219 results about "Compressor blade" patented technology

The invention provides a gas compressor blisk rupture rotation speed prediction method based on virtual-real fusion, belongs to the technical field of aero-engines, and particularly relates to an aero-engine gas compressor blisk rupture rotation speed prediction method by establishing a rupture rotation speed-oriented aero-engine gas compressor blisk digital twin model and performing double correction on the model and a failure criterion by using test real-time data. And in the simulation process, a parallel computing technology is used for acceleration, so that a high-precision fracture rotation speed prediction value which aims at a single individual of the aeroengine compressor blisk and accords with physical rules and actual test measurement can be quickly obtained.

The invention discloses a gas compressor blade state monitoring method and system based on a health management platform, and relates to the technical field of gas turbines. Comprising the following steps: acquiring real-time operating parameters of the gas turbine, and transmitting the real-time operating parameters to a health management platform at a fixed sampling frequency; performing data preprocessing and feature extraction on the real-time operation parameter data; calculating the performance parameter degradation quantity of the gas turbine according to the real-time operation parameter data; performing blade fault identification according to the parameter degradation quantity and a preset fault diagnosis criterion; calculating a fault early warning index through the real-time operation parameter data, and judging whether the gas turbine is in a normal operation condition or not; and judging a fault early warning level according to the parameter degradation quantity and sending out a corresponding early warning signal. According to the method, the blade fault type can be accurately predicted, false alarms and false alarms can be reduced, meanwhile, early warning level signals are sent out, measures can be taken, fault expansion is avoided, and downtime and maintenance cost are reduced.

The invention provides a communication and interaction method of a digital twinning system in a compressor blisk machining process, and relates to the technical field of digital twinning and aeronautical manufacturing, and the method comprises the steps: collecting vibration, temperature and acoustic emission data in the compressor blisk machining process in real time, analyzing machine tool machining data, carrying out the alignment and noise reduction of the collected data, and obtaining a digital twinning result; then time-frequency domain mixed feature extraction is carried out, and time-frequency domain mixed feature data are optimized and uploaded in a graded mode; a deterministic transmission network is constructed, data transmission priorities of 5G communication and LoRa dual channels are dynamically adjusted according to a network load, and intelligent switching between 5G communication and a LoRa protocol is realized based on a switching decision model; analyzing normal processing data of the whole stage of the processing process, and establishing a processing data thread; and predicting the abnormal processing risk in the processing process, and updating the model parameters. According to the method, the communication and interaction efficiency of the digital twin system in the machining process of the compressor blisk in the aero-engine can be improved.

The invention belongs to the field of aero-engines and turbines, and provides a quick prediction method, system and device for a two-dimensional viscous compressible flow field of a gas compressor and a medium, and the method comprises the steps: obtaining CFD simulation data of a two-dimensional blade profile of the gas compressor under multiple working conditions, and carrying out the preprocessing to obtain a standardized training data set; on the basis of the data set, constructing a hybrid neural network model which takes a blade profile geometric parameter and an operation condition parameter as input and takes a two-dimensional viscous compressible flow field as output; combining mean square error, gradient loss and physical consistency loss based on RANS equation residual error to construct a total loss function, and performing end-to-end training on the model; during reasoning, to-be-predicted parameters are input into the trained model according to the same preprocessing mode, and then a flow field prediction result can be rapidly output. According to the method, the problems of long consumed time, low precision and poor consistency of traditional CFD simulation can be solved, millisecond-level flow field prediction is achieved, high precision and high physical consistency are achieved, and the method is suitable for efficient optimization design of the blade profile of the gas compressor.

The invention belongs to the field of intelligent manufacturing of aero-engines, and provides a parametric modeling method for structural characteristics of typical parts of an aero-engine. According to the method, efficient precision machining of the aero-engine compressor blade is achieved through an intelligent system of parametric modeling, feature self-recognition and rapid programming. The specific implementation process comprises the steps that a three-dimensional model is generated through a parametric modeling module on the basis of a tenon structure parameter set and a blade body profile geometric feature parameter set input by a user; secondly, a feature recognition module is used for conducting multi-level machining feature decomposition on the three-dimensional model, and key machining features such as a tenon and a blade body profile are precisely recognized through a geometric topology analysis and depth feature learning fusion algorithm; on the basis, intelligent matching and optimization of the machining technology are achieved for the recognized machining features through a rapid programming module, the machining program is optimized through physical simulation, and finally the machining program applied to the production field is generated.

