269 results about "Compressor blade" patented technology

The invention discloses a gas compressor blade mean camber line design method, a gas compressor blade, a gas compressor and an aero-engine, and the method comprises the steps: S1, obtaining known design data, including an inlet blade angle theta in, an outlet blade angle theta out, a mean camber line starting point coordinate, a mean camber line end point abscissa, and turning angles delta theta 1, delta theta 2 and delta theta 4 of a first section arc, a second section arc and a fourth section arc, and abscissas of all tangency points; and S2, each section of arc is calculated through a geometric analysis method, a mean camber line is constructed after circle center coordinates, arc radiuses and tangent point coordinates of each section of arc are calculated, and the mean camber line is formed by sequentially and tangentially connecting four sections of arcs according to the sequence of convex, concave, convex and concave. The gas compressor blade mean camber line design method is good in adaptability and multiple in controllable point in the high Mach number state; therefore, the requirements of an advanced compressor are met; design parameters are greatly reduced, the optimizing period is greatly shortened, the adjusting efficiency is high, and the time is short.

The invention provides a gas compressor moving blade tip clearance measuring method and device, and belongs to the field of aero-gas turbine gas compressors. The problems of low measurement precision, low measurement efficiency and small measurement range in the prior art are solved. The method comprises the following steps that an angle encoder is installed on a main shaft, and a three-dimensional laser profile measuring instrument is arranged on the blade tip side of a blade at intervals; the impeller is rotated, the three-dimensional laser profile measuring instrument scans the blade tip, and three-dimensional point cloud data of the whole circle of blade tip profile are obtained; after the casing and the impeller are assembled, a plurality of continuous blades on the impeller are detached, and the three-dimensional laser profile measuring instrument is arranged at the vacant position of the impeller; the impeller is rotated, the three-dimensional laser profile measuring instrument scans the inner wall of the casing, and three-dimensional point cloud data of the whole circle of the inner wall of the casing are obtained; processing data; registering the processed three-dimensional point cloud data of the blade tip contour and the processed three-dimensional point cloud data of the inner wall in the same coordinate system to form a registration model; and the blade tip clearance value of the whole circle is calculated. The method is mainly used for measuring the blade tip clearance.

The invention belongs to the technical field of aircraft engine compressor blade detection, and discloses an aircraft engine compressor blade erosion damage prediction method and system based on a PINN. The method comprises the following steps: constructing an erosion damage rate equation and a flow field-particulate matter coupling equation aiming at the erosion damage judgment of the blade of the gas compressor; performing dual-drive loss function design, constructing a physical information neural network (PINN) model, and performing adaptive training on the constructed physical information neural network (PINN) model; the invention relates to gas compressor blade erosion damage prediction and engineering application. According to the method, the limitation of the prior art is overcome, a more accurate and reliable damage prediction model is provided, and powerful support is provided for design, operation and maintenance and safety evaluation of the aero-engine. According to the innovative method, the physical law and the data driving technology can be combined, and the damage prediction precision and efficiency of the aero-engine compressor blade under the complex working condition are improved.

The invention discloses a turbofan engine compressor blade crack multi-modal monitoring system, and belongs to the field of multi-modal monitoring, and the system comprises a multi-modal acquisition module which obtains the multi-modal data of each compressor blade, carries out the dynamic filtering denoising and temperature compensation preprocessing, completes the time axis alignment, and carries out the data reconstruction and feature extraction; and constructing a multi-modal feature vector. And the triggering module carries out preliminary abnormality diagnosis according to the judgment result, judges whether the blade is abnormal or not through a key feature threshold, determines whether to trigger depth evaluation or not by combining a periodic threshold, identifies a potential crack blade and generates an excitation decision. And when an excitation decision is executed, the intelligent diagnosis module monitors the strain gradient change to perform excitation monitoring, compares the vibration characteristics of a potential crack blade with the vibration characteristics of a contrast blade, judges whether a crack exists or not, and realizes multi-stage early warning, so as to solve the problem that the crack of the turbofan engine compressor blade is difficult to quickly detect in real time in the prior art. And safe and stable operation of the engine is guaranteed.

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 discloses a gas compressor rotor blade surface defect detection device and a detection method thereof, and belongs to the technical field of gas compressor rotor blade detection. The gas compressor rotor blade surface defect detection device comprises a detection platform, an electric index plate clamping device, a tip cone stopping device and a flaw detection device, the electric index plate clamping device and the tip cone stopping device are matched to perform axial fixation and precise index driving on the two ends of the to-be-tested compressor rotor; the flaw detection device is used for detecting surface defects of the rotor blade of the gas compressor on the basis of an ultrasonic flaw detection principle; compared with existing artificial flaw detection of aero-engine compressor rotor blades, the method can accurately collect, process and analyze the surface defects of each stage of blade of the detected compressor rotor, can realize the positioning of each stage of blade and the traceability of the detection result, ensures the integrity of the measurement result, and improves the detection accuracy. The problem that the detection result is influenced by subjective factors in the existing artificial flaw detection of the aero-engine compressor rotor blade is solved.

The disclosure concerns a compressor blade for gas turbine engine. Specifically the blades of the compressor are modified according to predetermined requirements for both aerodynamic stability and fuel economy in multiple planes.

