overall water temperature in the body and improve the heating effect. At the same time, through the driving turbine mechanism of boiling gas, the preheating coil pipe heat is delivered to the air inlet air bellow to realize preheating, effectively cooperates with the equipment waste heat to form multiple heating mechanism, effectively reduces the overall heating power consumption of the equipment under low temperature working condition, and ensures the stable preheating effect.

A heat exchanger for a turbomachine, configured to heat a first fluid from a second fluid, comprising a shell delimiting an internal volume with a first internal circulation space configured to convey the first fluid, a second internal circulation space configured to convey the second fluid, the heat exchanger comprising a plurality of tubes, the first internal circulation space comprising an inner area of ​​the plurality of tubes (2), characterized in that the tubes are each surrounded by a heat exchange sleeve, the second internal circulation space comprising an outer area of ​​the heat exchange sleeves. Figure 5

A geared gas turbine engine comprises a heat management system configured to provide lubrication and cooling to a power gearbox and turbomachinery bearings, and comprising a pipe assembly adapted to provide a lubricant flow to the power gearbox and turbomachinery bearings to remove the heat generated by the power gearbox and turbomachinery bearings, an air-lubricant heat exchanger to dissipate a first amount of heat, and a fuel-lubricant heat exchanger to dissipate a second amount of heat wherein the heat management system is configured to provide the first amount of heat and the second amount of heat such that at cruise conditions a proportion of heat generated by the gearbox and the turbomachinery and dissipated to air is in the range of from 0.35 to 0.80.

This invention relates to the field of turbomachinery data processing technology, and particularly to a high-precision non-contact blade vibration signal acquisition and feature extraction method and system. The method involves: step S1 to acquire the sensing distance; step S2 to acquire the blade arrival time and theoretical arrival time; step S3 to fit the blade vibration signal; step S4 to extract features from the blade vibration signal; step S5 to perform real-time correction during feature extraction; step S6 to analyze the processed blade vibration features; step S7 to increase the rationality of the blade vibration parameter model; step S8 to reduce the impact of excessively long blade usage time on the accuracy of result confidence determination; and step S9 to optimize the preset blade usage time. The system includes a signal acquisition module, a signal acquisition module, a feature extraction module, a real-time monitoring module, and a feedback optimization module.

This invention relates to the field of turbomachinery performance evaluation technology, and discloses a method for quantitatively evaluating the performance uncertainty of turbomachinery blade geometric deviations. The method includes parametrically modeling geometric deviation characteristics and sampling to construct a sample; generating a three-dimensional blade simulation model based on the sample by performing deviation modeling on the original design blade profile; performing aerodynamic simulation calculations on the simulation model; extracting aerodynamic indicators using a condition control strategy to construct a mapping database; training a surrogate model using the database to establish mapping relationships; and finally using the surrogate model to perform aerodynamic performance inference to complete the evaluation. Furthermore, interpretable machine learning methods are combined to calculate contribution to identify geometric features highly sensitive to aerodynamic performance. This invention overcomes the limitation of two-dimensional evaluation in failing to consider radial flow effects by reconstructing a three-dimensional simulation model, and eliminates flow deviation operating errors by employing a condition control strategy, providing a quantitative decision-making basis for robust design, machining tolerance setting, and cleaning and maintenance schemes for turbomachinery.

This invention discloses a collaborative design method for orthogonal blades and streamlined endwalls of a high-pressure ratio centrifugal compressor, relating to the field of turbomachinery aerodynamic design technology. The method includes: determining the aerodynamic parameter boundary conditions and geometric constraints at the inlet and outlet of the return flow unit according to the overall compressor design requirements; establishing a parameterized blade model, applying geometrically orthogonal constraints at the intersections of the blade leading edge with the hub and wheel cover to generate meridional aerodynamic orthogonal blades; establishing a parameterized endwall model, defining the meridional profiles of the hub and wheel cover using higher-order curves, and introducing nonlinear control points to adjust the cross-sectional area distribution of the flow channel; constructing a collaborative optimization platform and setting an objective function; and combining three-dimensional numerical calculations and optimization algorithms to iteratively optimize the blade orthogonal parameters and endwall shape parameters until the objective function converges, obtaining the optimal design scheme. This invention solves the problems of severe corner separation and poor inter-stage matching caused by acute angle effects and monotonous endwall shapes in existing technologies.