The invention provides a rotating speed determination method in aero-engine complete machine three-dimensional simulation, and belongs to the technical field of aero-engine simulation, and the method specifically comprises the steps: constructing a fluid-structure interaction control equation on each grid unit of an aero-engine complete machine geometric model; forming an integral coupling matrix, and solving the integral coupling matrix; simulating the calculated thermal state geometric model; the gas compressor and the turbine can only rotate around the rotating shaft; calculating aerodynamic forces acting on the compressor blades and the turbine blades; in each time step, the rotating speed is updated according to the angular acceleration; and after a plurality of iterations, when the difference value of the aerodynamic torque borne by the gas compressor and the aerodynamic torque borne by the turbine accords with a preset range, the corresponding rotating speed is the steady-state working rotating speed of the engine under the thermal state geometry. According to the processing scheme, the engine can automatically match and dynamically determine the working rotating speed in the simulation process, the inaccuracy of the preset rotating speed is avoided, and the precision and reliability of a simulation result are improved.

A turbine wheel may include an inner hub having a central axis, an outer ring concentric with the inner hub, and an intermediate ring concentrically disposed between the inner hub and the outer ring, and a plurality of turbine blades extending between the intermediate ring and the outer ring. The turbine blades may be oriented to be driven by a fluid stream flowing through the turbine wheel to rotate the turbine wheel in a first direction about the central axis. In addition, the turbine wheel may further include a plurality of compressor blades extending between the inner hub and the intermediate ring. Also, the compressor blades may be oriented to propel a fluid in a downstream direction when the turbine wheel is rotated in the first direction.

The invention relates to the technical field of fluid mechanics, in particular to an efficient simulation method and system for turbofan engine compressor blade sand dust erosion. The efficient simulation method for sand dust erosion of the blades of the turbofan engine compressor comprises the following steps of performing unsteady calculation on a full-stage and full-ring airflow field of the compressor until the airflow field converges, and extracting airflow field data of moving blade and stationary blade areas of each stage; according to the airflow field absolute speed data of each stage of moving blade area, obtaining airflow field relative speed data of each stage of moving blade area; setting the particle size, the initial flow, the initial position and the initial speed of sand dust at an inlet of the gas compressor; after entering the gas compressor, the sand dust sequentially passes through each stage of stationary blade and movable blade area in the gas compressor at a corresponding movement speed, and the absolute speed data and the relative speed data of the airflow field of each stage of movable blade and stationary blade area are combined to obtain the real-time movement track of the full-stage and full-ring sand dust in the gas compressor. Further obtaining the position, angle and speed of the impact point of the sand dust and the blade, and finally obtaining the erosion abrasion loss of the blade.

The invention discloses an axial flow compressor blade anti-erosion optimization method facing a sand and dust environment, and belongs to the technical field of compressor blade anti-erosion design and optimization. According to the method, on the basis of a gas-solid two-phase flow erosion mechanism in a sand and dust environment, links such as working condition analysis, parameterization platform building, constrained multi-objective optimization method design, model building and optimization and index verification are integrated, and an integrated anti-erosion optimization process is formed. CFturbo, AutoGrid, CFX and Matlab are called in a unified mode through a Python script, automatic coupling of geometric modeling, grid division, gas-solid coupling simulation and an intelligent optimization algorithm of the blade is achieved, and collaborative optimization of aerodynamic performance and erosion resistance can be achieved under the constraint of erosion wear. Compared with a traditional method which depends on experience modification or single simulation, the method has the advantages that the optimization efficiency is remarkably improved while the simulation precision is guaranteed, the personal error risk is reduced, and the method can flexibly adapt to different compressor models and diversified operation conditions.