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 discloses a compressor blade with a venturi tube micro-channel cooling structure, which realizes uniform distribution of cooling air and forms a stable air film cooling layer on the surface of the blade by arranging arc-shaped fine channels and fine pores on the tip of the blade and combining the series design of adjacent channels, thereby obviously improving the cooling effect. The inner wall of the channel is of a sharkskin microcosmic placoid scale structure, friction resistance is reduced, the airflow turbulence characteristic is enhanced, and therefore the heat exchange performance is optimized. The grooves penetrating through the surface of the blade prolong the service life of the blade by changing the flowing path of the air flow and destroying the formation of a liquid film cooling layer, and enhance the cooling effect and the contact between the air flow and the surface of the blade through linkage with the small channels, thereby further improving the heat exchange performance. And the micropores of the damping boss not only provide a damping effect, but also can effectively release cooling air, so that the combination of cooling and structural stability is improved. While the cooling effect is optimized, the working efficiency of the blade is remarkably improved, and the service life of the blade is remarkably prolonged.

The invention provides a test device for simulating volcanic ash erosion sample materials at normal temperature. The test device comprises a gas supply system, a dust supply system, a gas-solid mixing system and an erosion test system, the gas supply system comprises an air compressor, a high-pressure gas storage tank and an air dryer which are connected in sequence, and an outlet of the air dryer is connected with the gas-solid mixing system; the dust supply system comprises a solid particle storage tank, the upper end of the solid particle storage tank is provided with a feed port, and a lower end outlet is connected with the gas-solid mixing system; the washout test system comprises a test chamber, a fixed disc which is arranged at an inlet of the test chamber and is communicated with the gas-solid mixing system, a nozzle on the fixed disc, and a sample disc which is positioned in the test chamber and is opposite to the fixed disc, and the nozzle is used for uniformly spraying solid particles onto a sample material on the sample disc. The test device can accurately simulate the erosive wear condition of the air compressor blade by the volcanic ash with different concentrations and speeds at normal temperature, and can also simulate the erosive damage condition of the air compressor blade and a thin-wall piece sample by the volcanic ash at a high rotating speed.

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 optimization method of an aero-engine warm-up process, which comprises the following steps of: firstly, acquiring corresponding samples of a high-pressure turbine blade tip clearance delta 1, a high-pressure compressor blade tip clearance delta 2, a minimum thrust Fm in a take-off process, a total oil consumption TFC in the whole process, warm-up time delta t and a warm-up rotating speed NH in the engine process; then, on the basis of the samples, a prediction model for delta 1, delta 2, Fm and TFC by delta t and NH is established through an artificial neural network; and finally, by taking the minimum values of delta 1, delta 2 and Fm required by actual use as constraints, selecting a combination of the lowest required machine warming time and machine warming rotating speed by utilizing an epsilon-constraint method, and if the combination is not unique, selecting a combination corresponding to the minimum TFC from a result to finish optimization of the machine warming process. According to the method, the artificial neural network is used for establishing the relation of the related variables along with the engine warming time and the engine warming rotating speed, the epsilon-constraint method is used for selecting the minimum required engine warming time under multiple constraints, and the engine warming process is scientifically optimized.

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 simulation method and system for the influence of the geometric deviation of a compressor blade on the performance of an aero-engine. The method comprises the steps that stage superposition is conducted on the characteristics of a multi-stage compressor based on the influence of the geometric deviation; superposing the characteristic change of the gas compressor caused by the geometric deviation of the blade to the general characteristic of the downstream gas compressor to obtain the change of the working point of the downstream gas compressor; after the characteristics of the gas compressor are obtained under the condition of blade geometric deviation, the characteristics are introduced into an engine performance calculation model, and changes of engine thrust, oil consumption rate and exhaust temperature are obtained. Computational analysis of gas compressor characteristics and engine performance is realized, influences of typical geometric deviations on gas compressor blade channel flow, engine surge margin, rotor power capability and working efficiency, engine thrust, oil consumption rate, exhaust temperature and the like are disclosed, and rapid prediction of engine performance changes under different blade geometric deviations is realized. And formulating the use specifications of the aero-engine under the condition that the compressor blades have geometric deviation in the support service process.

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.

The invention discloses an electrically-assisted turbocharger rotor experiment table and application thereof, which is characterized in that an axial locking nut, a thrust bearing and a floating ring bearing are used for supporting a shafting structure, equivalent compressor blades and a turbine disc are used for simulating mass, and a middle motor structure is used for driving; according to the invention, the static contact parameter identification experiment and the dynamic characteristic and parameter matching relation experiment of the electrically-assisted turbocharger can be realized. According to the method, the axial locking device is used for fixing the shafting part, adjusting the tightening torque and the tightening step length to change the pre-tightening parameters, and the static contact parameters are identified by combining metal processing process parameters. The matching form and matching parameters of the middle motor and the rotating shaft are changed, so that a dynamic characteristic and parameter matching relation experiment can be realized; an equivalent simulation counterweight disc is arranged on the central rotating shaft, and the weight and the number of counterweight bolts are adjusted to simulate rotor unbalance and typical assembly fault experiments.

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 invention belongs to the technical field of laser strengthening, and discloses a method for improving the wear resistance of a vibration reduction boss through micro-scale laser shock peening, and the method comprises the following steps: carrying out preprocessing treatment on a to-be-strengthened surface of the vibration reduction boss of a compressor blade; micro-scale laser surface impact strengthening is conducted on the to-be-strengthened surface of the blade vibration reduction boss obtained after preprocessing treatment, and the strengthened surface of the blade vibration reduction boss is obtained; and femtosecond laser shock peening is conducted on the strengthened surface of the blade vibration reduction boss, and the blade vibration reduction boss subjected to compound peening is obtained. According to the method, the abrasion resistance of the surface of the blade vibration reduction boss can be remarkably improved, meanwhile, the surface of the blade vibration reduction boss has low roughness, and the method has the technical advantages of being high in machining precision, uniform in distribution of an introduced residual stress field, high in machining efficiency and the like.

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.