The invention provides an additive repair blade high-cycle fatigue limit prediction method based on random defect probability distribution, and relates to the field of aero-engine fan and compressor blade maintenance, and the method comprises the steps: obtaining the defect distribution of a real blade; establishing a prediction model based on a critical distance theory; according to the prediction model, critical distances of the multiple real blades are determined, and a first critical distance is generated; performing defect reconstruction based on the defect distribution of the real blade to generate a virtual blade; obtaining the first high-cycle fatigue strength of the virtual blade according to the defect distribution of the virtual blade, the first critical distance and the prediction model; respectively determining a plurality of defect parameters and a Pearson coefficient of the first high-cycle fatigue strength, and screening a first parameter significantly related to the first high-cycle fatigue strength; constructing a fatigue strength evaluation model; and inputting the first parameter into the trained fatigue strength evaluation model, and outputting the predicted high-cycle fatigue strength. According to the scheme, the high-cycle fatigue limit of the blade can be quickly predicted, and the prediction accuracy is improved.

The invention discloses an axial flow compressor blade profile optimization method based on main and auxiliary agent models and a related device, and relates to the field of compressor blade design, and the method comprises the steps: generating a plurality of design parameter groups based on to-be-optimized design parameters and a parameter adjustable range of a to-be-optimized axial flow compressor blade by using a data sampling algorithm, outputting a blade profile performance prediction value corresponding to each design parameter group by using the main and auxiliary combination agent model, screening a plurality of alternative design parameter groups based on a comparison result of each blade profile performance prediction value and an initial performance reference, and performing computational fluid dynamics simulation; and determining the alternative design parameter group corresponding to the optimal simulation performance parameter as a blade profile optimization parameter group. According to the method, the invalid risk of the design parameter group is reduced by using the adjustable range of the parameters, the performance prediction precision is improved based on the flow field characteristic parameters by using the main and auxiliary combination agent model, finally, the blade profile optimization parameter group is determined based on the simulation result of the screened alternative design parameter group, and the prediction precision is further improved.

The invention relates to the field of flow control of turbine mechanical parts of aviation gas turbine engines, in particular to a non-axisymmetric end wall transverse pressure distribution design method and system.The method comprises the steps that the range of a non-axisymmetric end wall profile covers the area from the front edge to the tail edge of a blade in the axial direction; a transverse end wall profile control line is arranged at the separation starting position of the flat end wall blade grid corner area; giving transverse distribution of static pressure coefficient variation on the end wall profile control line; calculating a target static pressure value of the non-axisymmetric end wall at a corresponding axial position based on the transverse distribution of the variable quantity of the static pressure coefficient and the transverse distribution of the static pressure coefficient of the flat end wall cascade at the corresponding position; and performing three-dimensional inverse problem design according to the target static pressure value of the non-axisymmetric end wall at the corresponding axial position to obtain a corresponding non-axisymmetric end wall profile. The method disclosed by the invention can be used as a non-axisymmetric end wall universal design method for improving the flow of the end area of the compressor blade.

The invention discloses a thermal flushing system design method for an aero-engine compressor blade, belongs to the field of aero-engine cleaning, and provides a thermal flushing system design method for removing salt deposited on the surface of the compressor blade for an aero-engine working under the marine environment condition. According to the design method, targeted cleaning of the cleaning fluid on the salt deposited on the surface of the blade of the gas compressor is innovatively provided, and the cleaning efficiency of desalting is improved. According to the method, a complete design process of the aero-engine thermal flushing system is formed.

The invention provides a gas compressor blade titanium fire friction test device and test method. The gas compressor blade titanium fire friction test device is used for testing the flame retardant property of a blade sample piece of a gas compressor, and is characterized by comprising a first rotating piece, a second rotating piece and a friction piece, wherein the friction piece is arranged on the first rotating piece; the blade sample piece is assembled on the second rotating piece, and the blade sample piece and the friction piece are arranged to be opposite to each other; the first rotating part can rotate in a first direction, and the second rotating part can rotate in a second direction opposite to the first direction; in the rotating process of the first rotating piece and the second rotating piece, the friction piece makes contact with the blade sample piece, and then the blade sample piece is rubbed. The test device can better simulate the working environment of the high-pressure compressor.

The invention relates to the technical field of aeronautical materials, in particular to a high-frequency titanium alloy compressor blade fatigue test piece and a high-frequency titanium alloy compressor blade fatigue test method. In order to obtain the fatigue strength of the high-frequency blade test piece, the frequency of the test piece needs to be reduced, so that the vibration stress generated by the high-frequency blade fatigue test piece is large enough. The invention provides a simple and effective fatigue test piece design method for high-frequency titanium alloy compressor blades. Argon arc welding is adopted, a titanium alloy block similar to the shape of a blade crown is welded to the middle of the blade tip of the high-frequency compressor blade, and the frequency of the blade is reduced under the condition that blade stress distribution and the maximum stress position are not changed basically. The invention discloses a method for reducing the inherent frequency of a blade by welding a balancing weight on a blade tip through argon arc welding. The invention discloses a method for reducing inherent frequency without changing basic modal stress distribution of a blade. The invention discloses a method for reducing the inherent frequency of a blade by welding a balancing weight on a blade tip through argon arc welding. The invention discloses a method for reducing inherent frequency without changing basic modal stress distribution of a blade.

A gas turbine engine has a compression system blade ratio defined as the ratio of the height of a fan blade to the height of the most downstream compressor blade in the range of from 45 to 95. This results in an optimum balance between installation benefits, operability, maintenance requirements and engine efficiency when the gas turbine engine is installed on an aircraft.

The present application relates to the field of gas turbine engine compressor test, and discloses a blade non-integer order vibration analysis method and system of multi-physical field dynamic signal, wherein the airflow excitation frequency and the circumferential mode order generated by unstable aerodynamic disturbance are obtained through dynamic pressure test, and the circumferential unstable flow mode characteristics causing blade non-integer order vibration are obtained; the blade vibration frequency and the pitch diameter are obtained through blade vibration dynamic test, and the test parameters of the two physical fields are comprehensively processed to realize the judgment of blade non-integer order vibration. The present application can effectively identify the occurrence of blade non-integer order vibration, can be used for the test of non-integer order vibration of aero-engine compressor blade and the mechanism analysis of related blade vibration failure fault, and has important significance for improving the overall safety and reliability of aero-engine compressor.

The invention discloses a blade surface identification and curved surface reconstruction method based on compressor blade deformation analysis, and belongs to the field of blade thermal deformation analysis. According to the method, the problems of thermal deformation and low accuracy in existing compressor blade reconstruction are solved. The method comprises the following steps: firstly, drawing a blade region as a structured grid, extracting a layer of grid on the surface of a blade, and defining a unit set and a node set; performing blade finite element conventional calculation, extracting corresponding data, and identifying a corner unit and a corner node as an initial unit; according to the topological relation of the units, the adjacent relation between the units is established, and a blade unit arrangement matrix is constructed according to the adjacent relation; determining inner and outer surfaces by using the arrangement matrix and the node relation in the unit; determining the surface of the blade according to the perimeter of a certain layer; and according to the node coordinates and the surface node matrix, expanding and writing into a point cloud format according to rows, and reconstructing a surface body. The method is used for blade thermal deformation reconstruction.

The invention provides a heavy type gas turbine gas compressor aerodynamic strength vibration structure coupling design method and system. The method comprises the steps that a first through-flow parameter and a first blade profile parameter of a gas compressor are determined according to aerodynamic performance indexes; optimizing a first flow parameter and a first blade profile parameter of the gas compressor according to the structure manufacturing and assembling index and the aerodynamic performance index to obtain a second flow parameter and a second blade profile parameter of the gas compressor; optimizing the geometric parameters of the blade root of the gas compressor according to the strength vibration life index and the structure manufacturing and assembling index, and designing and optimizing the geometric parameters of the blade locking structure of the gas compressor to obtain the optimized geometric parameters of the blade root and the blade locking structure of the gas compressor; and designing the heavy gas turbine compressor based on the second through-flow parameter and the second blade profile parameter. According to the technical scheme provided by the invention, the design efficiency is improved, the performance balance is optimized, and meanwhile, the design cost is reduced.

The application discloses a compressor blade phosphate aluminum coating precision grinding and polishing path generation method and system based on point cloud data processing, which comprises the following steps: obtaining point cloud data of the coating surface of the compressor blade by using a three-dimensional scanning device, filtering and denoising the point cloud data, and pre-processing the data, and aligning the pre-processed point cloud data with a robot machining reference coordinate system; based on the local curvature of the point cloud, the point cloud model is divided into multiple layers of contour slices by using a curvature adaptive slicing algorithm, wherein the slicing interval is negatively correlated with the local curvature radius; the contour points of each layer of slices are B-spline fitted to generate a continuous grinding and polishing path curve; the robot executes the grinding and polishing path curve, and integrates a real-time thickness detection device; when the thickness measurement value deviates from a preset threshold value, closed-loop feedback control is performed, and the grinding and polishing path and / or grinding and polishing parameters are dynamically compensated. The application has the advantages of high machining precision, high efficiency, high consistency and the like.

The invention discloses a gas compressor blade design method and device, electronic equipment and a storage medium, and the method comprises the following steps: S1, determining a boundary condition in the spanwise direction of a blade and a boundary condition in the flow direction of the blade, the boundary conditions in the flow direction of the blade comprise root section angle distribution, tip section angle distribution, root section thickness distribution and tip section thickness distribution; s2, closing the convective diffusion equation based on boundary conditions, giving a determined value of a control coefficient in the convective diffusion equation, and solving the convective diffusion equation to obtain angle distribution and thickness distribution of each section of the blade; compared with the prior art, the convection diffusion equation with clear physical significance is reasonably borrowed, the performance upper limit of the compressor blade is improved, meanwhile, the performance of the compressor blade is prevented from being reduced, practicability is high, and the method is suitable for wide application and popularization.

The invention belongs to the technical field of gas turbine gas compressor inspection and blade cleaning design, and particularly relates to a gas turbine gas compressor inspection and blade cleaning integrated structure which integrates a gas compressor inspection and blade cleaning structure on a gas compressor stator casing, so that holes in the gas compressor stator casing can be reduced, and the efficiency of gas turbine gas compressor inspection is improved. Compared with the prior art, large circumferential space is prevented from being occupied, interference between structures is reduced, the disassembly and assembly difficulty is reduced, the online cleaning function on the compressor blades is designed, the compressor blades can be kept clean and the pollution rate of the compressor blades can be reduced through daily online cleaning, the performance attenuation rate of the compressor can be further reduced, and the service life of the compressor is prolonged. The gas turbine can continuously work for a longer time, the maintenance cost is reduced, in addition, the cleaning nozzle is designed, the large spraying range and angle are achieved, the better atomization effect can be achieved, the cleaning efficiency and the cleaning effect can be improved, and deep cleaning of compressor blades can be achieved.

The present invention discloses a blade shaping method for controlling double leakage flow at the blade tip of a high-load compressor. By numerically simulating a high-load compressor blade grid, the transport characteristics of the blade tip leakage flow in the channel are analyzed, and the intersection position of the double leakage trajectory with the adjacent blade pressure surface is determined. With this as the center, a sinusoidal two-dimensional bulge is generated at different flow direction positions, which is then projected onto the pressure surface along the local normal. The starting / ending positions and peak values ​​of the two-dimensional bulges at different spanwise positions are adjusted, and stacked along the spanwise direction to obtain a blade tip three-dimensional geometric structure with the pressure surface belly bulge as the main feature. The blade shaping method of the present invention enhances the axial momentum of the blade tip through pressure field reconstruction, weakens the strength of the blade tip double leakage, and delays the upstream movement of the leakage flow-mainstream interface. While effectively widening the stable operating margin, it can also ensure the aerodynamic efficiency of the blade tip, and does not increase the complexity of the blade tip structure. It can be directly applied to the refined design of high-load compressor blade tips.

The application discloses a system identification method for mechanical damping of a compressor blade, and the method is as follows: a working condition is determined through steady CFD calculation of a multistage compressor, and is used as an initial field to carry out unsteady CFD calculation; blade modal force amplitude f and phase θ are calculated based on the results of the aerodynamic unsteady calculation and strength finite element analysis; aerodynamic damping under different pitch diameters is calculated by combining the influence coefficient method, and average aerodynamic damping coefficients of aerodynamic damping under all different pitch diameters are calculated; vibration response values of the blade under typical states in the test are measured, and are converted into blade generalized displacement q; the modal force amplitude f, the phase θ, the blade generalized displacement q and the average aerodynamic damping coefficients are brought into a formula to obtain mechanical damping of the blade. Advantageous effects: the real mechanical damping of the blade in a multistage environment is obtained, and based on the obtained mechanical damping, the prediction accuracy and reliability of forced response of the blade in subsequent blade type optimization and blade modification design of derivative types can be effectively improved.

The application discloses a compressor blade design method and device, electronic equipment and storage medium, including the following steps: S1: determining the boundary condition in the blade span direction and the boundary condition in the blade flow direction, wherein the boundary condition in the blade flow direction includes the root section angle distribution, the tip section angle distribution, the root section thickness distribution and the tip section thickness distribution; S2: closing the convection diffusion equation based on the boundary condition, giving the determined value of the control coefficient in the convection diffusion equation, solving the convection diffusion equation to obtain the angle distribution and thickness distribution of each section of the blade; compared with the prior art, the convection diffusion equation with clear physical meaning is reasonably borrowed, so that the performance upper limit of the compressor blade is improved, and the performance of the compressor blade is avoided from being reduced, the practicality is high, and the application is suitable for wide promotion and application.

The utility model relates to a special fixture for air conditioner compressor vane milling machine notch processing, including top head bottom plate one, top head bottom plate two, extruding piece, top head bottom plate one is connected with machine tool operation mesa through fixing piece, top head bottom plate two is detachably connected with top head bottom plate one through reciprocating friction cooperation, and extruding piece sets up on top head bottom plate two and is used for fixing work piece. The novel fixture provided by the utility model has simple structure, convenient operation, can effectively reduce the abrasion of machine tool mesa, prolongs machine tool service life, improves processing efficiency and precision simultaneously, and has higher practical value.

The invention discloses a gas compressor blade based on a distributed self-excitation adjustable fluid oscillator, and belongs to the technical field of aero-engines and gas turbines. Comprising a gas compressor blade body, a distributed self-excitation fluid oscillator and a parameter adjusting device based on piezoelectric ceramics. The multiple self-excitation fluid oscillators are arranged in the compressor blade body, each self-excitation fluid oscillator comprises a flow channel, an inlet hole, an outlet hole, a resonant cavity and a feedback channel, the resonant cavity communicates with the flow channel and the feedback channel, the outlet hole of the flow channel and an adjusting outlet of the resonant cavity are formed in the suction face, and the inlet hole of the resonant cavity is formed in the pressure face. Self-excitation is achieved through the pressure difference between the pressure surface and the suction surface of the blade, the flow of the resonant cavity is finely adjusted through the piezoelectric ceramic piece, the channel state and the outlet shape are fed back, the strong coupling limitation of the flow and the frequency of a traditional fluid oscillator is broken through, and autonomous control and three-in-one flexible and changeable control of different flows and frequencies of all outlets are achieved; the device has the advantages of simple structure, low energy consumption, high reliability and the like.

The invention belongs to the technical field of surface treatment after combined machining of parts, and relates to a method and related equipment for laser strengthening repair of easily-damaged parts of a gas compressor blade, a laser shock strengthening technology is adopted, the strength of the easily-damaged parts of a blade tip of the blade is enhanced, and the residual tensile stress of the blade is changed into residual compressive stress on the premise that the roughness is not changed. The invention provides a method for improving the surface integrity of the blade in order to ensure that the blade subjected to laser pulse shock peening treatment does not influence the flow field distribution of the working environment of the blade and under the condition that the surface roughness of the blade is not changed, the stress resistance is enhanced, the surface integrity of a part is improved, and the blade tip corner falling fault is reduced.

A gas turbine engine has a compression system radius ratio defined as the ratio of the radius of the tip of a fan blade to the radius of the tip of the most downstream compressor blade in the range of from 5 to 9. This results in an optimum balance between installation benefits, operability, maintenance requirements and engine efficiency when the gas turbine engine is installed on an aircraft.

The invention discloses a compressor blade axial locking structure, and relates to the technical field of gas turbine design. The impeller specifically comprises an impeller disc, blades, a check ring and a locking part between the impeller disc and the check ring, the root of each blade is a tenon, a plurality of mortises matched with the tenons are formed in the outer edge of the impeller disc, and the blades and the impeller disc are installed in a matched mode. The check ring is installed on the axial air outlet side of the wheel disc and used for limiting axial movement of the blades, the locking part is used for detachably connecting the check ring and the wheel disc, the locking part comprises a check ring locking block or a check ring hole on the check ring, and further comprises a lock head and a spring, and the outer edge, close to the check ring, of the wheel disc is provided with straight holes which are perpendicularly communicated with each other. And the wheel disc is detachably connected with the check ring by matching with the locking part. The axial locking structure of the compressor blade is easy to machine, few in assembly parts, convenient to disassemble and assemble and easy to repair after being damaged.